Patent Publication Number: US-11047241-B2

Title: Gas turbine engine airfoil having serpentine fed platform cooling passage

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to U.S. Provisional Application No. 61/879,736, which was filed on Sep. 19, 2013 and is incorporated herein by reference. 
    
    
     BACKGROUND 
     This disclosure relates to a gas turbine engine airfoil. More particularly, the disclosure relates to a cooling configuration in the airfoil. 
     Gas turbine engines typically include a compressor section, a combustor section and a turbine section. 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 are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
     Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades. 
     Many turbine blades having turns that provide a serpentine shape, which create undesired pressure losses. Some turbine blades use internally cored serpentine cavities to cool the mid-body section of the airfoil between the leading and trailing edges. The cooling flow is fed into a serpentine passage from the root of the blade. Most serpentine configurations use three to five passageways with the last passageway flowing radially outward from the root and finally terminating near the tip. If cooling air does not reach the tip of the last passage, the airfoil could develop a hot spot and burn through. Having the last passageway flow radially outward takes advantage of pumping action from the circumferential forces on the turbine blade, which ensures cooling air reaches the tip of the last passage. 
     SUMMARY 
     In one exemplary embodiment, a gas turbine engine airfoil includes a platform, and spaced apart walls that provide an exterior airfoil surface that extends radially from the platform to an end opposite the platform. A serpentine cooling passage is arranged between the walls and has a first passageway that extends from the platform toward the end and a second passageway fluidly connecting to the first passageway and extending from the end toward the platform to an end. A platform cooling passageway is fluidly connected to the end and extends transversely into the platform. A cooling hole fluidly connects the platform cooling passageway to an exterior surface. 
     In a further embodiment of the above, multiple cooling passages extend radially within the airfoil and are spaced apart from one another in a chord-wise direction. 
     In a further embodiment of any of the above, the multiple passages include a leading edge passage that is arranged near a leading edge of the exterior airfoil surface. 
     In a further embodiment of any of the above, the multiple passages include a trailing edge passage that is arranged near a trailing edge of the exterior airfoil surface. 
     In a further embodiment of any of the above, each of the multiple passages includes discrete inlets that provide cooling flow to the passage. 
     In a further embodiment of any of the above, the first and second passageways provide an up-pass passageway and a down-pass passageway that form a U-shaped cooling passage. 
     In a further embodiment of any of the above, the serpentine cooling passage includes a third passageway that is fluidly connected to the second passageway and extends from the platform toward the end. 
     In a further embodiment of any of the above, the serpentine cooling passage terminates at the end. 
     In a further embodiment of any of the above, the platform cooling passageway extends along a pressure side of the platform. 
     In a further embodiment of any of the above, the platform cooling passageway extends along a suction side of the platform. 
     In a further embodiment of any of the above, the platform cooling passageway extends along a pressure side and a suction side of the platform. 
     In a further embodiment of any of the above, multiple cooling holes fluidly connect the platform cooling passageway to the exterior surface. 
     In another exemplary embodiment, a core for a gas turbine engine airfoil includes a serpentine core portion that is configured to provide an inlet that extends to a first passageway. A second passageway is fluidly connected to the first passageway to form a U-shaped cooling passage. A platform core portion is configured to provide a platform cooling passageway that is arranged transverse and connected to the second passageway. 
     In a further embodiment of the above, the first passageway is an up-pass passageway and the second passageway is a down-pass passageway. The down-pass passageway terminates near the platform cooling passageway. 
     In a further embodiment of any of the above, the first passageway is an up-pass passageway and the second passageway is a down-pass passageway. The platform cooling passageway is generally normal to the down-pass passageway. 
     In a further embodiment of any of the above, the platform core portion extends in the opposite directions from the serpentine core portion. 
     In another exemplary embodiment, a method of cooling an airfoil comprising the steps of supplying a cooling fluid to an airfoil in a radial direction toward an end, turning the cooling fluid from the end back toward the root to a region near a platform, conveying the cooling fluid from the region to the platform and exiting the cooling fluid through a cooling hole to an exterior surface. 
     In a further embodiment of the above, the supplying step includes providing the cooling fluid through multiple discrete inlets to multiple cooling passages. 
     In a further embodiment of any of the above, the turning step includes flowing the cooling fluid along a U-shaped serpentine cooling passage. 
     In a further embodiment of any of the above, the conveying step includes conveying the cooling fluid to the platform on opposite sides of an 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. 2A  is a perspective view of the airfoil having the disclosed cooling passage. 
         FIG. 2B  is a plan view of the airfoil illustrating directional references. 
         FIG. 3  is a perspective view of an example airfoil having a serpentine cooling passage, with the cooling passages and core shown in phantom. 
         FIG. 4  is a cross-sectional view of the airfoil shown in  FIG. 2A  taken along line  4 - 4 . 
         FIG. 5  is a schematic perspective view of an example serpentine cooling passage with a platform cooling passageway. 
         FIG. 6  is a schematic view of a turbine blade illustrating a serpentine cooling passageway feeding fluid to platform cooling passageways having cooling holes. 
     
