Patent Publication Number: US-10316668-B2

Title: Gas turbine engine component having curved turbulator

Description:
BACKGROUND 
     This disclosure relates to a gas turbine engine, and more particularly to a gas turbine engine component that includes at least one curved turbulator. 
     Gas turbine engines typically include a compressor section, a combustor section and a turbine section. In general, during operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
     Due to exposure to hot combustion gases, numerous components of the gas turbine engine may include internal cooling passages that route cooling air through the part. A variety of interior treatments may be incorporated into the internal cooling passages to augment the heat transfer effect and improve cooling. For example, some cooling passages may include pedestals, air-jet impingement, or turbulator treatments. 
     SUMMARY 
     A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a wall that forms a portion of an outer periphery of at least one cavity and at least one curved turbulator that extends from said wall. 
     In a further non-limiting embodiment of the foregoing component for a gas turbine engine, the component is one of a blade and a vane. 
     In a further non-limiting embodiment of either of the foregoing components for a gas turbine engine, the component is a blade outer air seal (BOAS). 
     In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, a plurality of curved turbulators are spaced along the wall. 
     In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one curved turbulator includes a contiguous body having at least one peak and at least one valley. 
     In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the contiguous body provides a smooth surface that excludes any sharp transition areas. 
     In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one curved turbulator is sinusoidal shaped. 
     In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, a row of film cooling holes are spaced from the at least one curved turbulator. 
     In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the row of film cooling holes includes a first film cooling hole and a second film cooling hole staggered from said first film cooling hole. 
     In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, a second turbulator extends from the wall and includes a different shape from the at least one curved turbulator. 
     In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one curved turbulator extends across a width of said wall. 
     In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one curved turbulator extends perpendicular to a direction of flow of cooling airflow communicated through the at least one cavity. 
     A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a first curved turbulator that protrudes into a cavity flow path and a second curved turbulator that protrudes into the cavity flow path at a position that is spaced from the first curved turbulator. A row of film cooling holes are disposed between the first curved turbulator and the second curved turbulator. 
     In a further non-limiting embodiment of the foregoing component for a gas turbine engine, the row of film cooling holes includes a first film cooling hole and a second film cooling hole that is staggered from the first film cooling hole. 
     In a further non-limiting embodiment of either of the foregoing components for a gas turbine engine, the first curved turbulator and the second curved turbulator are sinusoidal shaped. 
     In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, a pitch between the first curved turbulator and the second curved turbulator is continuously varied. 
     Nom A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a compressor section, a combustor section in fluid communication with the compressor section and a turbine section in fluid communication with the combustor section. A component extends into a core flow path of at least one of the compressor section and the turbine section, The component includes a wall that forms a portion of an outer periphery of at least one cavity of the component. At least one curved turbulator extends from the wall. 
     In a further non-limiting embodiment of the foregoing gas turbine engine, the component is an airfoil of the turbine section. 
     In a further non-limiting embodiment of either of the foregoing gas turbine engines, the component is a blade outer air seal (BOAS). 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, the wall is part of a platform of the component. 
     The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  illustrates a schematic, cross-sectional view of a gas turbine engine. 
         FIG. 2  illustrates a component that can be incorporated into a gas turbine engine. 
         FIG. 3  illustrates a cross-sectional view of the component of  FIG. 2 . 
         FIG. 4  illustrates a portion of a cooling circuit that can be incorporated into a gas turbine engine. 
         FIG. 5  illustrates another embodiment. 
         FIG. 6  shows yet another embodiment. 
         FIGS. 7A and 7B  illustrate exemplary turbulators. 
         FIG. 8  illustrates another turbulator embodiment. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The exemplary gas turbine engine  20  is a two-spool turbofan engine 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 for features. The fan section  22  drives air along a bypass flow path B, while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26 . The hot combustion gases generated in the combustor section  26  are expanded through the turbine section  28 . Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to turbofan engines and these teachings could extend to other types of engines, including but not limited to, three-spool engine architectures. 
     The gas turbine engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine centerline longitudinal axis A. The low speed spool  30  and the high speed spool  32  may be mounted relative to an engine static structure  33  via several bearing systems  31 . It should be understood that other bearing systems  31  may alternatively or additionally be provided. 
     The low speed spool  30  generally includes an inner shaft  34  that interconnects a fan  36 , a low pressure compressor  38  and a low pressure turbine  39 . The inner shaft  34  can be connected to the fan  36  through a geared architecture  45  to drive the fan  36  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  35  that interconnects a high pressure compressor  37  and a high pressure turbine  40 . In this embodiment, the inner shaft  34  and the outer shaft  35  are supported at various axial locations by bearing systems  31  positioned within the engine static structure  33 . 
