Patent Publication Number: US-11655718-B2

Title: Blade with tip rail, cooling

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application claims benefit of the filing date of and is a divisional of U.S. patent application Ser. No. 16/793,138, filed Feb. 18, 2020, now allowed, now U.S. Pat. No. 11,035,237, which is a divisional of U.S. patent application Ser. No. 15/648,541, filed Jul. 13, 2017, now U.S. Pat. No. 10,605,098, both of which are incorporated herein in its entirety. 
    
    
     TECHNICAL FIELD 
     This disclosure relates generally to a blade, and more specifically to a blade with tip rail cooling. 
     BACKGROUND OF THE INVENTION 
     Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of combusted gases passing through the engine onto a multitude of rotating turbine blades, and in some cases, such as aircraft, generate thrust for propulsion. 
     Gas turbine engines for aircraft are designed to operate at high temperatures to maximize engine efficiency, so cooling of certain engine components, such as a high pressure turbine and a low pressure turbine, can be beneficial. Typically, cooling is accomplished by ducting cooler air from high and/or low pressure compressors to the engine components that require cooling. Temperatures in the high pressure turbine can be 1000° C. to 2000° C. and the cooling air from the compressor can be 500° C. to 700° C., enough of a difference to cool the high pressure turbine. 
     Contemporary turbine blades, as well as vanes or nozzles, generally include one or more interior cooling circuits for routing the cooling air through the blade to cool different portions of the blade, and can include dedicated cooling circuits for cooling different portions of the blade, such as the leading edge, trailing edge and tip of the blade. 
     Turbine blade tip rails in particular help to reduce aero losses and therefore increase the efficiency of turbine engines. The tip rail is subjected to a high heat loads and is difficult to effectively cool. It is frequently one of the hottest regions in the blade. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect, the disclosure relates to an airfoil comprising a body defining an interior, and extending axially between a leading edge and a trailing edge to define a chord-wise direction and radially between a root and a tip to define a span-wise direction, which terminates in a tip end wall and a tip rail extending form the tip end wall, a cooling passage formed in the interior, at least two radially-spaced cooling cavities within the tip rail, the at least two radially-spaced cooling cavities spaced with respect to an exterior surface of the tip rail, a cooling conduit fluidly coupling the cooling passage with at least one of the at least two radially-spaced cooling cavities, a connecting conduit fluidly coupling the radially-spaced cooling cavities, and a film-hole having an inlet fluidly coupled to at least one of the radially-spaced cooling cavities and an outlet provided on an exterior surface of the tip rail, wherein at least one of the connecting conduit or the cooling conduit is angled with respect to a radial axis. 
     In another aspect, the disclosure relates to a blade for a turbine engine, the blade comprising a body defining an interior, and extending axially between a leading edge and a trailing edge to define a chord-wise direction and radially between a root and a tip to define a span-wise direction, with the tip terminating in a tip end wall, and a tip rail extending from the tip end wall, at least one cooling passage formed in the interior, at least two radially-spaced cooling cavities within the tip rail, the at least two radially-spaced cooling cavities spaced with respect to an exterior surface of the tip rail, at least one cooling conduit fluidly coupling the cooling passage with at least one of the at least two radially-spaced cooling cavities, at least one connecting conduit fluidly coupling the at least two radially-spaced cooling cavities, and a set of film-holes, wherein each film-hole of the set of film-holes includes an inlet fluidly coupled to at least one of the radially-spaced cooling cavities and an outlet provided on at least a portion of an exterior surface of the tip rail, and wherein at least one of the at least one connecting conduit or the at least one cooling conduit is angled with respect to a radial axis. 
     In yet another aspect, the disclosure relates to a method of cooling a tip rail of an airfoil, with an internal cooling passage, for a turbine engine, the method comprising supplying cooling fluid to a first cooling cavity, located in the tip rail of the airfoil, through a first cooling conduit fluidly located within a wall of the airfoil and coupled to a cooling passage, and supplying cooling fluid to a second cooling cavity, located in the top of the airfoil and spaced from the first cooling cavity, through a second cooling conduit fluidly coupled to the first cooling cavity. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       In the drawings: 
         FIG.  1    is a schematic cross-sectional diagram of a portion of a turbine engine for an aircraft. 
         FIG.  2    is an isometric view of a blade for the engine of  FIG.  1    including a tip rail with cooling holes. 
