Patent Publication Number: US-11661853-B2

Title: Airfoil tip pocket with augmentation features

Description:
CROSS REFERENCE TO RELATED APPLICATION 
     This application is a divisional of U.S. patent application Ser. No. 16/858,955, filed on Apr. 27, 2020, which is a continuation of U.S. patent application Ser. No. 15/995,884, filed on Jun. 1, 2018, which is a continuation of U.S. patent application Ser. No. 14/704,022, filed on May 5, 2015, which claims priority to U.S. Provisional Application No. 61/994,270, filed on May 16, 2014. 
    
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This invention was made with government support under Contract No. N68335-13-C-0005, awarded by the United States Navy. The Government therefore has certain rights in this invention. 
    
    
     BACKGROUND 
     This disclosure relates to a gas turbine engine, and more particularly to a gas turbine engine component having an airfoil that includes a tip pocket. The tip pocket may employ one or more heat transfer augmentation devices. 
     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. 
     Because they are commonly exposed to hot combustion gases, many gas turbine engine components, such as blades and vanes, employ cooling circuits that channel dedicated cooling fluids for cooling the component. It can be difficult to communicate enough cooling fluid to cool airfoil tips of the components. This may lead to airfoil tip burning. 
     SUMMARY 
     A component according to an exemplary aspect of the present disclosure includes, among other things, an airfoil that includes a pressure sidewall and a suction sidewall that meet together at both a leading edge and a trailing edge. The airfoil extends to a tip. A tip pocket is formed in the tip and terminates prior to the trailing edge. A heat transfer augmentation device is formed in the tip pocket. 
     In a further non-limiting embodiment of the foregoing component, the tip pocket includes a suction side lip, a pressure side lip, a leading edge lip and a trailing edge lip that extend radially outwardly from a floor. 
     In a further non-limiting embodiment of either of the foregoing components, the heat transfer augmentation device axially extends across at least one of the suction side lip and the pressure side lip. 
     In a further non-limiting embodiment of any of the foregoing components, the heat transfer augmentation device extends from the leading edge lip to the trailing edge lip. 
     In a further non-limiting embodiment of any of the foregoing components, the heat transfer augmentation device extends radially outwardly from the floor across at least one of the suction side lip and the pressure side lip. 
     In a further non-limiting embodiment of any of the foregoing components, the heat transfer augmentation device is slanted toward either the leading edge lip or the trailing edge lip. 
     In a further non-limiting embodiment of any of the foregoing components, the heat transfer augmentation device is formed on the floor and extends between the suction side lip and the pressure side lip. 
     In a further non-limiting embodiment of any of the foregoing components, the heat transfer augmentation device is tapered. 
     In a further non-limiting embodiment of any of the foregoing components, the heat transfer augmentation device is a chevron or a trip strip. 
     In a further non-limiting embodiment of any of the foregoing components, the heat transfer augmentation device is formed on a suction side lip of the tip pocket and a second heat transfer augmentation device is formed on a pressure side lip of the tip pocket. 
     In a further non-limiting embodiment of any of the foregoing components, a third heat transfer augmentation device is formed on a floor of the tip pocket. 
     In a further non-limiting embodiment of any of the foregoing components, a plurality of cooling holes fluidly connect the tip pocket to at least one internal cooling cavity formed inside the airfoil. 
     In a further non-limiting embodiment of any of the foregoing components, the plurality of cooling holes are positioned relative to a central axis of the tip pocket, and a wall of the heat transfer augmentation device terminates prior to intersecting the central axis. 
     In a further non-limiting embodiment of any of the foregoing components, at least one of the plurality of cooling holes is angled relative to a floor of the tip pocket. 
     In a further non-limiting embodiment of any of the foregoing components, the heat transfer augmentation device is slanted at an angle relative to a floor of the tip pocket. 
     In a further non-limiting embodiment of any of the foregoing components, the heat transfer augmentation device includes a ramp that divides a floor of the tip pocket into radially offset floor portions. 
     A gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, an airfoil including a tip pocket formed at a tip of the airfoil. The tip pocket extends from a position near a leading edge of the airfoil to a position that is upstream from a trailing edge of the airfoil. A plurality of heat transfer augmentation devices are formed in the tip pocket. 
