Patent Publication Number: US-11391162-B2

Title: Spar with embedded plenum passage

Description:
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 may include low and high pressure compressors, and the turbine section may also include low and high pressure turbines. 
     Airfoils in the turbine section are typically formed of a superalloy and may include thermal barrier coatings to extend temperature capability and lifetime. Ceramic matrix composite (“CMC”) materials are also being considered for airfoils. Among other attractive properties, CMCs have high temperature resistance. Despite this attribute, however, there are unique challenges to implementing CMCs in airfoils. 
     SUMMARY 
     A spar for a vane arc segment of a gas turbine engine according to an example of the present disclosure includes an elongated spar leg that has a spar wall that circumscribes a core passage, and a plenum passage embedded in the spar wall between inner and outer portions of the spar wall. The outer portion of the spar wall has a plurality of cooling through-holes for emitting cooling air from the plenum passage. 
     In a further embodiment of any of the foregoing embodiments, the inner portion of the spar wall is fully solid such that the plenum passage is fluidly isolated from the core passage, and the plurality of cooling holes are exclusive outlets from plenum passage. 
     In a further embodiment of any of the foregoing embodiments, the spar leg includes ribs in the plenum passage that connect the inner and outer portions of the spar wall. 
     In a further embodiment of any of the foregoing embodiments, the ribs segregate the plenum passage into radial zones. 
     In a further embodiment of any of the foregoing embodiments, the ribs segregate the plenum passage into circumferential zones. 
     In a further embodiment of any of the foregoing embodiments, at least one of the ribs includes a through-hole. 
     A further embodiment of any of the foregoing embodiments includes a spar platform from which the spar leg extends, the spar platform having a cooling passage, and one of the ribs segregates the plenum passage from the cooling passage. 
     In a further embodiment of any of the foregoing embodiments, the ribs segregate the plenum passage into zones and define a manifold passage that interconnects the zones. 
     A vane arc segment for a gas turbine engine according to an example of the present disclosure includes an airfoil section that has an airfoil wall that defines a leading edge, a trailing edge, a pressure side, and a suction side. The airfoil section has an internal cavity, and a spar extends through the internal cavity for supporting the airfoil section. The spar is spaced from the airfoil wall. The spar has an elongated spar leg including a spar wall that circumscribes a core passage, and a plenum passage embedded in the spar wall between inner and outer portions of the spar wall. The outer portion of the spar wall has a plurality of cooling through-holes for emitting cooling air from the plenum passage toward the airfoil wall. 
     In a further embodiment of any of the foregoing embodiments, the inner portion of the spar wall is fully solid such that the plenum passage is fluidly isolated from the core passage, and the plurality of cooling holes are exclusive outlets from plenum passage. 
     In a further embodiment of any of the foregoing embodiments, the spar leg includes ribs in the plenum passage that connect the inner and outer portions of the spar wall. 
     In a further embodiment of any of the foregoing embodiments, the ribs segregate the plenum passage into radial zones. 
     In a further embodiment of any of the foregoing embodiments, the ribs segregate the plenum passage into circumferential zones. 
     In a further embodiment of any of the foregoing embodiments, at least one of the ribs includes a through-hole. 
     A further embodiment of any of the foregoing embodiments includes a spar platform from which the spar leg extends, the spar platform having a cooling passage, and one of the ribs segregates the plenum passage from the cooling passage. 
     In a further embodiment of any of the foregoing embodiments, the ribs segregate the plenum passage into zones and define a manifold passage that interconnects the zones. 
     In a further embodiment of any of the foregoing embodiments, the airfoil wall is ceramic and also defines first and second platforms. 
     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 vane arc segments disposed about a central axis of the gas turbine engine. The vane arc segments are as in 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 a gas turbine engine. 
         FIG. 2  illustrates an example vane arc segment of the gas turbine engine. 
         FIG. 3  a line representation of a sectioned view of the vane arc segment. 
         FIG. 4  illustrates a portion of a spar that has ribs that segregate a plenum passage in radial zones. 
         FIG. 5  illustrates a portion of a spar that has ribs that segregate a plenum passage in radial zones and a manifold passage. 
         FIG. 6  illustrates a portion of a spar that has ribs that segregate a plenum passage in circumferential zones. 
     
    
    
     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 housing  15  such as a fan case or nacelle, 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 through 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 example of select portions of a vane arc segment  60  from the turbine section  28  of the engine  20  (see also  FIG. 1 ). It is to be understood that although the examples herein are discussed in context of a vane from the turbine section, the examples are applicable to other cooled vanes that have support spars. 