         FIG.  3    is section view of the blade of  FIG.  2    taken across section illustrating cooling passages within the blade. 
         FIG.  4    is a section view of the tip rail of the blade of  FIG.  2    taken across section IV-IV illustrating cooling cavities within the tip rail. 
         FIG.  5    is the section view from  FIG.  4    illustrating a method of cooling the tip rail of the blade of  FIG.  2   . 
         FIG.  6    is a section view of the tip rail of the blade of  FIG.  2    taken across section IV-IV illustrating cooling cavities within the tip rail according to a second aspect of the disclosure described herein. 
         FIG.  7    is a section view of the tip rail of the airfoil of  FIG.  2    taken across section IV-IV illustrating cooling cavities within the tip rail according to a third aspect of the disclosure described herein. 
         FIG.  8    is a is a section side view of the tip rail of the airfoil of  FIG.  2    taken across section VIII-VIII illustrating cooling cavities within the tip rail according to a fourth aspect of the disclosure described herein. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Aspects of the disclosure described herein are directed to a tip of a blade including spaced cooling cavities having outlets formed in at least a portion of a tip rail. For purposes of illustration, the present disclosure will be described with respect to a blade for a turbine in an aircraft gas turbine engine. It will be understood, however, that aspects of the disclosure described herein are not so limited and may have general applicability within an engine, including compressors, as well as in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications. 
     As used herein, the term “forward” or “upstream” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” or “downstream” used in conjunction with “forward” or “upstream” refers to a direction toward the rear or outlet of the engine or being relatively closer to the engine outlet as compared to another component. 
     Additionally, as used herein, the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference. A “set” as used herein can include any number of a particular element, including only one. 
     All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, aft, etc.) are only used for identification purposes to aid the reader&#39;s understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of aspects of the disclosure described herein. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to one another. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto can vary. 
       FIG.  1    is a schematic cross-sectional diagram of a portion of a gas turbine engine  10  for an aircraft. The engine  10  has a longitudinally extending axis or centerline  12  extending from forward  14  to aft  16 . The engine  10  includes, in downstream serial flow relationship, a fan section  18  including a fan  20 , a compressor section  22  including a booster or low pressure (LP) compressor  24  and a high pressure (HP) compressor  26 , a combustion section  28  including a combustor  30 , a turbine section  32  including a HP turbine  34 , and a LP turbine  36 , and an exhaust section  38 . 
     The fan section  18  includes a fan casing  40  surrounding the fan  20 . The fan  20  includes a plurality of fan blades  42  disposed radially about the centerline  12  and rotatable within the fan casing  40 . The HP compressor  26 , the combustor  30 , and the HP turbine  34  form a core  44  of the engine  10 , which generates and extracts energy from combustion gases. The core  44  is surrounded by core casing  46 , which can be coupled with the fan casing  40 . 
     A HP shaft or spool  48  disposed coaxially about the centerline  12  of the engine  10  drivingly connects the HP turbine  34  to the HP compressor  26 . A LP shaft or spool  50 , which is disposed coaxially about the centerline  12  of the engine  10  within the larger diameter annular HP spool  48 , drivingly connects the LP turbine  36  to the LP compressor  24  and fan  20 . The spools  48 ,  50  are rotatable about the engine centerline and couple to a plurality of rotatable elements, which can collectively define a rotor  51 . 
     The LP compressor  24  and the HP compressor  26  respectively include a plurality of compressor stages  52 ,  54 , in which a set of compressor blades  56 ,  58  rotate relative to a corresponding set of static compressor vanes  60 ,  62  (also called a nozzle) to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage  52 ,  54 , multiple compressor blades  56 ,  58  can be provided in a ring and can extend radially outwardly relative to the centerline  12 , from a blade platform to a blade tip, while the corresponding static compressor vanes  60 ,  62  are positioned upstream of and adjacent to the rotating blades  56 ,  58 . It is noted that the number of blades, vanes, and compressor stages shown in  FIG.  1    were selected for illustrative purposes only, and that other numbers are possible. 
     The blades  56 ,  58  for a stage of the compressor can be mounted to a disk  61 , which is mounted to the corresponding one of the HP and LP spools  48 ,  50 , with each stage having its own disk  61 . The vanes  60 ,  62  for a stage of the compressor can be mounted to the core casing  46  in a circumferential arrangement. 