     In a further non-limiting embodiment of the foregoing gas turbine engine, the tip pocket extends from a position near the leading edge to a position near a mid-span of the airfoil. 
     In a further non-limiting embodiment of either of the foregoing gas turbine engines, the plurality of heat transfer augmentation devices are formed on a floor of the tip pocket. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, the plurality of heat transfer augmentation devices are formed on a suction side wall of the tip pocket. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, the plurality of heat transfer augmentation devices are formed on a pressure side wall of the tip pocket. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, the plurality of heat transfer augmentation devices are formed on at least two of a floor, a suction side wall and a pressure side wall of the tip pocket. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, a cooling hole fluidly connects the tip pocket to at least one internal cooling cavity formed inside the airfoil. The plurality of heat transfer augmentation devices terminate prior to intersecting a central axis of the tip pocket. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, the cooling hole is angled relative to a floor of the tip pocket. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, the plurality of heat transfer augmentation devices are slanted at an angle relative to a floor of the tip pocket. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, the plurality of heat transfer augmentation devices includes ramps that divide a floor of the tip pocket into radially offset floor portions. 
     A method of cooling a gas turbine engine component according to another exemplary aspect of the present disclosure includes, among other things, communicating a cooling fluid into a tip pocket formed at a tip of an airfoil, temporarily blocking the cooling fluid within the tip pocket with at least one heat transfer augmentation device, and expelling the cooling fluid from the tip pocket into a gas stream. 
     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. 
     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 gas turbine engine component. 
         FIG.  3    illustrates a tip of an airfoil of a gas turbine engine component. 
         FIG.  4    illustrates a cross-sectional view through section A-A of  FIG.  3   . 
         FIG.  5    illustrates a cross-sectional view through section B-B of  FIG.  3   . 
         FIG.  6    illustrates a tip of an airfoil of a gas turbine engine component according to a second embodiment of this disclosure. 
         FIG.  7    illustrates a cross-sectional view through section A-A of  FIG.  6   . 
         FIG.  8    illustrates a cross-sectional view through section B-B of  FIG.  6   . 
         FIG.  9    illustrates a tip of an airfoil of a gas turbine engine component according to another embodiment of this disclosure. 
         FIG.  10    illustrates a cross-sectional view through section A-A of  FIG.  9   . 
         FIG.  11    illustrates a cross-sectional view through section B-B of  FIG.  9   . 
         FIG.  12    illustrates a tip of an airfoil of a gas turbine engine component according to yet another embodiment of this disclosure. 
         FIG.  13    illustrates a cross-sectional view through section A-A of  FIG.  12   . 
         FIG.  14    illustrates a cross-sectional view through section B-B of  FIG.  12   . 
         FIG.  15    illustrates a tip of an airfoil of a gas turbine engine component according to yet another embodiment of this disclosure. 
         FIG.  16    illustrates a cross-sectional view through section A-A of  FIG.  15   . 
         FIG.  17    illustrates a cross-sectional view through section B-B of  FIG.  15   . 
         FIG.  18    illustrates an airfoil tip according to another embodiment of this disclosure. 
         FIG.  19    illustrates another airfoil tip. 
         FIG.  20    illustrates a tip of an airfoil of a gas turbine engine component according to yet another embodiment of this disclosure. 
         FIG.  21    illustrates a tip of an airfoil of a gas turbine engine component according to yet another embodiment of this disclosure. 
         FIG.  22    illustrates a cross-sectional view through section A-A of  FIG.  21   . 
         FIG.  23    illustrates a cross-sectional view through section B-B of  FIG.  21   . 
         FIG.  24    illustrates a tip of an airfoil of a gas turbine engine component according to yet another embodiment of this disclosure. 
     
    
    
     DETAILED DESCRIPTION 
     This disclosure relates to a gas turbine engine component having an airfoil. A tip pocket is formed at a tip of the airfoil. The tip pocket may include one or more heat transfer augmentation devices, such as trip strips, chevrons, or the like, disposed within the tip pocket. The heat transfer augmentation devices may be formed on a suction or pressure side lip of the tip pocket, a floor of the tip pocket, or any combination of locations. These and various other features are discussed in greater detail herein. 