     A combustor  42  is arranged between the high pressure compressor  37  and the high pressure turbine  40 . A mid-turbine frame  44  may be arranged generally between the high pressure turbine  40  and the low pressure turbine  39 . The mid-turbine frame  44  can support one or more bearing systems  31  of the turbine section  28 . The mid-turbine frame  44  may include one or more airfoils  46  that extend within the core flow path C. 
     The inner shaft  34  and the outer shaft  35  are concentric and rotate via the bearing systems  31  about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor  38  and the high pressure compressor  37 , is mixed with fuel and burned in the combustor  42 , and is then expanded over the high pressure turbine  40  and the low pressure turbine  39 . The high pressure turbine  40  and the low pressure turbine  39  rotationally drive the respective high speed spool  32  and the low speed spool  30  in response to the expansion. 
     The pressure ratio of the low pressure turbine  39  can be pressure measured prior to the inlet of the low pressure turbine  39  as related to the pressure at the outlet of the low pressure turbine  39  and prior to an exhaust nozzle of the gas turbine engine  20 . In one non-limiting embodiment, the bypass ratio of the gas turbine engine  20  is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  38 , and the low pressure turbine  39  has a pressure ratio that is greater than about five (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines, including direct drive turbofans. 
     In this embodiment of the exemplary gas turbine engine  20 , a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section  22  of the gas turbine engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine  20  at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. 
     Fan Pressure Ratio is the pressure ratio across a blade of the fan section  22  without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine  20  is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of [(Tram° R)/(518.7° R)] 0.5 , where T represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine  20  is less than about 1150 fps (351 m/s). 
     Each of the compressor section  24  and the turbine section  28  may include alternating rows of rotor assemblies and vane assemblies (shown schematically) that carry airfoils that extend into the core flow path C. For example, the rotor assemblies can carry a plurality of rotating blades  25 , while each vane assembly can carry a plurality of vanes  27  that extend into the core flow path C. The blades  25  create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine  20  along the core flow path C. The vanes  27  direct the core airflow to the blades  25  to either add or extract energy. 
     Various components of the gas turbine engine  20 , including but not limited to the airfoils of the blades  25  and the vanes  27  of the compressor section  24  and the turbine section  28 , may be subjected to repetitive thermal cycling under widely ranging temperatures and pressures. The hardware of the turbine section  28  is particularly subjected to relatively extreme operating conditions. Therefore, some components may require internal cooling circuits for cooling the parts during engine operation. 
     This disclosure relates to curved turbulators that can be incorporated into the walls of internal cooling cavities of gas turbine engine components. Among other benefits, the exemplary curved turbulators provide reduced stress concentrations and increased flexibility of film cooling hole placement as compared to prior art interior treatments. 
       FIGS. 2 and 3  illustrate a component  50  that can be incorporated into a gas turbine engine, such as the gas turbine engine  20  of  FIG. 1 . The component  50  may include a body portion  52  that axially extends between a leading edge portion  54  and a trailing edge portion  56 . The body portion  52  may additional include a first wall  58  (e.g., a pressure side wall) and a second wall  60  (e.g., a suction side wall) that are spaced apart from one another and that join at each of the leading edge portion  54  and the trailing edge portion  56 . 
     In this embodiment, the body portion  52  is representative of an airfoil. For example, the body portion  52  could be an airfoil that extends between inner and outer platforms (not shown) where the component  50  is a vane, or could extend from platform and root portions (also not shown) where the component  50  is a blade. Alternatively, the component  50  could be a non-airfoil component, including but not limited to a blade outer air seal (BOAS), a combustor liner, a turbine exhaust case liner, or any other part that may require dedicated cooling. 
     A gas path  62  is communicated axially downstream through the gas turbine engine  20  along the core flow path C (see  FIG. 1 ) in a direction that extends from the leading edge portion  54  toward the trailing edge portion  56  of the body portion  52 . The gas path  62  represents the communication of core airflow along the core flow path C. 
     One or more cavities  72  may be disposed inside of the body portion  52  as part of an internal cooling circuit for cooling portions of the component  50 . The cavities  72  may extend radially, axially and/or circumferentially inside of the body portion  52  to establish cooling passages for receiving a cooling airflow  68  to cool the component  50 . The cooling airflow  68  may be communicated into one or more of the cavities  72  from an airflow source  70  that is external to the component  50 . 