     The HP turbine  34  and the LP turbine  36  respectively include a plurality of turbine stages  64 ,  66 , in which a set of turbine blades  68 ,  70  are rotated relative to a corresponding set of static turbine vanes  72 ,  74  (also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage  64 ,  66 , multiple turbine blades  68 ,  70  can be provided in a ring and can extend radially outwardly relative to the centerline  12 , from a blade platform to a blade tip, while the corresponding static turbine vanes  72 ,  74  are positioned upstream of and adjacent to the rotating blades  68 ,  70 . It is noted that the number of blades, vanes, and turbine stages shown in  FIG.  1    were selected for illustrative purposes only, and that other numbers are possible. 
     The blades  68 ,  70  for a stage of the turbine can be mounted to a disk  71 , which is mounted to the corresponding one of the HP and LP spools  48 ,  50 , with each stage having a dedicated disk  71 . The vanes  72 ,  74  for a stage of the compressor can be mounted to the core casing  46  in a circumferential arrangement. 
     Complementary to the rotor portion, the stationary portions of the engine  10 , such as the static vanes  60 ,  62 ,  72 ,  74  among the compressor and turbine section  22 ,  32  are also referred to individually or collectively as a stator  63 . As such, the stator  63  can refer to the combination of non-rotating elements throughout the engine  10 . 
     In operation, the airflow exiting the fan section  18  is split such that a portion of the airflow is channeled into the LP compressor  24 , which then supplies pressurized air  76  to the HP compressor  26 , which further pressurizes the air. The pressurized air  76  from the HP compressor  26  is mixed with fuel in the combustor  30  and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine  34 , which drives the HP compressor  26 . The combustion gases are discharged into the LP turbine  36 , which extracts additional work to drive the LP compressor  24 , and are ultimately discharged from the engine  10  via the exhaust section  38 . The driving of the LP turbine  36  drives the LP spool  50  to rotate the fan  20  and the LP compressor  24 . 
     A portion of pressurized airflow  76  generated in the compressor section  22  can be drawn from the compressor section  22  as bleed air  77 . The bleed air  77  can be drawn from the pressurized airflow  76  and provided to engine components requiring cooling. The temperature of pressurized airflow  76  entering the combustor  30  is significantly increased. As such, cooling provided by the bleed air  77  is necessary for operating of such engine components in the heightened temperature environments. 
     A remaining portion of airflow  78  from the fan section  18  bypasses the LP compressor  24  and engine core  44  and exits the engine assembly  10  through a stationary vane row, and more particularly an outlet guide vane assembly  80 , comprising a plurality of airfoil guide vanes  82 , at a fan exhaust side  84 . More specifically, a circumferential row of radially extending airfoil guide vanes  82  is utilized adjacent the fan section  18  to exert some directional control of the airflow  78 . 
     The airflow  78  can be a cooling fluid used for cooling of portions, especially hot portions, of the engine  10 , and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor  30 , especially the turbine section  32 , with the HP turbine  34  being the hottest portion as it is directly downstream of the combustion section  28 . Other sources of cooling fluid can be, but are not limited to, fluid discharged from the LP compressor  24  or the HP compressor  26 . 
     Referring to  FIG.  2   , an engine component in the form of one of the turbine blades  68  includes a dovetail  86  and an airfoil  88 . The airfoil  88  includes a tip  90  and a root  92  defining a span-wise direction there between. A tip wall  94  is provided at the tip  90 , with a tip rail  96  having an exterior surface  98  and extending from the tip wall  94  to define a tip plenum  100 . The airfoil  88  further includes a leading edge  104  and a trailing edge  106  defining a chord-wise direction there between. A plurality of film-holes  112  are provided along a distal end  111  of the tip rail  96  and can also be provided in the span-wise direction along the trailing edge  106  of the airfoil  88 . Furthermore, a second set of film-holes  113  can be provided along the exterior surface  98  of the tip rail  96 . 