     The vane arc segment  60  includes an airfoil section  62  that is formed by an airfoil wall  63 . The airfoil section  62  defines a leading edge  62   a , a trailing edge  62   b , and first and second sides  62   c / 62   d  that join the leading edge  62   a  and the trailing edge  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 section  62  generally extends in a radial direction relative to the central engine axis and spans from a first end  62   e  at an inner or first platform  64  to a second end  62   f  at a second or outer platform  66 . The terms “inner” and “outer” refer to location with respect to the central engine axis A, i.e., radially inner or radially outer. 
     The airfoil wall  63  is continuous in that the platforms  64 / 66  and airfoil section  62  constitute a one-piece body. As an example, the airfoil wall  63  is formed of a ceramic material, an organic matrix composite (OMC), or a metal matrix composite (MMC). For instance, the ceramic material is a monolithic ceramic or a ceramic matrix composite (CMC) that is formed of ceramic fibers that are disposed in a ceramic matrix. The monolithic ceramic may be, but is not limited to, SiC or other silicon-containing ceramic. The ceramic matrix composite may be, but is not limited to, a SiC/SiC ceramic matrix composite in which SiC fibers are disposed within a SiC matrix. Example organic matrix composites include, but are not limited to, glass fiber, carbon fiber, and/or aramid fibers disposed in a polymer matrix, such as epoxy. Example metal matrix composites include, but are not limited to, boron carbide fibers and/or alumina fibers disposed in a metal matrix, such as aluminum. The fibers may be provided in fiber plies, which may be woven or unidirectional and may collectively include plies of different fiber weave configurations. 
     The airfoil section  62  circumscribes an interior through-cavity  68 . The airfoil section  62  may have a single through-cavity  68  or, as shown, a rib  70  that that divides the interior through-cavity  68  into a forward cavity that is bound by the leading edge  62   a  portion of the airfoil wall  63  and an aft cavity that is bound by the trailing edge  62   b  portion of the airfoil wall  63 . It is to be appreciated that the airfoil section  62  may have additional ribs that further divide the through-cavity  68 . 
     The airfoil wall  63  is supported by a spar  72  and support hardware  73 .  FIG. 3  shows a line representation of a sectioned view of the vane arc segment  60  taken in the plane of the chord of the airfoil section  62 . The spar  72  includes a spar platform  72   a  and a spar leg  72   b  that extends from the spar platform  72   a  into the through-cavity  68 . The spar  72  is generally radially elongated and is secured with structural support S (e.g., a case) and support hardware  73 . The spar  72  thereby traps the airfoil wall  63  between the spar platform  72   a  and support hardware  73  to mechanically support the airfoil wall  63  and react out loads, such as aerodynamic loads. In this regard, the spar  72  may be formed of a relatively high temperature resistance, high strength material, such as a single crystal metal alloy (e.g., a single crystal nickel- or cobalt-alloy). 
     The spar  72  is formed by a spar wall  76  that has inner and outer wall portions  76   a / 76   b . The wall portions  76   a / 76   b  define a plenum passage  78  there between. The spar wall  76  circumscribes a core passage  72   c . The inner wall portion  76   a  is solid such that the plenum passage  78  is substantially or fully fluidly isolated from the core passage  72   c . The outer wall portion  76   b  has a plurality of cooling through-holes, represented by arrows  80 . The end  81  of the plenum passage  78  distal from the spar platform  72   a  is closed such that the holes  80  are the exclusive outlets of the plenum passage  78 . In inner wall portion  76   a  extends beyond the outer wall portion  76   b  and is secured to the support hardware  73 . 
     Cooling air F, such as bleed air from the compressor section  24 , is conveyed into and through the through-passage  72   c  of the spar  72 . This cooling air is destined for a downstream cooling location, such as a tangential onboard injector (TOBI). As indicated above, the through-passage  72   c  is substantially or fully isolated from the plenum passage  78 . Cooling air F is also conveyed through an inlet into the plenum passage  78  as a source of air for impingement cooling of the airfoil wall  63 . As the only exits from the plenum passage  78  are through the holes  80 , all of the cooling air in the plenum passage  78  is emitted as impingement cooling onto the airfoil wall  63 . For example, the impingement holes  80  are directed toward the leading edge  62   a . Alternatively or additionally, the cooling holes  78  may be directed toward the pressure side  62   c  and/or suction side  62   d . As the plenum passage  78  is isolated from the core passage  72   c , the cooling air F in the core passage  72   c  does not intermix with cooling air in the plenum passage  78 . Conversely, there may also be portions in the plenum passage  78  that receive less air flow than other regions. In that case, one or more orifices can be provided through the inner wall portion  76   a . The orifice or orifices serve as flow sinks that draw cooling air from the plenum passage  78  into the core passage  72   c , thereby pulling a greater amount of cooling air flow to the region that would otherwise have low flow. In general, however, due to desired flow margins of the cooling air for impingement and film cooling, the core passage  72   c  and the plenum passage  78  will exclude such orifices and be fully fluidly isolated. 
