Patent Publication Number: US-2019170001-A1

Title: Impingement cooling of a blade platform

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is the US National Stage of International Application No. PCT/EP2017/067938 filed Jul. 14, 2017, and claims the benefit thereof. The International Application claims the benefit of European Application No. EP16179848 filed Jul. 18, 2016. All of the applications are incorporated by reference herein in their entirety. 
    
    
     FIELD OF INVENTION 
     The present invention relates to turbomachine components having an aerofoil, and more particularly to cooling of platform of a turbomachine component having an aerofoil, particularly a vane platform or a blade platform, in gas turbine engines. 
     BACKGROUND OF INVENTION 
     To effectively use cooling air for cooling of gas turbine components is a constant challenge and an important area of interest in gas turbine engine designs. For example, for cooling different parts of a turbomachine component having an aerofoil, such as a vane or a blade, conventional design uses various ways including film cooling and circulation of cooling fluid through different parts of the vane or the blade. However, the conventional designs are inefficient in effectively cooling all parts of the vane or the blade of the turbomachine, for example the conventional designs are inept at cooling certain parts of the platform of the vanes and/or the blades. 
     A turbine vane generally includes an inner platform and an outer platform, whereas a turbine blade usually has only one platform and may optionally have a shroud. When installed in a gas turbine engine, the inner platform of the turbine vane is usually connected to or fixed to a stationary turbine component positioned towards the rotational axis of the turbine such as a turbine vane carrier ring or a stator. Several turbine vanes may be fixed to a given turbine vane carrier ring. Similarly, the outer platform of the turbine vane is fixed to another stationary component of the turbine towards an outer casing of the turbine. Similarly, the platform of the turbine blade is fixed to rotating disks or discs mounted on a main shaft of the turbine. Several turbine blades are fixed to a given rotating disc. To be arranged properly around a given turbine vane carrier ring or a given rotating disc, the platforms of the turbine vanes or the turbine blades are usually axially extending beyond the region of the platform required to support the aerofoil and thus forming platform overhangs next to the leading edge and/or the trailing edge of the aerofoil. Such platform overhangs are prominently present in guide vanes of a gas turbine. Usually, in a gas turbine, any platform in a turbomachine component having an aerofoil has one or more platform overhangs. 
     U.S. Pat. No. 4,573,865 discloses a multiple-impingement cooled structure, such as for use as a turbine shroud assembly. The structure includes a plurality of baffles which define with an element to be cooled, such as a shroud, a plurality of cavities. Impingement cooling air is directed through holes in one of the baffles to impinge upon only the portion of the shroud in a first cavity. That cooling air is then directed to impinge again upon the portion of the shroud in a second cavity. 
     In the present description the turbine vane of a gas turbine has been used as an example of a turbomachine component having an aerofoil, however, it may be noted that for the purposes of the present technique, examples of the turbomachine component having an aerofoil also include the blade of a gas turbine. In the conventional design certain regions of the platforms of such turbomachine component having an aerofoil, hereinafter also referred to as the vane or the turbomachine component, are cooled, for example the region of the platform that is directly covered by the aerofoil has cavities through which cooling fluid flows into the aerofoil and thus the region of the platform bordering the cavity is cooled by the flow of the cooling fluid. However, the platform overhangs adjacent to the region of the platform directly beneath or above the aerofoil are not subjected to efficient cooling and thus prone to failure under the high operational temperatures and corroding effects of the hot gases coming from the combustor section when the turbine is operated. Thus there is a need to provide a technique to cool the platform overhangs, particularly side of the platform overhang that are in or towards hot gas path in the gas turbine. 
     SUMMARY OF INVENTION 
     Thus an object of the present disclosure is to provide a technique wherein the platform overhangs are cooled efficiently. It is desirable to cool side of the platform overhangs that are in or towards the hot gas path in the gas turbine. 
     The above objects are achieved by a turbomachine component and an array of turbomachine components of the present technique. Advantageous embodiments of the present technique are provided in dependent claims. Features of independent claims may be combined with features of dependent claims, and features of dependent claims can be combined together. 
     In an aspect of the present technique, a turbomachine component, particularly a blade or a vane for a gas turbine engine, is presented. The turbomachine component includes an aerofoil and a first platform. The first platform extends both circumferentially and axially. The aerofoil has a pressure side and a suction side that meet at a trailing edge and a leading edge. The first platform includes an aerofoil side wherefrom the aerofoil extends radially, an opposite side of the aerofoil side, and a first-platform cavity positioned in a first overhang region of the first platform. The first-platform cavity extends within the first platform and includes an aerofoil-side cavity wall along the aerofoil side and a plurality of impingement plates. The first platform cavity extends circumferentially and axially. The impingement plates are arranged successively in an axial direction within the first-platform cavity. Each impingement plate includes an aerofoil-side part, a flow-input-side part and a central plate. 
     The aerofoil-side part extends towards and is connected to the aerofoil-side cavity wall of the first-platform cavity. The flow-input-side part extends towards a direction opposite to the aerofoil-side cavity wall of the first-platform cavity. The central plate is between the aerofoil-side part and the flow-input-side part, and is suspended by the aerofoil-side part and the flow-input-side part in the first-platform cavity. The central plate is suspended, extending circumferentially and axially, along the aerofoil-side cavity wall such that the impingement plate defines, within the first-platform cavity in a radial direction, an aerofoil-side segment and a flow-input-side segment corresponding to said impingement plate. The central plate has impingement holes such that cooling air entering the first-platform cavity flows within the first-platform cavity from the flow-input-side segment of one impingement plate through the impingement holes to the aerofoil-side segment of said impingement plate as impingement jets, and thus cooling the aerofoil-side cavity wall along the aerofoil side of the first platform, which in turn results in the cooling of the aerofoil side of the first platform. Subsequently, the cooling air from the aerofoil-side segment of said impingement plate flows to the flow-input-side segment of a following impingement plate. From the flow-input-side segment of the following impingement plate the cooling air flows through the impingement holes of said following impingement plate as impingement jets towards the aerofoil-side cavity wall of the first-platform cavity, thus cooling of the aerofoil side of the first platform, and therefrom to the flow-input-side segment of a subsequent following impingement plate. 