     The airfoil  88  mounts to the dovetail  86  by way of a platform  114  at the root  92 . The platform  114  helps to radially contain a turbine engine mainstream airflow driven by the blade  68 . The dovetail  86  can be configured to mount to a turbine rotor disk on the engine  10  to drive the blade  68 . The dovetail  86  further includes at least one inlet passage  116 , with the exemplary dovetail  86  shown as a having three inlet passages  116 . The inlet passages  116  extend through the dovetail  86  and the platform  114  to provide internal fluid communication with the airfoil  88  at corresponding passage outlets  118 . A flow of cooling fluid C, such as airflow  77  and/or airflow  78  can be provided to the airfoil  88  through the inlet passage  116 . It should be appreciated that the dovetail  86  is shown in cross-section, such that the inlet passages  116  are enclosed within the body of the dovetail  86 . 
     Referring now to  FIG.  3   , the airfoil  88  includes an outer wall  120  with a concave-shaped pressure side  122  and a convex-shaped suction side  124  joined together to define the shape of airfoil  88 . During operation, the airfoil  88  rotates in a direction such that the pressure side  122  follows the suction side  124 . Thus, as shown in  FIG.  3   , the airfoil  88  would rotate upward toward the top of the page in the direction of arrow (A). 
     An interior  130  is defined by the outer wall  120 . One or more interior walls shown as ribs  132  can divide the interior  130  into multiple cooling passages  119 . Each of the passage outlets  118  can be fluidly coupled to one or more internal cooling passages  119 . The inlet passages  116 , passage outlets  118 , internal cooling passages  119 , and film-holes  112 , can be fluidly coupled to each other and form one or more cooling circuits  121  within the airfoil  88 . 
     It should be appreciated that the interior structure of the airfoil  88  is exemplary as illustrated. The interior  130  of the airfoil  88  can be organized in a myriad of different ways, and the cooling passages  119  can include single passages extending in the span-wise direction, or can be complex cooling circuits, having multiple features such as passages, channels, inlets, outlets, ribs, pin banks, circuits, sub-circuits, film-holes, plenums, mesh, turbulators, or otherwise in non-limiting examples. Preferably, the cooling passages  119  will be in fluid communication with the inlet passages  116  of the dovetail  86 . At least one of the cooling passages  119  is in fluid communication with the film-holes  112 . 
     As can be seen more clearly in  FIG.  4   , a cross-section of the tip rail  96  taken across IV-IV of  FIG.  2    shows a series  140  of spaced cooling cavities  140   a ,  140   b ,  140   c  stacked within the tip rail  96  to define a portion of the cooling circuit  121  as described herein. In a first aspect of the disclosure discussed herein, the spaced cooling cavities are radially-spaced cooling cavities  140   a ,  140   b ,  140   c . While three radially-spaced cooling cavities  140   a ,  140   b ,  140   c  are illustrated, at least two are recommended, and it is further contemplated that four or more are possible. A plurality of connecting conduits  142   a  and  142   b  can fluidly couple the radially-spaced cooling cavities  140   a ,  140   b , and  140   c  to each other. It is contemplated that the connecting conduits  142   a  and  142   b  can each having respective intermediate inlets  143   a ,  143   b  and outlets  145   a ,  145   b.    
     A cooling conduit  144  having an inlet  146  fluidly coupled to the cooling passage  119  and an outlet  148  fluidly coupled to a first cooling cavity  140   a  connecting the series  140  of radially-spaced cooling cavities  140   a ,  140   b ,  140   c  to the cooling passage  119 . It is contemplated that the cooling conduit  144  can be multiple cooling conduits  144  such that a plurality of cooling conduits  144  are formed between the cooling passage  119  and at least one of the cooling cavities  140   a.    
     The plurality of film-holes  112  provided along the distal end  111  of the tip rail  96  can include a film-hole inlet  150   a  fluidly coupled to radially-spaced cooling cavity  140   c  and a film-hole outlet  152   a  fluidly coupled to an air source  154  surrounding the airfoil  88 . 
     The second set of film-holes  113  provided along the exterior surface  98  of the tip rail  96  can include a film-hole inlet  150   b  fluidly coupled to radially-spaced cooling cavity  140   c  and a film-hole outlet  152   b  fluidly coupled to an air source  154  surrounding the airfoil  88  within the tip plenum  100 . 
     It is further contemplated that the blade  68  can be located radially below a shroud segment  156 . The shroud segment  156  can be a plurality of shroud segments  156  circumferentially arranged around the blades  68 . 
