Patent Publication Number: US-2023160308-A1

Title: Airfoil with rib having connector arms

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     The present disclosure is a continuation of U.S. Pat. Application No. 16/545,417 filed Aug. 20, 2019. 
    
    
     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-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 high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction. 
     SUMMARY 
     An airfoil according to an example of the present disclosure includes an airfoil wall that defines a leading end, a trailing end, and first and second sides that join the leading end and the trailing end. A rib connects the first and second sides of the airfoil wall and defines a tube portion that circumscribes a rib passage. First and second connector arms solely join the tube portion to, respectively, the first and second sides of the airfoil wall. 
     In a further embodiment of any of the foregoing embodiments, the airfoil wall and the rib bound a cooling channel there between, and the cooling channel is flow isolated from the rib passage. 
     In a further embodiment of any of the foregoing embodiments, the airfoil wall and the rib bound a cooling channel there between, and the rib includes at least one cooling aperture connecting the cooling channel and the rib passage. 
     In a further embodiment of any of the foregoing embodiments, the tube portion includes forward and aft walls and first and second side walls joining the forward and aft walls. The first connector arm projects from the first side wall and the second connector arm projects from the second side wall. 
     In a further embodiment of any of the foregoing embodiments, the airfoil wall and the rib bound a cooling channel there between, and at least one of the first and second side walls includes at least one cooling aperture connecting the rib passage and the cooling channel. 
     In a further embodiment of any of the foregoing embodiments, the airfoil wall and the rib bound a cooling channel there between, at least one of the first and second side walls includes at least one cooling aperture connecting the rib passage and the cooling channel, and the at least one cooling aperture is aft of the first and second connector arms. 
     In a further embodiment of any of the foregoing embodiments, the airfoil wall and the rib bound a cooling channel there between, and forward wall includes at least one cooling aperture connecting the rib passage and the cooling channel. 
     An airfoil according to an example of the present disclosure includes an airfoil wall that defines a leading end, a trailing end, and first and second sides that join the leading end and the trailing end. The first and second sides span in a radial direction, and first, second, and third ribs connect the first and second sides of the airfoil wall. Each of the first, second, and third ribs define a tube portion that circumscribes a rib passage. First and second connector arms solely join the tube portion to, respectively, the first and second sides of the airfoil wall. The first rib and the airfoil wall bound a first cooling channel there between. The first rib, the second rib, and the airfoil wall bound a second cooling channel there between. The second rib, the third rib, and the airfoil wall bound a third cooling channel there between. 
     In a further embodiment of any of the foregoing embodiments, the rib passages are flow isolated from each of the first cooling channel, the second cooling channel, and the third cooling channel. 
     In a further embodiment of any of the foregoing embodiments, the second cooling channel and the third cooling channel are connected in a serpentine flow pattern. 
     In a further embodiment of any of the foregoing embodiments, the first cooling channel is flow isolated from the second cooling channel and the third cooling channel. 
     In a further embodiment of any of the foregoing embodiments, the first cooling channel, the second cooling channel, and the third cooling channel are flow isolated from each other. 
     A further embodiment according to an example of the present disclosure includes a fourth rib that also connects the first and second sides of the airfoil wall and also defines a respective tube portion that circumscribes a respective rib passage, with respective first and second connector arms that solely join the tube portion to, respectively, the first and second sides of the airfoil wall. 
     In a further embodiment of any of the foregoing embodiments, the third rib, the fourth rib, and the airfoil wall bound a fourth cooling channel there between. 
     In a further embodiment of any of the foregoing embodiments, the fourth rib includes at least one cooling aperture aft of the respective connector arms of the fourth rib. 
     In a further embodiment of any of the foregoing embodiments, the tube portion of each of the first, second, and third ribs includes at least one cooling aperture. 
     In a further embodiment of any of the foregoing embodiments,, except for connection through the first and second wall of the airfoil wall, the first rib, the second rib, and the third rib are disjoined from each other. 
     In a further embodiment of any of the foregoing embodiments, the airfoil wall includes cooling apertures connecting each of the first cooling channel, the second cooling channel, and the third cooling channel to an exterior gaspath. 
     In a further embodiment of any of the foregoing embodiments, at least one of the first cooling channel, the second cooling channel, and the third cooling channel is connected to an exterior gaspath through cooling apertures in the first side of the airfoil wall but not the second side of the airfoil wall. 
     A gas turbine engine according to an example of the present disclosure includes a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section has an airfoil according to any of the foregoing embodiments. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The various features and advantages of the present 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. 
         FIG.  1    illustrates an example gas turbine engine. 
         FIG.  2    illustrates a sectioned view of an example airfoil of the engine. 
         FIG.  3    illustrates a perspective view of the airfoil of  FIG.  2   . 