     In turbomachine component, particularly in the first-platform cavity, as a result of the serially arranged impingement plates, two pockets of air corresponding to each impingement plate are created in sections of the first-platform cavity corresponding to each of the serially arranged impingement plates, namely the flow-input-side segment and the aerofoil-side segment. The flow-input-side segment and the aerofoil-side segment are in fluid communication through the impingement holes of the impingement plate creating the flow-input-side and the aerofoil-side segments. As a net result of all the impingement plates, a series of flow-input-side segments and aerofoil-side segments are created i.e. for example a flow-input-side segment of a first impingement plate fluidly connected to an aerofoil-side segment of the first impingement plate which in turn is fluidly connected to a flow-input-side segment of a second impingement plate which in turn is fluidly connected to an aerofoil-side segment of the second impingement plate which in turn is fluidly connected to a flow-input-side segment of a third impingement plate and so on and so forth. As an effect of the flow of the cooling air serially flowing through the impingement plates so arranged in the first-platform cavity buildup of strong cross flow with respect to impingement jets corresponding to a given impingement plate is minimized and thus the impingement jets are able to reach the aerofoil-side cavity wall of the first-platform cavity and provide effective cooling to the aerofoil side within the first overhang region of the first platform. Furthermore, sizes of the impingement holes can be controlled individually for different impingement plates and thus parameters of the impingement jets produced by different impingement plates, such as velocity of the impingement jets, can be controlled and thereby different degrees of cooling can be achieved locally for different impingement plates. 
     Moreover, since all the cooling air passes through the impingement holes of every impingement plate, individually and serially, the entire volume of the cooling air is used to serially cool each of the different sections of the aerofoil side within the first overhang region of the first platform created by the different impingement plates, and thus less cooling air is required to cool the aerofoil side within the first overhang region of the first platform. 
     In an embodiment of the turbomachine component, the first-platform cavity includes an opposite-side cavity wall along the opposite side of the first platform and the flow-input-side part of the impingement plate arranged within the first-platform cavity is connected to the opposite-side cavity wall. 
     In another embodiment of the turbomachine component, the first platform includes an additional first-platform cavity positioned in a second overhang region of the first platform. The additional first-platform cavity extends circumferentially and axially within the first platform and includes an aerofoil-side cavity wall along the aerofoil side and a plurality of impingement plates arranged similarly as the impingement plates are arranged in the first-platform cavity. Thus cooling is provided to second overhang region of the first platform. 
     In another embodiment of the turbomachine component, the additional first-platform cavity includes an opposite-side cavity wall along the opposite side of the first platform and the flow-input-side part of each of the impingement plates arranged within the additional first-platform cavity is connected to the opposite-side cavity wall. 
     In another embodiment of the turbomachine component, the first overhang region of the first platform is downstream of the trailing edge when viewed from the leading edge towards the trailing edge, and optionally the second overhang region of the first platform is upstream of the leading edge. In another embodiment of the turbomachine component, the first overhang region of the first platform is downstream of the leading edge when viewed from the trailing edge towards the leading edge, and optionally the second overhang region of the first platform is upstream of the leading edge. 
     In another embodiment of the turbomachine component, such as when the turbomachine component is a turbine vane, the turbomachine component includes a second platform. The second platform extends circumferentially and axially. The second platform includes an aerofoil side whereto the radially extending aerofoil extends, an opposite side of the aerofoil side, and a second-platform cavity positioned in a first overhang region of the second platform. The second-platform cavity extends circumferentially and axially within the second platform and includes an aerofoil-side cavity wall along the aerofoil side, and a plurality of impingement plates arranged similarly as the impingement plates are arranged in the first-platform cavity of the first platform. Thus cooling is provided to the second platform, for example the outer platform of a turbine vane. 
     In another embodiment of the turbomachine component, the second-platform cavity includes an opposite-side cavity wall along the opposite side of the second platform and the flow-input-side part of the impingement plate arranged within the second-platform cavity is connected to the opposite-side cavity wall. 
     In another embodiment of the turbomachine component, the second platform includes an additional second-platform cavity positioned in a second overhang region of the second platform. The additional second-platform cavity extends circumferentially and axially within the second platform and includes an aerofoil-side cavity wall along the aerofoil side and a plurality of impingement plates arranged similarly as the impingement plates are arranged in the second-platform cavity. 
     In another embodiment of the turbomachine component, the additional second-platform cavity includes an opposite-side cavity wall along the opposite side of the second platform and the flow-input-side part of each of the impingement plates arranged within the additional second-platform cavity is connected to the opposite-side cavity wall. 
     In another embodiment of the turbomachine component, the first overhang region of the second platform is downstream of the trailing edge when viewed from the leading edge towards the trailing edge, and optionally the second overhang region of the second platform is upstream of the leading edge. 
     In another embodiment of the turbomachine component, the first overhang region of the second platform is downstream of the leading edge when viewed from the trailing edge towards the leading edge, and optionally the second overhang region of the second platform is upstream of the leading edge. 