     Turning to  FIG.  5    a method of cooling the tip rail  96  of the airfoil  88  is illustrated. Some numbers from  FIG.  4    have been removed for clarity. The method includes supplying cooling air (C) through a series of the radially-spaced cooling cavities  140   a ,  140   b , and  140   c  in the tip rail  96 . The method can further include supplying the cooling air (C) sequentially first to cooling cavity  140   a , then cooling cavity  140   b , and finally cooling cavity  140   c . It is further contemplated that the supplying can be provided to two or more cooling cavities  140   a ,  140   b , and  140   c  and is not limited to the three illustrated. It is further contemplated that the method can include impinging (I) the cooling fluid (C) onto the shroud segment  156  to cool the shroud segment  156 . Likewise, it is further contemplated that the method can include emitting the cooling fluid (C) through the outlet  152   b  into the tip plenum  100  to cool the tip rail  96 . 
     During operation it is contemplated that the tip rail  96  can contact the shroud segment  156  and over a lifetime of the blade  68  contact between the tip rail  96  and the shroud segment  156  can cause rubbing away of portions of the tip rail  96 . Forming a series  140  of radially-spaced cooling cavities  140   a ,  140   b ,  140   c  enables continued operation of the portion of the cooling circuit  121  having the radially-spaced cooling cavities  140   a  and  140   b . In other words, as the tip rail  96  wears away, cooling cavity  140   b  replaces cooling cavity  140   c  in operation. In the event cooling air (C) can no longer be provided to cooling cavity  140   c , the method simply includes providing cooling air (C) sequentially to a first cooling cavity  140   a  and then to the second cooling cavity  140   b.    
       FIG.  6    illustrates a series  240  of radially-spaced cooling cavities  240   a ,  240   b  according to another aspect of the disclosure described herein. The series  240  of radially-spaced cooling cavities  240   a ,  240   b  has similarities to the series  140 , therefore, like parts will be identified with like numerals increased by 100. It should be understood that the description of the like parts of the series  140  apply to those of the series  240 , unless otherwise noted. 
     A cooling conduit  244  can have an inlet  246  fluidly coupled to a cooling passage  219  and an outlet  248  fluidly coupled to a first cooling cavity  240   a . The inlet  246  is spaced from the outlet  248  such that the cooling conduit  244  there between is angled with respect to a radial axis (R). The angled orientation of the cooling conduit  244  can be formed such that a portion (I) of cooling air (C) impinges on an interior surface  260  of the cooling cavity  240   a  before moving into a subsequent cooling cavity  240   b.    
     At least one connecting conduit  242   a  can fluidly couple the radially-spaced cooling cavities  240   a  and  240   b  to each other. It is contemplated that the connecting conduit  242  can extend from an inlet  243  to an outlet  248  such that the cooling conduit  244  there between is angled with respect to the radial axis (R). The angled orientation of the connecting conduit  242  can be formed such that a portion (I) of cooling air (C) impinges on an interior surface  260  of the cooling cavity  240   b  before exhausting through a film-hole  212 . 
     Turning to  FIG.  7   , it is further contemplated that the spaced cavities as discussed herein are circumferentially-spaced with respect to the engine centerline  12  rather than radially-spaced. A series  340  of spaced cooling cavities  340   a ,  340   b  has similarities to the series  140 , therefore, like parts will be identified with like numerals increased by 200. It should be understood that the description of the like parts of the series  140  apply to those of the series  340 , unless otherwise noted. 
     In this aspect of the disclosure, a first cooling cavity  340   a  is circumferentially spaced from a second cooling cavity  340   b  to define the series  340 . A cooling conduit  344  can have an inlet  346  fluidly coupled to a cooling passage  319  and an outlet  348  fluidly coupled to the first cooling cavity  340   a . A connecting conduit  342 , illustrated as three connecting conduits  342   a ,  342   b ,  342   c , can each extend from an inlet  343   a ,  343   b ,  343   c  fluidly coupled to the first cooling cavity  340   a  to an outlet  345   a ,  345   b ,  345   c  fluidly coupled to the second cooling cavity  340   b.    
     At least one film-hole  312  can be along an outer wall  320  of tip rail  296  and can include a film-hole inlet  350  fluidly coupled to the second cooling cavity  340   b  and a film-hole outlet  352  fluidly coupled to an air source  354  surrounding the airfoil  88 . 
     The method as discussed herein can further include impinging (I) cooling fluid (C) both radially and circumferentially to maximize the amount of heat removed by the cooling fluid (C). In this aspect of the disclosure, it is further contemplated that the method can also include emitting the cooling fluid (C) through the at least one film-hole  312  along the outer wall  320 . 