         FIG.  4    illustrates another example airfoil for use in the engine and that has multiple ribs. 
         FIGS.  5 A and  5 B  illustrate a radial cooling air flow pattern of the airfoil of  FIG.  4   . 
         FIGS.  6 A and  6 B  illustrate a modified airfoil that is configured for a serpentine cooling air flow pattern. 
         FIGS.  7 A and  7 B  illustrate a modified airfoil that is configured for an impingement/axial cooling air flow pattern. 
         FIGS.  8 A and  8 B  illustrate a modified airfoil that is configured for a full impingement cooling air flow pattern. 
     
    
    
     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 . 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 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 a 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  may be 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 the low pressure compressor, or aft of the combustor section  26  or even aft of turbine section  28 , and fan  42  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 and less than about 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 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). 
       FIG.  2    illustrates a representative sectioned view of an example of an airfoil  60  used in the turbine engine  20  (see also  FIG.  1   ), and  FIG.  3    illustrates a perspective view of the airfoil  60 . As shown, the airfoil  60  is a turbine blade; however, it is to be understood that, although the examples herein may be described and shown with reference to turbine blades or vanes, this disclosure is also applicable to cooled blades or vanes in other locations than shown. 
     The airfoil  60  includes an (outer) airfoil wall  62  that delimits the aerodynamic profile of the airfoil  60 . In this regard, the wall  62  defines a leading end  62   a , a trailing end  62   b , and first and second sides  62   c / 62   d  that join the leading end  62   a  and the trailing end  62   b . In this example, the first side  62   c  is a pressure side and the second side  62   d  is a suction side. The airfoil wall  62  generally extends in an axial direction relative to the central engine axis A. For a blade, the airfoil wall  62  will typically span in a longitudinal direction from an inner platform to a free tip end. In a vane, the airfoil wall  62  will typically span in a longitudinal direction from an inner platform to an outer platform. 
     The airfoil  60  further includes at least one rib  64  that connects the first and second sides  62   c / 62   d  of the airfoil wall  62 . The rib  64  is generally longitudinally elongated between an inner diameter and outer diameter such that it spans the full or substantially full longitudinal distance of the airfoil wall  62 . The term substantially full refers to at least 70% of the longitudinal distance between the inner diameter and outer diameter. The rib  64  defines a tube portion  66  that circumscribes a rib passage  68 , and first and second connector arms  70   a / 70   b  that solely join the tube portion  66  to, respectively, the first and second sides  62   c / 62   d  of the airfoil wall  62 . As used herein, the phrase “solely join” or variations thereof refers to the arm  70   a  being the exclusive structural attachment of the tube portion  66  to the first side  62   c  and the arm  70   b  being the exclusive structural attachment of the tube portion  66  to the second side  62   d . Such an attachment configuration permits the rib  64  to reinforce the sides  62   c / 62   d  to facilitate reduction in bulging from internal pressure, while still permitting the rib  64  to move and thermally expand and contract at a different rate than the sides  62   c / 62   d  during thermal cycling. 
     In the illustrated example, the tube portion  66  is generally rectangular and includes forward and aft walls  66   a / 66   b  and first and second side walls  66   c / 66   d  that join the forward and aft walls  66   a / 66   b . However, other shapes, such as ellipses or triangles may also be used. The first connector arm  70   a  projects from the first side wall  66   c  and the second connector arm  70   b  projects from the second side wall  66   d . 
     The rib  64  partitions the interior cavity of the airfoil  60  such that the airfoil wall  62  and the rib  64  bound cooling channels  72  there between. Cooling air, such as bleed air from the compressor section  24  of the engine  20  can be provided to the cooling channels  72  and the rib passage  68 . The cooling air can be fed from a radially inner or radially outer location into the cooling channels  72  and rib passage  68 . For example, the tube portion  66  is continuous such that the cooling channels  72  are flow isolated from the rib passage  68 . As used herein, the phrase “flow isolated” or variations thereof refers to passages, channels, or both that are not fluidly connected to each other within the airfoil  60  such that air cannot flow within the airfoil  60  from one passage or channel to the other passage or channel. For instance, such flow isolation permits air in the channels and passages to be used for different purposes or at differential pressures. In this regard, cooling air in the cooling channels  72  can be discharged through cooling holes or the like in the side walls  62   c / 62   d  to serve for cooling the side walls  62   c / 62   d , while cooling air in the rib passage  68  can serve to cool a blade tip or platform or be provided to other downstream structures. 