       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 the 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 the 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 gear system  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 and turboshafts. 
     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, 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 1,150 ft/second (350.5 meters/second). 
     Each of the compressor section  24  and the turbine section  28  may include alternating rows of rotor assemblies and vane assemblies (shown schematically). 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  may either create or extract energy in the form of pressure from the core airflow as it is communicated along the core flow path C. The vanes  27  direct the core airflow to the blades  25  to either add or extract energy. 
       FIG.  2    illustrates a component  58  that can be employed for use within a gas turbine engine, such as the gas turbine engine  20  of  FIG.  1   . In one embodiment, the component  58  is a turbine blade. Although the component  58  is illustrated as turbine blade, the various features of this disclosure are applicable to other components located elsewhere within a gas turbine engine, such as vanes or compressor airfoils. Thus, some features that are particularly relevant to the illustrated turbine blade are to be considered optional and are not necessary to practice this disclosure. 
     In one embodiment, the component  58  includes a platform  60 , an airfoil  62  that extends in a first direction from the platform  60 , and a root  64  that extends in a second, opposite direction from the platform  60 . The airfoil  62  includes a leading edge  66 , a trailing edge  68 , a pressure sidewall  70  and a suction sidewall  72 . The pressure sidewall  70  and the suction sidewall  72  are spaced apart and generally meet together at both the leading edge  66  and the trailing edge  68 . 
     The airfoil  62  connects to the platform  60  at a fillet  69 . The root  64  connects to the platform  60  at buttresses  71 . The root  64  may include a neck  73  and one or more serrations  75  for securing the component  58  to a disk (not shown). 
     Although shown schematically in  FIG.  2   , the component  58  may include multiple internal cooling cavities  74 A,  74 B and  74 C. The internal cooling cavities  74 A,  74 B and  74 C are core formed cavities that extend inside of the airfoil  62 . The internal cooling cavities  74 A,  74 B and  74 C may form part of an internal cooling circuit  81  for cooling the component  58 . The exemplary internal cooling circuit  81  of the component  58  represents but one non-limiting example of the many potential cooling circuits that may be formed inside the component  58 . In other words, the component  58  could be cast to include various alternative cooling cavities and internal circuitry configurations within the scope of this disclosure 
     With reference to the engine  20  of  FIG.  1   , the airfoil  62  extends to a tip  77 . The tip  77  can be referred to as the outer diameter portion of the component  58 . The root  64  can be referred to as the inner diameter portion of the component  58 . 
     A tip pocket  80  may be formed in the tip  77  of the airfoil  62 . The tip pocket  80  may also be referred to as a squealer pocket. In one embodiment, the tip pocket  80  is part of the internal cooling circuit  81  of the component  58 . As discussed in greater detail below, the tip pocket  80  introduces a cooling fluid at the tip  77  of the airfoil  62  to cool the tip  77  and avoid airfoil tip burning. 
       FIGS.  3 ,  4  and  5    illustrate an exemplary tip  77  of an airfoil  62 . A tip pocket  80  is formed within the tip  77 . A cooling fluid F may be directed into the tip pocket  80  to cool the tip  77 . 
     The tip pocket  80  terminates prior to, or upstream from, the trailing edge  68  of the airfoil  62 . In one embodiment, the tip pocket  80  extends from a position near the leading edge  66  of the airfoil  62  to a position near a mid-span M of the airfoil  62  (see  FIG.  3   ). However, the axial, circumferential and radial dimensions of the tip pocket  80  may vary within the scope of this disclosure. 
     The tip pocket  80  may include a suction side lip  82 , a pressure side lip  84 , a leading edge lip  86  and a trailing edge lip  88 . The suction side lip  82 , the pressure side lip  84 , the leading edge lip  86  and the trailing edge lip  88  extend radially outwardly from a floor  90  of the tip pocket  80 . 