     In yet another aspect of the disclosure discussed herein as depicted in  FIG.  8    a side cross-section taken across VIII-VIII of  FIG.  2    of the tip rail  96  with at least two radially-spaced cooling cavities  440   a ,  440   b . The first set of cavities  440   a  can include at least two first and second axially-spaced cavities  441   a ,  441   b  and the second set of cavities  440   b , radially-spaced from the first set of cavities  440   a , can include third and fourth axially-spaced cavities  441   c ,  441   d . The first and second axially-spaced cavities  441   a ,  441   b  are radially stacked below the third and fourth axially-spaced cavities  441   c ,  441   d  to define a series  440  of radially-spaced cooling cavities  440   a ,  440   b . The series  440  has similarities to the series  140  therefore, like parts will be identified with like numerals increased by 300. 
     The first set of radially-spaced cavities  440   a  can include a cooling conduit  444  having an inlet  446  fluidly coupled to a cooling passage  419  and an outlet  448  fluidly coupled to a first axially-spaced cavity  441   a . A first connecting conduit  442   a , oriented primarily in an axial direction, can fluidly couple the first axially-spaced cavity  441   a  to the second axially-spaced cavity  441   b . A second connecting conduit  442   b , oriented primarily in a radial direction, can fluidly couple the second axially-spaced cavity  441   b  to the third axially-spaced cavity  441   c . In doing so, the second connecting conduit  442   b  couples the first set of radially-spaced cavities  440   a  to the second set of radially-spaced cavities  440   b . A third connecting conduit  442   c  can fluidly couple the third axially-spaced cavity  441   c  to the fourth axially-spaced cavity  441   d . Film-holes  412  can fluidly couple the fourth axially-spaced cavity  441   d  to an exterior air source  454 . It is further contemplated that the second set of film-holes are connected to one or more of the cavities as illustrated, by way of non-limiting example, in  FIGS.  4 ,  6 , and  7   . 
     The method as discussed herein can further include impinging (I) cooling fluid (C) both radially and axially to maximize the amount of heat removed by the cooling fluid (C). 
     While illustrated as four cavities, it should be understood that the description and orientation of the conduits, film-holes, and cavities as described herein is for illustrative purposes and not meant to be limiting. For example each set of spaced cavities could include three or more axially-spaced cavities each fluidly coupled to proximate cooling cavities. The arrangement of cooling cavities as illustrated in  FIG.  8    can be applied to the arrangements discussed regarding  FIG.  4   ,  FIG.  6    and  FIG.  7    or in any combination discussed herein. 
     Furthermore, the film-holes and second set of holes as discussed herein are not limited to straight holes. Impingement holes can be curved or can be angled, as illustrated in  FIG.  6    to cool the sides of the tip rail. 
     The tip rail of the blade is subjected to high heat loads and frequently rubs against the shroud during operation. Designs accounting for the entire rail rubbing away during operation are necessary to continue cooling in this region as rub can remove cooling features or close cooling holes, completely cutting off coolant flow from this region. As discussed herein a series of cavities spaced in the radial direction enables impingement cooling to be delivered directly to the top of the rail, where a high demand for cooling exists. Additionally, if the top cavity rubs away, there is one below it to take its place in order to continue impingement cooling. 
     In the event that the rail rubs away and the top cooling hole closes, cooling flow can still be exhausted through the second set of holes, allowing the top of the rail to still be cooled. As more of the rail rubs away, the top impingement cavity will completely open up and the one below it will take its place. In this way, a normal or near normal level of cooling effectiveness can be maintained until the rail is almost completely gone. 
     To the extent not already described, the different features and structures of the various embodiments can be used in combination with each other as desired. That one feature is not illustrated in all of the embodiments is not meant to be construed that it cannot be, but is done for brevity of description. Thus, the various features of the different embodiments can be mixed and matched as desired to form new embodiments, whether or not the new embodiments are expressly described. All combinations or permutations of features described herein are covered by this disclosure. 
     It should be appreciated that application of the disclosed design is not limited to turbine engines with fan and booster sections, but is applicable to turbojets and turbo engines as well. 
     This written description uses examples to describe aspects of the disclosure described herein, including the best mode, and also to enable any person skilled in the art to practice aspects of the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of aspects of the disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.