     Another aspect of the present technique presents an array of turbomachine components, such as turbine vanes or turbine blades for a gas turbine. The array includes a plurality of turbomachine components having aerofoils and a turbomachine components carrying ring. Each of the turbomachine components having aerofoils is circumferentially arranged on the turbomachine components carrying ring. The plurality of turbomachine components having aerofoils includes at least one turbomachine component according to the aspect of the present technique presented hereinabove. 
     In an embodiment of the array, the turbomachine components having aerofoils are blades for the gas turbine engine and the turbomachine components carrying ring is a rotor disc for the gas turbine engine. 
     In another embodiment of the array, the turbomachine components having aerofoils are vanes of the gas turbine engine and the turbomachine components carrying ring is a vane carrier ring of the gas turbine engine. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The above mentioned attributes and other features and advantages of the present technique and the manner of attaining them will become more apparent and the present technique itself will be better understood by reference to the following description of embodiments of the present technique taken in conjunction with the accompanying drawings, wherein: 
         FIG. 1  shows part of an exemplary turbine engine in a sectional view and in which an exemplary embodiment of a turbomachine component of the present technique is to be incorporated; 
         FIG. 2  schematically illustrates an exemplary embodiment of a segment of the turbine engine of  FIG. 1  in a sectional view and in which an exemplary embodiment of the turbomachine component of the present technique is to be incorporated; 
         FIG. 3  schematically illustrates an exemplary embodiment of a segment of the turbine engine of  FIG. 2  in a sectional view and in which an exemplary embodiment of the turbomachine component of the present technique is incorporated; 
         FIG. 4  schematically illustrates another exemplary embodiment of the turbomachine component with a first-platform cavity according to the present technique; 
         FIG. 5  schematically illustrates another exemplary embodiment of the turbomachine component with an additional first-platform cavity according to the present technique; 
         FIG. 6  schematically illustrates another exemplary embodiment of the turbomachine component with the first-platform cavity having another shape as compared to the first-platform cavity of  FIG. 4 ; 
         FIG. 7  schematically illustrates another exemplary embodiment of the turbomachine component with the first-platform cavity having another shape as compared to the first-platform cavity of  FIG. 6 ; 
         FIG. 8  schematically illustrates a cross-sectional view of an exemplary embodiment of a first platform of the turbomachine component when viewed in a radial direction; 
         FIG. 9  schematically illustrates a cross-sectional view of another exemplary embodiment of the first platform of the turbomachine component when viewed in the radial direction; 
         FIG. 10  schematically illustrates another exemplary embodiment of the turbomachine component with a second-platform cavity according to the present technique; 
         FIG. 11  schematically illustrates another exemplary embodiment of the turbomachine component with an additional second-platform cavity according to the present technique; 
         FIG. 12  schematically illustrates a cross-sectional view of an exemplary embodiment of a second platform of the turbomachine component when viewed in the radial direction; 
         FIG. 13  schematically illustrates a cross-sectional view of another exemplary embodiment of the second platform of the turbomachine component when viewed in the radial direction; 
         FIG. 14  schematically illustrates cooling air flow within the first-platform cavity of the exemplary embodiment of the turbomachine component depicted in  FIG. 3 ; 
         FIG. 15  schematically illustrates an exemplary embodiment of an arrangement of impingement plates within the first-platform cavity of the exemplary embodiment of the turbomachine component depicted in  FIG. 3 ; 
         FIG. 16  schematically illustrates an exemplary embodiment of an impingement plate from the arrangement of impingement plates within the first-platform cavity as depicted in  FIG. 15 ; 
         FIG. 17  schematically illustrates another exemplary embodiment of the impingement plate; 
         FIG. 18  schematically illustrates arrangement of impingement plates in the second-platform cavity and cooling air flow within the second-platform cavity of the exemplary embodiment of the turbomachine component depicted in  FIG. 10 ; 
         FIG. 19  schematically illustrates an array of turbomachine components; and 
         FIG. 20  schematically illustrates the first platforms of the turbomachine components of the array; in accordance with aspects of the present technique. 
     
    
    
     DETAILED DESCRIPTION OF INVENTION 
     Hereinafter, above-mentioned and other features of the present technique are described in details. Various embodiments are described with reference to the drawing, wherein like reference numerals are used to refer to like elements throughout. In the following description, for purpose of explanation, numerous specific details are set forth in order to provide a thorough understanding of one or more embodiments. It may be noted that the illustrated embodiments are intended to explain, and not to limit the invention. It may be evident that such embodiments may be practiced without these specific details. 
       FIG. 1  shows an example of a gas turbine engine  10  in a sectional view. The gas turbine engine  10  comprises, in flow series, an inlet  12 , a compressor or compressor section  14 , a combustor section  16  and a turbine section  18  which are generally arranged in flow series and generally about and in the direction of a rotational axis  20 . The gas turbine engine  10  further comprises a shaft  22  which is rotatable about the rotational axis  20  and which extends longitudinally through the gas turbine engine  10 . The shaft  22  drivingly connects the turbine section  18  to the compressor section  14 . 
     In operation of the gas turbine engine  10 , air  24 , which is taken in through the air inlet  12  is compressed by the compressor section  14  and delivered to the combustion section or burner section  16 . The burner section  16  comprises a burner plenum  26 , one or more combustion chambers  28  extending along a longitudinal axis  35  and at least one burner  30  fixed to each combustion chamber  28 . The combustion chambers  28  and the burners  30  are located inside the burner plenum  26 . The compressed air passing through the compressor  14  enters a diffuser  32  and is discharged from the diffuser  32  into the burner plenum  26  from where a portion of the air enters the burner  30  and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas  34  or working gas from the combustion is channelled through the combustion chamber  28  to the turbine section  18  via a transition duct  17 . An inner surface  55  of the transition duct  17  defines a part of the hot gas path. 