     Alternatively, the tube portion  66  of the rib  64  can include one or more cooling apertures  74  that connect the rib passage  68  with one or both of the cooling channels  72 . For instance, the cooling apertures  74  open toward the side walls  62   c / 62   d  to provide impingement cooling onto the interior surfaces of the side walls  62   c / 62   d . In this regard, cooling air flows out from the rib passage  68  through cooling apertures  74  and into cooling channels  72 . One or both of the forward and aft walls  66   a / 66   b  of the tube portion  66  can exclude any cooling apertures, as these walls are not adjacent the airfoil wall  62 . In one further example, the tube portion  66  only includes cooling apertures  74  aft of the connector arms  70   a / 70   b  such that the forward one of the cooling channels  72  has flow that is isolated from the rib passage  68  and the aft one of the cooling channels  72 . Alternatively, cooling apertures  74  may only be included on the forward side of connector arms  70   a / 70   b  such that the aft cooling channel  72  is isolated from rib passage  68 . In another embodiment, cooling apertures  74  may only be included on one of the rib sidewalls; rib second sidewall  66   d , for example. 
       FIG.  4    illustrates another example airfoil  160 . In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding elements. In this example, the airfoil  160  includes multiple ribs similar to the rib  64  described above, including first, second, third, and fourth ribs  164   a / 164   b / 164   c / 164   d . Like rib  64 , each of the ribs  164   a / 164   b / 164   c / 164   d  connects the first and second sides  62   c / 62   d  of the airfoil wall  62 . Except for connection through the first and second wall  62   c / 62   d  of the airfoil wall  62 , the ribs  164   a / 164   b / 164   c / 164   d  are disjoined from each other. As used herein, the term “disjoined” refers to the ribs  164   a / 164   b / 164   c / 164   d  excluding any structural attachments to each other. Such an attachment configuration permits the ribs  164   a / 164   b / 164   c / 164   d  to reinforce the sides  62   c / 62   d  to facilitate reduction in bulging from internal pressure, while still permitting the ribs  164   a / 164   b / 164   c / 164   d  to move and thermally expand and contract at a different rate than the sides  62   c / 62   d  during thermal cycling. It is to be appreciated that, although four ribs are shown in the illustrated example, the airfoils herein may include fewer or more than four ribs, such as a single rib, two ribs, three ribs, or five or more ribs. 
     Each of the ribs  164   a / 164   b / 164   c / 164   d  includes a respective tube portion  66  that circumscribes a respective rib passage  68 , and first and second connector arms  70   a / 70   b  that solely join the tube portion  66  to the first and second sides  62   c / 62   d  of the airfoil wall  62 . In this example, the tube portions  66  of the first, second, and third ribs  164   a / 164   b / 164   c  are generally rectangular, similar to the tube portion  66  of the rib  64  described above. However, the fourth, aft-most rib  164   d  tapers and is generally triangular, to follow the taper of the airfoil  160  toward the trailing end  62   b . 
     The first rib  164   a  and the airfoil wall  62  bound a first cooling channel  172   a  there between. The first rib  164   a , the second rib  164   b , and the airfoil wall  62  bound a second cooling channel  172   b  there between. The second rib  164   b , the third rib  164   c , and the airfoil wall  62  bound a third cooling channel  172   c  there between. The third rib  164   c , the fourth rib  164   d , and the airfoil wall  62  bound a fourth cooling channel  172   d  there between. The fourth rib  164   d  and the airfoil wall  62  also bound a fifth cooling channel  172   e . For example, the cooling channels  172   a / 172   b / 172   c / 172   d / 172   e  are flow isolated from each other. 
     The tube portions  66  can be adapted to provide a desired cooling air flow pattern. In one example, the tube portions  66  are continuous such that all of the cooling channels  172   a / 172   b / 172   c / 172   d  are flow isolated from each other and from all of the rib passages  68 .  FIGS.  5 A and  5 B  illustrate an example configured for a radial cooling air flow pattern. It is to be appreciated that the numbering of many of the structures from  FIG.  4    has been excluded in order to more clearly show the flow pattern. In this example, cooling air is provided into the rib passages  68  and flows longitudinally (radially), as indicated by flow arrows F1. Cooling air is also provided to the cooling channels  172   a / 172   b / 172   c / 172   d  and flows longitudinally (radially), as indicated by flow arrows F2. The cooling air in the cooling channels  172   a / 172   b / 172   c / 172   d  can be discharged through cooling holes or the like in the airfoil wall  62 , as indicated by flow arrows F 3 . The cooling flow in the cooling channel  172   e  can also flow axially and be discharged through holes or slots  76  in the trailing end  62   b , as indicated by flow arrows F4. In the illustrated example the cooling air is fed from the bottom of the airfoil  160 , which is the radially inner end of the airfoil  160  and is consistent with flow provided to a blade. Alternatively, for a vane, the cooling air could be fed from the top of the airfoil, which is the radially outer end. 
