Patent Publication Number: US-2018038236-A1

Title: Gas turbine engine stator vane baffle arrangement

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a divisional of U.S. patent application Ser. No. 14/708,830 filed May 11, 2015 which claims priority to U.S. Provisional Application No. 62/001,939 which was filed on May 22, 2014 and is incorporated herein by reference. 
    
    
     BACKGROUND 
     This disclosure relates to a gas turbine engine turbine stator vane with a baffle. 
     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 combustor 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. 
     Some stator vane cooling configurations include a cooling cavity with a baffle arranged within the cavity. The baffle may be constructed from a sheet steel and is supported relative to the exterior wall of the stator vane by radially extending ribs in the exterior wall of the airfoil from the inner platform toward the outer platform. 
     To enhance cooling within the cooling cavity in the area between the baffle and the exterior wall, trip strips may be provided on the exterior wall. The trip strips increase the turbulence of the cooling fluid to enhance heat transfer. 
     SUMMARY 
     In one exemplary embodiment, a method of flowing cooling fluid through a stator vane in a gas turbine engine includes the step of providing an airfoil that has an exterior wall that provides a cooling cavity. The exterior surface has an interior surface that has multiple pin fins that extend therefrom. A baffle is arranged in the cooling cavity and supported by the pin fins. A perimeter cavity is provided between the baffle and the exterior wall. The pin fins are arranged in the perimeter cavity. Cooling fluid flows through a region in the perimeter cavity. The pin fins are arranged in the region having a low Reynolds number and through which the cooling fluid flows. 
     In a further embodiment of any of the above, the baffle is sheet steel. 
     In a further embodiment of any of the above, the exterior wall provides pressure and suction sides joined at leading and trailing edges. The baffle includes impingement holes configured to provide impingement cooling fluid onto the exterior wall at the leading edge. 
     In a further embodiment of any of the above, the baffle includes a generally smooth outer contour free of protrusions. 
     In a further embodiment of any of the above, the outer contour is provided by plastic deformation. 
     In a further embodiment of any of the above, cooling holes are provided by at least one of drilling, laser drilling, or electro discharge machining. 
     In a further embodiment of any of the above, the perimeter cavity circumscribes the baffle. 
     In a further embodiment of any of the above, the pin fins provide the sole support for the baffle in the perimeter cavity. 
     In a further embodiment of any of the above, the pin fins are arranged in rows. 
     In a further embodiment of any of the above, the pin fins are radially spaced from one another. 
     In a further embodiment of any of the above, a rib separates the cooling cavity from a trailing edge cooling cavity and the rib includes holes. 
     In a further embodiment of any of the above, the pin fins are integral with the exterior wall. 
     In a further embodiment of any of the above, the airfoil is a nickel alloy. 
     In a further embodiment of any of the above, the low Reynolds number corresponds to a laminar or near-laminar flow of the cooling fluid. 
     In a further embodiment of any of the above, the Reynolds number is less than 4000. 
     In a further embodiment of any of the above, the Reynolds number is less than 1500. 
     In a further embodiment of any of the above, the region has a Nusselt number less than 40. 
     In another exemplary embodiment, an assembly for a gas turbine engine includes an airfoil that has an exterior wall that provides a cooling cavity. The exterior wall has an interior surface that has multiple pin fins extending therefrom. A baffle is arranged in the cooling cavity and supported by the pin fins. The pin fins are arranged in a region with a low Reynolds number. A cooling source is in fluid communication with one side of the baffle. A component is in fluid communication with another side of the baffle. Cooling fluid is configured to flow from the cooling source through the baffle to the component. 
     In a further embodiment of any of the above, the component is a downstream airfoil. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
         FIG. 1  schematically illustrates a gas turbine engine embodiment. 
         FIG. 2  is a schematic view through an engine section including a fixed stage and a rotating stage. 
         FIG. 3  is a schematic view of a stator vane and associated cooling path. 
         FIG. 4  is a cross-sectional view through an airfoil depicted in  FIG. 3  taken along line  4 - 4 . 
         FIG. 5  is a cross-sectional view through the airfoil shown in  FIG. 4  taken along line  5 - 5 . 
     