     This exemplary gas turbine engine  10  has a cannular combustor section arrangement  16 , which is constituted by an annular array of combustor cans  19  each having the burner  30  and the combustion chamber  28 , the transition duct  17  has a generally circular inlet that interfaces with the combustor chamber  28  and an outlet in the form of an annular segment. An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine  18 . 
     The turbine section  18  comprises a number of blade carrying discs  36  attached to the shaft  22 . In the present example, two discs  36  each carry an annular array of turbine blades  38 . However, the number of blade carrying discs could be different, i.e. only one disc or more than two discs. In addition, guiding vanes  40 , which are fixed to a stator  42  of the gas turbine engine  10 , are disposed between the stages of annular arrays of turbine blades  38 . Between the exit of the combustion chamber  28  and the leading turbine blades  38  inlet guiding vanes  44  are provided and turn the flow of working gas onto the turbine blades  38 . 
     The combustion gas  34  from the combustion chamber  28  enters the turbine section  18  and drives the turbine blades  38  which in turn rotate the shaft  22 . The guiding vanes  40 ,  44 , hereinafter also referred to as the vanes  40 , 44 , serve to optimise the angle of the combustion or working gas  34  on the turbine blades  38 . 
     The turbine section  18  drives the compressor section  14 . The compressor section  14  comprises an axial series of vane stages  46  and rotor blade stages  48 . The rotor blade stages  48  comprise a rotor disc supporting an annular array of blades. The compressor section  14  also comprises a casing  50  that surrounds the rotor stages and supports the vane stages  48 . The guide vane stages include an annular array of radially extending vanes that are mounted to the casing  50 . The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point. Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions. 
     The casing  50  defines a radially outer surface  52  of the passage  56  of the compressor  14 . A radially inner surface  54  of the passage  56  is at least partly defined by a rotor drum  53  of the rotor which is partly defined by the annular array of blades  48 . 
     The present technique is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine. However, it should be appreciated that the present technique is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications. Furthermore, the cannular combustor section arrangement  16  is also used for exemplary purposes and it should be appreciated that the present technique is equally applicable to annular type and can type combustion chambers. 
     The terms upstream and downstream refer to the flow direction of the airflow and/or working gas flow  34  through the engine unless otherwise stated. The terms forward and rearward refer to the general flow of gas through the engine. The terms axial, axially, axial direction, radial, radially, radial direction, circumferential, circumferentially and circumferential direction are made with reference to the rotational axis  20  of the engine, unless otherwise stated. The phrase a first element “along” a second element, and like phrases, means the first element runs or extends or is arranged in the same directions as the second element, i.e. for example to explain further, if the second element is a surface or a side and extends in x-z coordinates in Cartesian coordinate system then the first element “along” the second element means the first element also extends in x-z coordinate albeit the first element may be removed by a distance from the second element in x coordinate and/or in z coordinate. Simply put, the first element “along” the second element may be understood as the first element extending in such dimensions as to be parallel or substantially parallel to the second element for example the first element and the second element may form an angle between 0 degree and 20 degree. 
       FIG. 2  provides a more detailed view of a region ‘A’ of  FIG. 1  and gives an exemplary position in the turbine section  18 , including a junction of the combustor  16  and the turbine section  18 , where the present technique may be implemented. In  FIG. 2 , turbomachine components having aerofoil for example the inlet guiding vane  44 , the turbine blade  38 , and the guiding vane  40  are represented schematically in parts. In the gas turbine engine  10 , the inlet guiding vane  44  is fixed to a vane carrying ring  70  which may be part of the stator  42  and the turbine blade  38  is fixed to the blade carrying disc  36 . Hereinafter for purposes of explanation the inlet guiding vane  44  has been used but it may be appreciated by one skilled in the art of turbomachines that the present technique is also applicable to the turbine blade  38 , and the guiding vane  40 . 
     The inlet guiding vane  44 , hereinafter also referred to the vane  44 , has an aerofoil  110  extending from an inner platform  61 , arranged towards the rotational axis  20 , which in turn is adapted to be connected, or is connected when the vane  44  is installed within the gas turbine engine  10 , to the vane carrying ring  70 . The aerofoil  110  has a leading edge  58  and a trailing edge  60 . The aerofoil  110  covers a part  91  of the inner platform  61 , i.e. the part of the inner platform  61  that lies directly beneath the aerofoil  110 , however one or more other parts  62 ,  63  of the inner platform  61  extend beyond the part  91  of the inner platform  61  that lies directly beneath, or in direct contact with, the aerofoil  110  and thereby form a first overhang  62  downstream of the trailing edge  60  and a second overhang  63  upstream of the leading edge  58 . Similarly the turbine blade  38  has a platform  39  and the guiding vane  40  has an inner platform  71  and one or both of the platform  39  and the inner platform  71  may have one or more overhangs (not shown). The turbine blade  38  may have a heat shield  37  on the other end. 
     Conventionally, cooling air is fed from internal cooling channels (not shown) and through the platforms  61 ,  39 ,  71 , into the aerofoils  110  of the vane  40 , turbine blade  38  and the guiding vane  40 , for example through a space  77  beneath the platform  61  and then through part  91  into the aerofoil  110  of the vane  44 , though it has not been depicted in  FIG. 2  for sake of simplicity. 