       FIGS.  6 A and  6 B  illustrate a modified example of an airfoil  260 . The airfoil  260  is similar to the airfoil  160  but is configured for a serpentine cooling air flow pattern. In this example, rather than purely radial flow in the second, third, and fourth cooling channels  172   b / 172   c / 172   d , the second, third, and fourth cooling channels  172   b / 172   c / 172   d  are connected in a serpentine flow pattern, indicated at flow arrow F5. In this regard, the airfoil  260  may include connector or turn passages  78  at the radial ends of the airfoil  260  (or in the platforms for a vane) to transfer flow between the cooling channels  172   b / 172   c / 172   d . The cooling air in the cooling channels  172   b / 172   c / 172   d  can be discharged through cooling holes or the like in the airfoil wall  62 , as indicated by flow arrows F 3 . 
       FIGS.  7 A and  7 B  illustrate a modified example of an airfoil  360 . The airfoil  360  is similar to the airfoil  160  but is configured for an impingement/axial cooling air flow pattern. In this example, the first rib  164   a  and the fourth rib  164   d  include cooling apertures  74  such that cooling air from the respective rib passages  68  impinges onto the interior surfaces of the side walls  62   c / 62   d , as indicated by flow arrows F6. The impingement air discharged into the fifth cooling channel  172   e  then flows axially and is discharged through holes or slots  76  in the trailing end  62   b . The impingement air discharged into the first cooling channel  172   a  impinges at the interior surface of the leading end  62   a  and is discharged through cooling holes or the like in the airfoil wall  62  as indicated by flow arrows F 3 . 
       FIGS.  8 A and  8 B  illustrate a modified example of an airfoil  460 . The airfoil  460  is similar to the airfoil  360  but is configured for a full impingement cooling air flow pattern. In this example, each of the ribs  164   a / 164   b / 164   c / 164   d  includes cooling apertures  74  such that cooling air in the respective rib passages  68  impinges onto the interior surfaces of the side walls  62   c / 62   d , as indicated by flow arrows F6. 
     The above-described configurations of the ribs  164   a / 164   b / 164   c / 164   d  and the cooling channels  172   a / 172   b / 172   c / 172   d / 172   e  facilitates cooling of the airfoil  160 / 260 / 360 / 460 . For example, the cooling channels  172   b / 172   c / 172   d  are generally “H” shaped, with the side lobe portions of the “H” extending along, respectively, the first and second sides  62   c / 62   d  of the airfoil wall and the middle portion of the “H” extending between the respective ribs  164   a / 164   b / 164   c / 164   d . The side lobe portions of the “H” have a low dimensional aspect ratio, meaning the axial distance of the side lobe portions of the “H” between adjacent rib connector arms is larger than the circumferential distance between the sides  62   c / 62   d  of the airfoil wall  62  and the rib sidewalls  66   c / 66   d . Such a shape facilitates providing much of the cooling air flow along the sides  62   c / 62   d  of the airfoil wall  62  to enhance cooling. The middle portion of the “H” also connects the side lobe portions of the “H” such that cooling flow can migrate from the second side  62   d  to the first side  62   c . For instance, the flow can be discharged exclusively or primarily through cooling holes on the higher pressure first side  62   c , which facilitates reducing gaspath mixing losses that may be incurred if the cooling air was instead discharged to the lower gaspath pressure at the second side  62   d . Additionally, if cooling apertures are only included on the second rib sidewall  66   d , the cooling flow may be concentrated to impinge on the second airfoil sidewall  62   c  before flowing through the middle portion of the “H” and out through cooling holes on the first side  62   c . Moreover, since the tube portions  66  are only connected to the sides  62   c / 62   d  via the connector arms  70   a / 70   b , the rib passages  68  are substantially thermally isolated from the sides  62   c / 62   d . Cooling air in the rib passages  68  thus remains relatively cool, to enhance cooling of the tip of the airfoil, platform of the airfoil, or other downstream structure. 
     The airfoils described herein may be fabricated from superalloys using such processes as investment casting or additive manufacturing. For example, in an investment casting process, an investment core is fabricated and then used in the casting of the superalloy to define internal features in the airfoil. Such an investment core can be formed from a ceramic or other suitable material in a molding process in which the ceramic or other material is injected into the cavity of a molding die. To form the ribs described herein, a sacrificial body with the shape of the ribs is inserted into the cavity and the ceramic or other material is molded around the sacrificial body. The sacrificial body may be formed of a thermoplastic or other material that can readily later be removed without damaging the core. The sacrificial body is then removed, such as by melting, leaving open the pattern of the ribs in the investment core. Subsequently, when the core is used in the investment casting, the superalloy fills the open pattern of the ribs left by the sacrificial body in the core, thereby forming the walls of the ribs. 
     Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments. 
     The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.