     The cooling airflow  68  is generally of a lower temperature than the airflow of the gas path  62  that is communicated across the body portion  52 . In one particular embodiment, the cooling airflow  68  is a bleed airflow that can be sourced from the compressor section  24  or any other portion of the gas turbine engine  20  that includes a lower temperature and higher pressure than the component  50 . The cooling airflow  68  can be circulated through the cavities  72 , such as along a serpentine path, to transfer thermal energy from the component  50  to the cooling airflow  68  thereby cooling the component  50 . The cooling circuit can include any number of cavities  72 . The cavities  72  may be in fluid communication with one another or could alternatively be isolated from one another. 
     One or more ribs  74  may extend between the first wall  58  and the second wall  60  of the body portion  52 . The rib(s)  74  divide the cavities  72  from one another. 
     As discussed in greater detail below, at least one of the cavities  72  can include one or more curved turbulators  80  that protrude into a cavity flow path  82  of the cavity  72  to disrupt the thermal boundary layer of the cooling airflow  68  and increase the cooling effectiveness of the internal cooling circuit of the component  50 . In one embodiment, the curved turbulators  80  are miniature walls protruding into the cavity flow path  82 . The design, configuration and placement of the numerous curved turbulators  80  shown by  FIGS. 2 and 3  are exemplary only and are not intended to limit this disclosure. 
       FIG. 4  illustrates a wall  84  of a cavity  72  of a component (e.g., the component  50 ). The wall  84  forms a portion of an outer periphery of the cavity  72 . The wall  84  could be an internal surface of either the first wall  58  or the second wall  60  (see  FIGS. 2 and 3 ) that faces into the cavity  72 , or could extend along one of the ribs  74 . 
     A curved turbulator  80  may extend from the wall  84 . In this embodiment, the wall  84  of the cavity  72  includes a plurality of curved turbulators  80 . The curved turbulators  80  can span a width W of the wall  84  and extend substantially perpendicular to the direction of flow of the cooling airflow  68  within a cavity flow path  82  of the cavity  72 . Due to the continuous curvature of the curved turbulators  80 , a pitch P (e.g., a spacing) between each adjacent curved turbulator  80  is continuously varied. 
     A row of film cooling holes  86  can be disposed between radially adjacent curved turbulators  80 . In this embodiment, each row of film cooling holes  86  includes a first film cooling hole  86 A and a second film cooling hole  86 B that is radially staggered from the first film cooling hole  86 A. Of course, additional film cooling holes than are shown in this embodiment could be disposed through the wall  84  in each row of film cooling holes  86 . The film cooling holes  86 A,  86 B do not intersect through any curved turbulator  80  because of the wavy design of the curved turbulators  80 . Other portions of the wall  84  may exclude film cooling holes  86  between adjacent curved turbulators  80 . 
     The curved turbulators  80  are configurable in a variety of patterns. For example, as shown in  FIG. 4 , a plurality of curved turbulators  80  can be radially disposed along the wall  84 . In another embodiment, the wall  84  can include a combination of alternating curved turbulators  80 A and V-shaped turbulators  80 B (see  FIG. 5 ). In yet another embodiment, the wall  84  could include a first cluster C 1  of curved turbulators  80 A and a second cluster C 2  of turbulators  80 B embodying a different design than the curved turbulators  80 A (see  FIG. 6 ). Other configurations and patterns are also contemplated. The configuration of the various wall treatments can vary based on streamwise profiles, height, spacing, boundary layer shape and other design criteria. 
       FIG. 7A  illustrates one exemplary curved turbulator  80  that can be incorporated into a gas turbine engine component cooling circuit. In this embodiment, the curved turbulator  80  includes a contiguous body  90  that includes at least one peak  92  and at least one valley  94 . The contiguous body  90  includes a completely smooth surface that excludes any sharp transition areas. The curved turbulator  80  could also exclude any peak  92  (see  FIG. 7B ). 
       FIG. 8  illustrates another curved turbulator  180 . The curved turbulator  180  of this embodiment is sinusoidal shaped. The curved turbulator  180  may include a plurality of peaks  192  and a plurality of valleys  194  extending along a smooth, contiguous body  190 . 
     The curved turbulators of this disclosure may embody any curved or wavy geometry that provides a smooth transition surface that is capable of accommodating relatively large variations in the streamwise positioning of the turbulators relative to the cooling airflow that flows within the cavities. The exemplary curved turbulators also provide reduced stress concentrations as compared to treatments having more angular designs, such as V-shaped turbulators. 
     Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
     It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure. 
     The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would understand that certain modifications could come within the scope of this disclosure. For these reasons, the following claims should be studied to determine the true scope and content of this disclosure.