    
    
     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 augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flowpath B while the compressor section  24  drives air along a core flowpath C (as shown in  FIG. 2 ) 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 low pressure compressor  44  and a 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 high pressure compressor  52  and 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  supports one or more 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 C 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. 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 disclosed serpentine cooling passage may be used in various gas turbine engine components. For exemplary purposes, a turbine blade  64  is described. It should be understood that the cooling passage may also be used in vanes, blade outer air seals, and turbine platforms, for example. 
     Referring to  FIGS. 2A and 2B , a root  74  of each turbine blade  64  is mounted to the rotor disk. The turbine blade  64  includes a platform  76 , which provides the inner flow path, supported by the root  74 . An airfoil  78  extends in a radial direction R from the platform  76  to a tip  80 . It should be understood that the turbine blades may be integrally formed with the rotor such that the roots are eliminated. In such a configuration, the platform is provided by the outer diameter of the rotor. The airfoil  78  provides leading and trailing edges  82 ,  84 . The tip  80  is arranged adjacent to a blade outer air seal (not shown). 
     The airfoil  78  of  FIG. 2B  somewhat schematically illustrates exterior airfoil surface extending in a chord-wise direction C from a leading edge  82  to a trailing edge  84 . The airfoil  78  is provided between pressure (typically concave) and suction (typically convex) wall  86 ,  88  in an airfoil thickness direction T, which is generally perpendicular to the chord-wise direction C. Multiple turbine blades  64  are arranged circumferentially in a circumferential direction A. The airfoil  78  extends from the platform  76  in the radial direction R, or spanwise, to the tip  80 . 
     The airfoil  78  includes multiple cooling passages  90  provided between the pressure and suction walls  86 ,  88 . The exterior airfoil surface may include multiple film cooling holes (not shown) in fluid communication with the cooling passage  90 . Flow through the cooling passage  90  illustrated in  FIG. 2A  is shown in more detail in  FIG. 3 . 
     Referring to  FIG. 3 , a core  112  is shown in phantom within the turbine blade  64 . The core  112  produces correspondingly shaped passages within the turbine blade using known casting techniques. Alternatively, the airfoil  64  may be constructed using an additive manufacturing technique in which the cooling passages are formed while constructing the blade layer-by-layer. 
     The turbine blade  64  includes multiple cooling passages  90 A,  90 B,  90 C. The cooling passage  90 A corresponds to a leading edge cooling passage, and the cooling passage  90 C corresponds to a trailing edge cooling passage. In one type of cooling configuration, a serpentine cooling passage  90 B is provided in the mid-body section of the airfoil  78  between the leading and trailing edge cooling passages  90 A,  90 C, as shown in  FIG. 4 . In one example, each of the cooling passages  90 A,  90 B,  90 C is fed by discrete inlets  92 A,  92 B,  92 C, respectively, which are joined at a sprue ( FIG. 3 ) for handling during the casting process. 
     Typically, a serpentine cooling passage  90 B has at least one up-pass connected to at least one down-pass interconnected to one another by a bend to provide a U-shaped passage, as shown in  FIGS. 3 and 5 . In the example shown, the cooling passage  90 B includes a first passageway  94  extending radially outward from the inlet  92 B toward an end, in the example, the tip  80 . A second passageway  96  is interconnected to the first passageway  94  at a first bend  100  and extends radially downward away from the tip  80  toward the platform  76 . In one example, a third passageway  98  is interconnected to the second passageway  96  at a second bend  102  and extends radially upward from the platform  76  toward the tip where the passageway terminates. 
     The mid-body of the airfoil  78  may be susceptible to developing a hot spot if the pumping action of the fluid is ineffective. Thus, the disclosed cooling configuration provides a cooling flow exit at a location on the turbine blade  64  with a low dump pressure, which ensures that the fluid continues to flow through the serpentine cooling passage  90 B. To this end, a platform passageway  104  is arranged within the platform  76  and is fluidly interconnected to the second passageway  96  at an end  106 , which is generally arranged near the second bend  102  in the example. The platform passageway  104  is generally normal to the second passageway  94 . At least one cooling hole  108  fully connects the platform passageway  104  to an exterior surface  110  to provide an exit for the cooling flow near the inner gas flow path, which has a relatively low pressure as compared to the fluid pressure at the inlet  92 B. The cooling holes  108  may be any suitable shape, for example, slots, circular, non-circular, linear, non-linear and others. The exterior surface  110  may be provided in on the platform and or blade necks, for example. 
     The core  112  includes a serpentine core  114  providing the first, second and third passageways  94 ,  96 ,  98 . The core  112  also includes a platform core  116  corresponding to the platform passageway  104 . The serpentine core portion  114  includes an up-pass portion  118  and a down-pass portion  120  that respectively provide the first and second passageways  94 ,  96 . The platform core portion  116  is interconnected to the down pass portion  120  at an intersection  122 . 
     In one an example, the platform passageway  104  is generally perpendicular to the second passageway  96 . The first, second and third passageways  94 ,  96 ,  98  extend in a radial direction and the platform passageway  104  extends in the circumferential direction A. The serpentine cooling passage  90 B may be provided by any number of passes. For example, two passes are shown in  FIG. 5  and three passes are shown in  FIG. 3 . 
     A platform passageway may be provided on either or both of the pressure and suction side portions of the platform  76 , as shown in  FIG. 6 . 
     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 example embodiments have 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 and other reasons, the following claims should be studied to determine their true scope and content.