     In one example, the core passage  72   c  is provided with first pressurized air and the plenum passage  78  is provided with second pressurized air. The first and second pressurized air may differ in Mach number and thus also in pressure. For instance, the Mach number of the first pressurized air is greater than the Mach number of the second pressurized air, e.g., by a factor of 2-3 or more. At the expected Mach number of the first pressurized air, the air in the through-passage  72   c  is of insufficient pressure for impingement cooling. The pressurized air can come from the different sources (e.g., bleed air from different compressor stages) or the same source (same bleed air) that is divided into streams but that vary in pressure due to flow/exit paths. 
     In general, a leading edge of a turbine vane needs to be cooled. This is challenging in a two-cavity design with a forward spar that carries cooling air that is of insufficient pressure for impingement cooling. However, by substantially or fully isolating the core passage  72   c  from the plenum passage  78  and providing separate cooling air to the plenum passage  78 , the leading edge  62   a  portion of the airfoil wall  63  is provided with cooling, while maintaining the ability of the core passage  72   c  to convey cooling air for downstream use. 
     As discussed above, the spar  72  supports the airfoil wall  63 . The spar platform  72   a  and the spar leg  72   b  are thus structural. In this regard, the spar leg  72   b  may be of robust size in order to handle the structural loads. The robust size of the spar leg  72   b  takes up much of the volume of the through-cavity  68  in the airfoil section  62 , thus limiting the space available for a baffle to facilitate cooling of the airfoil wall  63 . In this regard, rather than a separate baffle, the plenum passage  78  is embedded in the spar wall  76  in order to deliver impingement cooling air to the airfoil wall  63 . Although the spar  72  in this example has a single one of the spar legs  72   b , it is to be appreciated that the spar may have one or more additional spar legs according to any of the examples herein, in which the one or more additional spar legs extend through other portions of the through cavity  68  (which may or may not be divided by ribs). 
     The spar  72  may be fabricated by investment casting to form the spar wall  76  around an investment skincore that forms the plenum passage  78 . A skincore is a thin, low aspect ratio investment casting core that that is typically injection molded from a material that contains ceramic or metal alloy. 
     As the investment minicore can be virtually any shape, the investment casting process permits the plenum passage  78  to include features that would be difficult or impossible to fabricate in a conventional sheet metal baffle. For example, as shown in  FIG. 3 , the spar leg  72   b  includes one or more internal ribs  82  in the plenum passage  78  that connect the inner and outer portions  76   a / 76   b  of the spar wall  76 . The ribs  82  serve to segregate the plenum passage  78  into two or more zones that provide cooling air to two or more zones of the airfoil wall  63 . Thus, the ribs  82  can be used to divide the flow of the cooling air F between the zones in order to tailor the cooling in each of the zones. Moreover, if as in the example in  FIG. 3 , the spar platform  72   a  has an internal cooling passage  72   d , one of the ribs  82 - 1  may segregate the plenum passage  78  from the cooling passage  72   d  such that the passages  72   d / 78  are fluidly isolated from each other. Additional ribs  82 - 1  may be provided in the internal cooling passage  72   d  in the platform the segregate it into zones. Additionally, as the spar  72  is structural, the ribs  82  also serve to mechanically reinforce the spar leg  72   b.    
       FIG. 4  illustrates one example in which the ribs  82  segregate the plenum passage  78  into radial zones Z 1 /Z 2 /Z 3 . In this example, the ribs  82  have through-holes  82   a  that serve to regulate flow of the cooling air between the zones Z 1 /Z 2 /Z 3 . 
       FIG. 5  includes a similar example except that the ribs  82  also define a radially elongated manifold passage  82   b  that interconnects the zones Z 1 /Z 2 /Z 3  via through-holes  82   a.    
       FIG. 6  illustrates another example, in which the view is of the plenum passage  78  radially looking inwards (toward the engine central axis A). In this example, the ribs  82  are radially elongated and segregate the plenum passage  78  into circumferential zones CZ 1 /CZ 2 /CZ 3 . As will be appreciated, the examples herein are not limiting, and various orientations of ribs  82  may be used to provide combinations of radial/circumferential zones and manifold passages. 
     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.