     The vane  44  also has an outer platform  64  to which the aerofoil  110  extends. The aerofoil  110  covers a part  94  of the outer platform  64 , i.e. the part of the outer platform  64  that lies directly above, or in direct contact with, the aerofoil  110 , however one or more other parts  65 ,  66  of the outer platform  64  extend beyond the part  94  of the outer platform  64  and thereby form a first overhang  65  downstream of the trailing edge  60  and a second overhang  66  upstream of the leading edge  58 . Similarly the guiding vane  40  has an outer platform  72  and may have similar overhangs in the outer platform  72 . 
     The present technique is implemented in one or more overhangs  62 , 63 , 65 , 66  of the vane  44  or similar overhangs (not shown) of the platforms  39 ,  71 ,  72  of the turbine blade  38  and the guiding vane  40 . 
       FIG. 3  in combination with  FIGS. 4 to 9 , schematically presents an exemplary embodiment of a turbomachine component  100  according to an aspect of the present technique. The turbomachine component  100  is implemented in the one or more of the overhangs  62 , 63 , 65 , 66  of the vane  44  or similar overhangs (not shown) of the platforms  39 ,  71 ,  72  of the turbine blade  38  and the guiding vane  40  of  FIG. 2 . 
     As shown in  FIG. 4  in combination with  FIG. 3 , the turbomachine component  100 , particularly a blade or a vane for the gas turbine engine  10 , includes the aerofoil  110  and a first platform  120  extending axially and circumferentially i.e. the first platform  120  extends in an axial direction shown in  FIG. 4  represented by an axis  98  and in a circumferential direction shown in  FIG. 4  represented by an axis  97  mutually perpendicular to the axis  98  and an axis  99 , wherein the axis  99  represents a radial direction. The aerofoil  110  includes a generally concave side also called pressure side  114 , and a generally convex side also called suction side  116 . The pressure side  114  and the suction side  116  meet at a trailing edge  112  and a leading edge  118 . The first platform  120  is similar to the inner platform  61  of the vane  44  of  FIG. 2 . 
     The first platform  120  has generally two sides along the radial direction  99  i.e. an aerofoil side  122  from which the aerofoil  110  extends radially and an opposite side  124  which is positioned towards the vane carrying ring  70  or the blade carrying disc  36  i.e. towards the rotational axis  20  when the turbomachine component  100 , hereinafter also referred to as the component  100 , is installed in the gas turbine engine  10 , hereinafter also referred to as the gas turbine  10 . The component  100  includes a first-platform cavity  125  positioned in a first overhang region  128  of the first platform  120 . The first overhang region  128  may be understood as any of the overhangs  62 , 63 , 65 , 66  of the vane  44  of  FIG. 2 , although for the purposes of the present exemplary embodiment, the first overhang region  128  of  FIG. 3  is similar to the overhang  62  of  FIG. 2  i.e. when viewed from the leading edge  118  towards the trailing edge  112 , the first overhang region  128  is present downstream of the trailing edge  112 . However, in  FIG. 4  the first overhang region  128  is similar to the overhang  63  of  FIG. 2  i.e. when viewed from the trailing edge  112  towards the leading edge  118 , the first overhang region  128  is present downstream of the leading edge  118 . As shown in  FIGS. 4, 6 and 7 , the first-platform cavity  125  may have different configurations such as a rectangular cross-section as shown in  FIG. 4 , having four walls i.e. one wall along the side  122 , another wall along the side  124 , also called as the opposite-side cavity wall  127  (also shown in  FIG. 3 ) and two side walls thereinbetween; or a semi-rectangular cross-section as shown in  FIG. 6 , having three walls i.e. one wall along the side  122  and two side walls; or may just have one wall along the side  122  as shown in  FIG. 7 . 
     As shown in  FIGS. 3, 4, 6 and 7 , the first-platform cavity  125  of  FIG. 3  extends axially i.e. along the axis  98 , and circumferentially i.e. along the axis  97 , within the first platform  120  and includes an aerofoil-side cavity wall  126  along the aerofoil side  122 . Within the first-platform cavity  125  a plurality of impingement plates  80  are arranged (not shown in  FIGS. 4,6 and 7 ). The impingement plates  80  are arranged successively in an axial direction i.e. along the axis  98  and, each impingement plate  80  extends along the axial direction  98  and the circumferential direction  97  within the first-platform cavity  125 . The cooling air or any other cooling fluid is fed into the first-platform cavity  125  through a cooling fluid channel  75  that in turns receives the cooling air or the other cooling fluid from the cooling passage  77  as shown in  FIG. 3 . The structure of the impingement plates  80  and the flow of the cooling air through the impingement plates  80  has been explained hereinafter later particularly with reference to  FIG. 3  and  FIGS. 14 to 18 . 
     As shown in  FIG. 5 , the first platform  120  may also include an additional first-platform cavity  135  positioned in a second overhang region  129  of the first platform  120 . The second overhang region  129  of the first platform  120  may be understood as the second overhang  63  of the inner platform  61  of the vane  44  as shown in  FIG. 2 . As shown in  FIG. 5  in combination with  FIG. 4 , the second overhang region  129  is present downstream of the trailing edge  112 , as shown in  FIG. 5 , when the first overhang region  128  is present upstream of the leading edge  118 , as shown in  FIG. 4 . In other words, there may be only one platform cavity  125  in the first platform  120  and the platform cavity  125  may be present either downstream of the trailing edge  112  or upstream of the leading edge when viewed from the leading edge  118  towards the trailing edge  112 , or may have two cavities  125 ,  135  whereby one is present downstream of the trailing edge  112  and other is present upstream of the leading edge  118  when viewed from the leading edge  118  towards the trailing edge  112 . The additional first-platform cavity  135  extends circumferentially and axially within the first platform  120  and includes an aerofoil-side cavity wall  136  along the aerofoil side  122  and a plurality of impingement plates  80  arranged similarly as the impingement plates  80  are arranged in the first-platform cavity  125 . The additional first-platform cavity  135  may include an opposite-side cavity wall  137  along the opposite side  124  of the first platform  120 . 