     The tip pocket  80  may include one or more heat transfer augmentation devices  92 . In one non-limiting embodiment, the heat transfer augmentation device  92  is a trip strip. However, other augmentation devices are also contemplated as being within the scope of this disclosure (see, for example,  FIGS.  18  and  19   ). 
     In one embodiment, the heat transfer augmentation device  92  axially extends across the suction side lip  82  of the tip pocket  80  between the leading edge lip  86  and the trailing edge lip  88 . However, other configurations are also contemplated. For example, the pressure side lip  84  could alternatively or additionally include a heat transfer augmentation device  92 . 
     As best illustrated in the cross-sectional views of  FIGS.  4  and  5   , cooling holes  94  may extend through the floor  90  of the tip pocket  80  to fluidly connect the tip pocket  80  with one or more internal cooling cavities  74  that are formed inside the airfoil  62 . The cooling holes  94  may communicate a cooling fluid F from the internal cooling cavities  74  into the tip pocket  80 . A portion of the cooling holes  94  may be angled relative to the floor  90  and the internal cooling cavities  74 . 
     The heat transfer augmentation device(s)  92  are adapted to temporarily trap the cooling fluid F inside the tip pocket  80 . For example, the heat transfer augmentation device(s)  92  may temporarily block the cooling fluid F prior to its ejection into a gas stream GS (see  FIG.  4   ). In addition, some of the cooling fluid F may ricochet back toward the floor  90  after contacting the heat transfer augmentation device(s)  92 , thereby increasing the amount of time the cooling fluid F stays within the tip pocket  80 . The heat transfer augmentation device(s)  92  also augment the heat transfer of the cooling fluid F in the tip pocket  80 . 
     This disclosure is not intended to be limited to the exact configuration of the tip pocket  80  of  FIGS.  3 - 5   . Indeed,  FIGS.  6 - 20    illustrate various other non-limiting embodiments of an airfoil tip pocket for a gas turbine engine component. 
       FIGS.  6 ,  7  and  8    illustrate another tip  177  of an airfoil  162 . In this disclosure, like reference numbers designate like elements where appropriate and reference numerals with the addition of 100 or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding original elements. A tip pocket  180  is formed at the tip  177  and includes multiple heat transfer augmentation devices  192 . 
     In this embodiment, the heat transfer augmentation devices  192  are formed on both a suction side lip  182  and a pressure side lip  184  of the tip pocket  180  and may extend radially outwardly from a floor  190  of the tip pocket  180 . In other words, the heat transfer augmentation devices  192  of this embodiment extend vertically. The heat transfer augmentation devices  192  may extend to the same height as the suction side lip  182  and the pressure side lip  184 , in one embodiment. 
     A plurality of cooling holes  194  extend through the floor  190  of the tip pocket  180 . The cooling holes  194  may be positioned along a central axis CA of the tip pocket  180  (see  FIG.  6   ) or could be offset from the central axis (see  FIGS.  21  and  24   ). Each heat transfer augmentation device  192  may include a wall  195  that faces toward the cooling holes  194 . In one embodiment, the walls  195  terminate prior to intersecting the central axis CA. Put another way, the thickness of each heat transfer augmentation device  192  does not extend beyond the cooling holes  194 . 
       FIGS.  9 ,  10 ,  11 ,  12 ,  13  and  14    illustrate another tip  277  of an airfoil  262 . A tip pocket  280  is formed at the tip  277  and includes multiple heat transfer augmentation devices  292 . In this embodiment, the heat transfer augmentation devices  292  are formed on both a suction side lip  282  and a pressure side lip  284  of the tip pocket  280  and may extend radially outwardly from a floor  290  of the tip pocket  280 . The heat transfer augmentation devices  292  may be slanted toward either a leading edge lip  286  of the tip pocket  280  (see  FIGS.  9 - 11   ) or a trailing edge lip  288  of the tip pocket  280  (see  FIGS.  12 - 14   ). In one non-limiting embodiment, the heat transfer augmentation devices  292  are slanted at about a 45° angle relative to the floor  290 . Of course, other slant angles are also contemplated. 