    
    
     The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible. 
     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 augmenter 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  15 , 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  57  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  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 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 (‘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). 
     Referring to  FIG. 2 , a portion of an engine section is shown, for example, a turbine section. It should be understood, however, that disclosed section also may be provided in a compressor section. 
     The section includes a fixed stage  60  that provides a circumferential array of vanes  63  arranged axially adjacent to a rotating stage  62 . In the example, the vanes  63  include an outer diameter portion  64  having hooks  65  that support the array of vanes  63  with respect to a case structure. An airfoil  68  extends radially from the outer platform  64  to an inner diameter portion or platform  66 . It should be understood that the disclosed vane arrangement could be used for vane structures cantilevered at the inner diameter portion of the airfoil. 
     Referring to  FIG. 3 , a cooling source  70 , such as bleed air from the compressor section, provides a cooling fluid to a baffle  72  arranged within a cooling cavity of the stator vane  63 . In the example, the cooling fluid flows into the baffle  72  through the outer platform  64 . Cooling fluid exits the baffle  72  through the inner platform  66  and flows to a component  102 . In the example, the component is a downstream airfoil. 
     Since the cooling fluid to the stator vane  63  is used to provide cooling fluid to another component, a very low flow may be provided to the baffle  72 , resulting in low Reynolds number. In this disclosure, a low Reynolds number corresponds to laminar or near-laminar flow. In one example, the Reynolds number is less than 4000. In another example, the Reynolds number is less than 1500. 
     Referring to  FIG. 4 , an exterior wall  82  provides pressure and suction sides  78 ,  80  that are joined at leading and trailing edges  74 ,  76 . The exterior wall  82  provides a cooling cavity  84  within which the baffle  72  is arranged. A perimeter cavity  86  is provided between the baffle  72  and the exterior wall  82 . 
     One or more radially extending ribs  90  are provided between and connect the pressure and suction sides  78 ,  80 . The ribs  90  separate a trailing edge cooling cavity  88  from the perimeter cavity  86 . In one example, holes  91  may be provided in the ribs  90  to provide cooling fluid from the perimeter cavity  86  into the trailing edge cooling cavity  88 , as shown in  FIG. 5 . Fluid exits the trailing edge  76  as is known. 
     An impingement cooling arrangement  92  is provided to cool the leading edge  74 . In the example, a portion of the baffle  72  includes impingement cooling holes  94  that provide impingement cooling fluid to an interior or backside of the exterior wall  82  at the leading edge  74 . 
     In one example, the baffle  72  is provided by sheet steel, for example, a single sheet, and includes an outer contour generally free of protrusions. The outer contour is provided by plastic deformation, as opposed to, for example, casting. The cooling holes, such as the impingement cooling holes  94 , are provided in the baffle  72  using at least one of drilling, laser drilling, or electro discharge machining. 
     The exterior wall  82  includes an interior surface  98  from which multiple pin fins extend to a terminal end. The terminal end supports the baffle  72 . In one example, the pin fins  96  are arranged in rows and radially spaced from one another, as best shown in  FIG. 5 . If the trips touch the baffle the flow can be blocked. Instead, with pin-fins the flow will go around not affecting the vane coolant flow rate. The pin fins  96  are integrally formed with the exterior wall, which may be formed from a nickel alloy. In one example, the pin fins  96  provide the sole support for the baffle  72  in the perimeter cavity  86 . 
     The perimeter cavity  86  circumscribes the baffle  72 . The region provided within the perimeter cavity  86  provides a Nusselt number of less than 40. In one example, the region is free of trip strips. 
     The disclosed vane and baffle arrangement provides improved convective cooling at very low Reynolds numbers as compared to trip strips. The disclosed configuration replaces trip strips with pin-fins to eliminate heat transfer decay at low Reynolds numbers. With trip strips under laminar flow, heat transfer decay is observed at the beginning of the passage and prior to reach fully developed flow. Moreover, heat transfer decay depends on the passage distance and will result in regions with improper convective cooling. Otherwise, pin-fins heat transfer coefficients are uniform at low Reynolds numbers, eliminating concern of low convective cooling in trip strips prior to reach the fully developed flow. 
     In addition, the simple design will reduce scrap rate and cost when manufacturing small airfoils. For small applications, too complicated cooling schemes are more prone to scrap due to tight manufacturing tolerances. 
     It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present invention. 
     Although the different examples have specific components shown in the illustrations, embodiments of this invention 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. 
     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 the claims. For that reason, the following claims should be studied to determine their true scope and content.