       FIGS. 8 and 9  schematically represent the positions of the first-platform cavity  125  and the additional first-platform cavity  135  with respect to the aerofoil  110 . As depicted in  FIG. 8 , in an exemplary embodiment of the component  100 , the first overhang region  128  of the first platform  120  wherein the first-platform cavity  125  is present is downstream of the trailing edge  112  when viewed from the leading edge  118  in direction of the trailing edge  112 , and the second overhang region  129  of the first platform  120  wherein the additional first-platform cavity  135  is located, when present, is upstream of the leading edge  118 , when viewed from the leading edge  118  in direction of the trailing edge  112 . In an alternate embodiment of the component  100 , as depicted in  FIG. 9 , the first overhang region  128  of the first platform  120  wherein the first-platform cavity  125  is present is upstream of the leading edge  118  when viewed from the leading edge  118  in direction of the trailing edge  112 , and the second overhang region  129  of the first platform  120  wherein the additional first-platform cavity  135  is located, when present, is downstream of the trailing edge  112 , when viewed from the leading edge  118  in direction of the trailing edge  112 . 
     As shown in  FIG. 10 , the turbomachine component  100  may also include a circumferentially and axially extending second platform  140 . The second platform  140  includes an aerofoil side  142  whereto the radially extending aerofoil  110  extends, an opposite side  144  of the aerofoil side  142 , and a second-platform cavity  145  positioned in a first overhang region  148  of the second platform  140 . The first overhang region  148  of the second platform  140  may be understood as the first overhang  65  of the outer platform  64  of the vane  44  as shown in  FIG. 2 . The second-platform cavity  145  extends axially and circumferentially within the second platform  140  and includes an aerofoil-side cavity wall  146  along the aerofoil side  142 , and a plurality of impingement plates  80  arranged similarly as the impingement plates  80  are arranged in the first-platform cavity  125  of the first platform  120 . 
     As shown in  FIG. 11 , the second platform  140  may also include an additional second-platform cavity  155  positioned in a second overhang region  149  of the second platform  140 . The second overhang region  149  of the second platform  140  may be understood as the second overhang  66  of the outer platform  64  of the vane  44  as shown in  FIG. 2 . The additional second-platform cavity  155  extends circumferentially and axially within the second platform  140  and includes an aerofoil-side cavity wall  156  along the aerofoil side  142  and a plurality of impingement plates  80  arranged similarly as the impingement plates  80  are arranged in the first-platform cavity  125 . The additional second-platform cavity  155  may include an opposite-side cavity wall  157  along the opposite side  144  of the second platform  140 . 
       FIGS. 12 and 13  schematically represent the positions of the second-platform cavity  145  and the additional second-platform cavity  155  with respect to the aerofoil  110 . As depicted in  FIG. 12 , in an exemplary embodiment of the component  100 , the first overhang region  148  of the second platform  140  wherein the second-platform cavity  145  is present is downstream of the trailing edge  112  when viewed from the leading edge  118  in direction of the trailing edge  112 , and the second overhang region  149  of the second platform  140  wherein the additional second-platform cavity  155  is located, when present, is upstream of the leading edge  118 , when viewed from the leading edge  118  in direction of the trailing edge  112 . In an alternate embodiment of the component  100 , as depicted in  FIG. 13 , the first overhang region  148  of the second platform  140  wherein the second-platform cavity  145  is present is upstream of the leading edge  118  when viewed from the leading edge  118  in direction of the trailing edge  112 , and the second overhang region  149  of the second platform  140  wherein the additional second-platform cavity  155  is located, when present, is downstream of the trailing edge  112 , when viewed from the leading edge  118  in direction of the trailing edge  112 . 
     Hereinafter, the impingement plates  80  and the flow of the cooling air within the cavities  125 ,  135 ,  145 ,  155  is explained. The flow of the cooling air within the cavities  125 ,  135 ,  145 ,  155  has been depicted by arrows marked with reference numeral  9 . 
     As shown in  FIGS. 3 and 14 , the component  100  further includes the plurality of impingement plates  80 . The impingement plates  80  are successively arranged in the axial direction within the first-platform cavity  125 , i.e. along the axis  98  of  FIG. 4 . It may be noted that  FIGS. 3 and 14  represent cross-sectional views of the component  100  which has three impingement plates  80  serially arranged and spanning different sections of the first-platform cavity  125 . However, the three impingement plates  80  depicted in  FIGS. 3 and 14  are only for exemplary purposes and the component  100  may include impingement plates  80  which are more than or less than three. 
     As depicted in  FIGS. 15 to 17  in combination with  FIGS. 3 and 14 , each impingement plate  80  includes an aerofoil-side part  86 , a flow-input-side part  87  and a central plate  82  structurally in-between the aerofoil-side part  86  and the flow-input-side part  87 . The aerofoil-side part  86  extends towards and is connected to the aerofoil-side cavity wall  126  of the first-platform cavity  125 . The flow-input-side part  87  extends towards a direction opposite to the aerofoil-side cavity wall  126  of the first-platform cavity  125  and may be connected to the opposite-side cavity wall  127  or to a part of the vane carrying ring  70  when the opposite-side cavity wall  127  is not present. The central plate  82  is suspended by the aerofoil-side part  86  and the flow-input-side part  87  in the first-platform cavity  125  such that the central plate  82  extends along the aerofoil-side cavity wall  126 . The parts  86  and  87  may be connected or joint or fixedly attached to the wall  126  and the wall  127 , respectively, and may even be connected or positioned by interference fit. 