       FIGS.  15 ,  16  and  17    illustrate yet another tip  377  of an airfoil  362 . A tip pocket  380  is formed at the tip  377  and includes a plurality of heat transfer augmentation devices  392 . In this embodiment, the heat transfer augmentation devices  392  are formed on a floor  390  of the tip pocket  380 . The heat transfer augmentation devices  392  help mitigate aftward flow of the cooling fluid F and increase surface area of the tip pocket  380 . 
     The heat transfer augmentation devices  392  may extend from a suction side lip  382  to a pressure side lip  384  of the tip pocket  380 . That is, the heat transfer augmentation devices  392  may span an entire distance between the suction side lip  382  and the pressure side lip  384 . 
     In one embodiment, the heat transfer augmentation devices  392  are tapered. For example, as best illustrated in  FIG.  17   , the heat transfer augmentation devices  392  may taper toward a central axis CA of the tip pocket  380 . This taper may simplify the casting of the heat transfer augmentation devices  392  within the tip pocket  380 . 
       FIGS.  18  and  19    illustrate yet another tip  477  of an airfoil  462 . A tip pocket  480  is formed at the tip  477  and includes a plurality of heat transfer augmentation devices  492 . Like the embodiment of  FIGS.  15 - 17   , the heat transfer augmentation devices  492  are formed on a floor  490  of the tip pocket  480 . However, rather than a trip strip configuration, the heat transfer augmentation devices  492  of this embodiment are chevrons. The heat transfer augmentation devices  492  may point toward a leading edge lip  486  (see  FIG.  19   ) or a trailing edge lip  488  (see  FIG.  18   ) of the tip pocket  480 . 
     Yet another tip pocket  580  is illustrated by  FIG.  20   . The tip pocket  580  is disposed at a tip  577  of an airfoil  562 . In this embodiment, the tip pocket  580  includes a plurality of heat transfer augmentation devices  592  that are disposed at each of a floor  590 , a suction side lip  582  and a pressure side lip  584  of the tip pocket  580 . 
       FIGS.  21 ,  22  and  23    illustrate another tip  677  of an airfoil  662 . A tip pocket  680  is formed at the tip  677  and may include one or more heat transfer augmentation devices  692 . In one embodiment, the heat transfer augmentation device  692  helps mitigate aftward flow of the cooling fluid F and increases surface area of the tip pocket  680 . 
     The heat transfer augmentation device  692  may extend at any angle between a suction side lip  682  and a pressure side lip  684  of the tip pocket  680 . In one embodiment, the heat transfer augmentation device  692  is formed by a stepped portion of a floor  690  of the tip pocket  680 . For example, the heat transfer augmentation device  692  may include a ramp  691  that extends between a first portion  693  and a second portion  695  of the floor  690 . The first portion  693  and the second portion  695  of the floor  690  are radially offset from one another. In other words, the first portion  693  and the second portion  695  of the floor  690  extend in different planes. 
     Yet another tip pocket  780  is illustrated by  FIG.  24   . The tip pocket  780  is disposed at a tip  777  of an airfoil  762 . In this embodiment, the tip pocket  780  includes heat transfer augmentation devices  792  that extend from a suction side lip  782  and a pressure side lip  784  of the tip pocket  780 . Cooling holes  794  may extend through a floor  790  of the tip pocket  780  to fluidly connect the tip pocket  780  with an internal cooling cavity  774 . The cooling holes  794  may communicate a cooling fluid F from the internal cooling cavity  774  into the tip pocket  780 . Cooling hole portions P 1  and P 2  of the cooling holes  794  may be angled relative to the floor  790  and the internal cooling cavity  774 . In one embodiment, the angled cooling holes  794  are angled at a relatively shallow angle α. Cooling holes of such a shallow angle cannot typically be cast or drilled into the tip pocket  780 . However, these angled cooling holes  794  can be manufactured by using an additive manufacturing process. 
     The tip wall configurations of  FIGS.  3 - 24    are intended as non-limiting embodiments only. Any configuration shown by these Figures may be used in combination with a configuration of any other Figure to create a desired tip cooling effect. In all embodiments, the heat transfer augmentation devices increase the amount of time the cooling fluid is maintained within the tip pocket and augment the heat transfer of the cooling fluid. 
     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.