     As a result of attaching the part  86  to the wall  126  and the part  87  to the wall  127  or a part of the vane carrying ring  70 , the central plate  82  between the part  86  and the part  87  is suspended in the first-platform cavity  125 . Referring again to  FIGS. 14 and 15 , spatial arrangement of the central plate  82  within the first-platform cavity  125  is depicted. As a result of suspension of the central plate  82  in first-platform cavity  125 , hereinafter also referred to the cavity  125 , and connection of the part  86  and the part  87  to the wall  126  and the wall  127  or a part of the vane carrying ring  70 , respectively, each impingement plate  80  divides a section of the cavity  125  and thus defines within the cavity  125 , in the radial direction  99 , an aerofoil-side segment  6  or compartment  6  and a flow-input-side segment  7  or compartment  7 . In other words, one segment  6  and one segment  7  are created by each of the impingement plates  80  and are said to be corresponding to the impingement plate  80  that creates said segment  6  and said segment  7 . 
     The central plate  82  has impingement holes  84  as depicted in  FIGS. 16 and 17 . In the central plate  82  the impingement holes  84  are located as the array  85 . The array  85  may span entire area of the central plate  82  between the part  86  and the part  87 , as shown in  FIG. 16 . Alternatively, the array  85  may not span the entire expanse of the central plate  82  and may be limited to a portion of the central plate  82  for example a region  88  of the central plate  82 . As shown in  FIG. 14 , the cooling air entering the first-platform cavity  125  flows within the first-platform cavity  125  from the flow-input-side segment  7  of one impingement plate  80  through the impingement holes  84  to the aerofoil-side segment  6  of said impingement plate  80  as impingement jets, and then from the aerofoil-side segment  6  of said impingement plate  80  to the flow-input-side segment  7  of a following impingement plate  80 . From the flow-input-side segment  7  of the following impingement plate  80  the cooling air flows through the impingement holes  84  of said following impingement plate  80  as impingement jets towards the aerofoil-side cavity wall  146  of the first-platform cavity  125  and therefrom to the flow-input-side segment  7  of a subsequent following impingement plate  80 , and so on and so forth. 
     Similarly for the impingement plates  80  arranged in the additional first-platform cavity  135 , the aerofoil-side part  86  of the impingement plate  80  extending towards and is connected to the aerofoil-side cavity wall  136  of the additional first-platform cavity  135 ; and the flow-input-side part  87  extends towards a direction opposite to the aerofoil-side cavity wall  136  of the additional first-platform cavity  135  and is connected to the opposite-side cavity wall  137  or to a part of the vane carrying ring  70 . The impingement plates  80  are similarly arranged in the additional first-platform cavity  135  as explained for the impingement plates  80  arranged in the first-platform cavity  125  and create similarly the segments  6  and  7  and have a direction of flow of cooling air similar to that of the direction of flow of cooling air explained hereinabove for  FIG. 14 , i.e. from the segment  7  towards the segment  6  for a corresponding impingement plate  80 . 
       FIG. 18  schematically depicts the impingement plates  80  arranged in the second-platform cavity  145 . The impingement plates  80  are successively arranged in the axial direction  98  within the second-platform cavity  145 , with the aerofoil-side part  86  extending towards and connected to the aerofoil-side cavity wall  146  of the second-platform cavity  145  and the flow-input-side part  87  extending towards and connected to the opposite-side cavity wall  147  or to another stationary part of the stator  42  when the opposite-side cavity wall  147  is not present. As a result of attaching the part  86  to the wall  146  and the part  87  to the wall  147 , the central plate  82  between the part  86  and the part  87  is suspended in the second-platform cavity  145 , and as a result of suspension of the central plate  82  in second-platform cavity  145 , hereinafter also referred to the cavity  145 , and connection of the part  86  and the part  87  to the wall  146  and the wall  147 , respectively, each impingement plate  80  divides a section of the cavity  145  and thus defines within the cavity  145 , in the radial direction  99 , the segment  6  and the segment  7  similar to the segment  6 ,  7  explained hereinabove with reference to  FIGS. 3 and 14 . The flow of cooling air within the cavity  145  is similar to the flow of cooling air explained hereinabove with reference to  FIGS. 3 and 14 . 
     Similarly for the impingement plates  80  arranged in the additional second-platform cavity  155 , the aerofoil-side part  86  of the impingement plate  80  extends towards and is connected to the aerofoil-side cavity wall  156  of the additional second-platform cavity  155 ; and the flow-input-side part  87  extends towards and is connected to the opposite-side cavity wall  157 . The impingement plates  80  are similarly arranged in the additional second-platform cavity  155  as explained for the impingement plates  80  arranged in the first-platform cavity  125  and create similarly the segments  6  and  7  and have a direction of flow of cooling air similar to that of the direction of flow of cooling air explained hereinabove for  FIG. 14 , i.e. from the segment  7  towards the segment  6  for a corresponding impingement plate  80 . 
     Furthermore, referring to  FIG. 18  another embodiment of the component  100  has been explained. The component  100  includes an array  67  of turbulators  68  positioned on the aerofoil-side cavity wall  146 . The component  100  may also includes an array  67  of turbulators  68  positioned on the aerofoil-side cavity walls  136 ,  146  and  156 . The turbulators  68  increase the turbulence in the cooling air when the cooling air passes over the aerofoil-side cavity wall  126 , 136 , 146 , 156  having the turbulators  68 . The turbulators  68  depicted in  FIG. 18  are rib shaped. However, it may be noted that it is well within the scope of the present technique, that the turbulators  68  may have variety of different shapes, for example but not limited to split-rib shaped i.e. rib shapes that are split, wedge shaped, split-wedge shaped, pin fin shaped i.e. cylindrical individual protrusions, conical shaped, conical frustum shaped, spherical dome shaped, tetrahedron shaped, tetrahedral frustum shaped, pyramidal shaped, and pyramidal frustum shaped. 
       FIG. 18  depicts the turbulators  68  to be limited to a part  79  of the aerofoil-side cavity wall  146  whereas another part  78  of the aerofoil-side cavity wall  146  is free of the turbulators  68 , however, the turbulators  68  may be present over the entire expanse of the aerofoil-side cavity wall  126  within the cavity  145 . 
     In an exemplary embodiment of the component  100 , one or more of the cavities  125 , 135 , 145 , 155  is completely limited to the overhang regions  128 , 129 , 148 , 149 , respectively does not extend to the part of the platforms  120 , 140  that are directly beneath or above the aerofoil  110 . The advantage is that the cooling air directed to the aerofoil cavity through the part of the platforms  120 , 140  that are directly beneath or above the aerofoil  110  is not affected by the flow of the cooling air into the cavities  125 , 135 , 145 , 155 . The cooling air after flowing through the cavities  125 , 135 , 145 , 155  is exited in the hot gas flow path from the platform  120 , 140  directly or into a rim seal cavity  73  as depicted in  FIG. 3 . 
     Referring now to  FIGS. 19 and 20 , another aspect of the present technique is described according to which an array  300  of turbomachine components such as the turbine vanes  44 ,  40  or the turbine blades  38  is presented. The array  300  includes a plurality of turbomachine components such as the turbine vanes  44 ,  40  or the turbine blades  38  and a turbomachine components carrying ring such as the vane carrying ring  70  or the blade carrying disc  36 . The turbine vanes  44 ,  40  or the turbine blades  38  are circumferentially arranged on the vane carrying ring  70  or the blade carrying disc  36 , respectively to form a circular array around the rotational axis  20 . The plurality of the turbine vanes  44 ,  40  or the turbine blades  38  includes at least one turbomachine component  110  according to the aspect of the present technique presented hereinabove with reference to  FIGS. 2 to 17 . 
     An advantage of the present cooling arrangement is that it is compact and can provide a thin impingement cooling arrangement. In other words, the present cooling arrangement is thin or has a relatively small thickness in a direction perpendicular to the plane of the surface or wall  126  being cooled. This is particularly helpful in applications such as a blade or vane where thicknesses of parts, such as wall  126  defining a gas-washed surface, are important to minimize aerodynamic losses. The thickness or the distance between the walls  126  and  127  can be a minimum whilst maintaining sufficient impingement cooling. Thus for the platform  120  shown in  FIG. 14  the blade&#39;s aerodynamics are not compromised, there is minimal weight increase and the surrounding engine architecture is unaffected so the blade can fit into the existing space provided and can be retrofitted. 
     Another advantage of the present arrangement is that the distance from the central plate  82  to the cooled wall  126  may be an optimum distance for maximum impingement cooling effect for the impingement cooling jets. The central plate  82  may be located nearer to the wall  126 , on to which the impingement jets strike, than the wall  127 . In other examples, the central plate  82  may be located nearer to the wall  127  than the wall  126 . Thus the wall  126  may be optimally cooled. Bespoke cooling arrangements are then possible for many different applications of the present invention. For optimum cooling the impingement jets&#39; effectiveness can be dependent on the pressure of the cooling fluid, the size of the impingement hole and the distance from the impingement hole in the central plate  82  to the target surface such as the wall  126 . 
     Furthermore, each consecutive impingement plate  80  may have its central plate  82  located at a different distance from the cooled wall  126  compared to one or more of the other central plates  82 . The different distances of each central plate  82  may be dependent on a number of factors such as the pressure of the cooling air  9  immediately adjacent each central plate  82  and/or the temperature of the wall  126  and/or the temperature of the cooling air  9 . For example, and with respect to the direction of the cooling flow  9 , a first central plate  82  is a first distance away from the cooled wall  126  and a downstream central plate  82  is a second distance from the cooled wall  126 ; the second distance is smaller than the first distance. Further, each consecutive central plate  82 , after the first central plate  82 , may be closer to the cooled wall  126  than its immediately upstream neighbour. In another example, the second distance from the cooled wall  126  is greater than the first distance. Further, each consecutive central plate  82 , after the first central plate  82 , may be further from the cooled wall  126  than its immediately upstream neighbour. 
     Yet further the two walls  126 ,  127  may not be parallel and may converge or diverge such that the aerofoil-side part  86  and the flow-input-side part  87  are different lengths. Thus, where the two walls  126 ,  127  are converging or diverging the central plate  82  may be parallel to the cooled wall  126  and not parallel to the wall  127 . Alternatively, the central plate  126  may converge or diverge with respect to the cooled wall  126 . 
     It should be appreciated that two, three or more impingement plates  80  may be sequentially or consecutively located to use and reuse cooling air  9 . 
     In the present disclosure, orientation terms such as “radial”, “inner”, “outer”, “circumferential”, “beneath” “below” and the like are to be taken relative to a turbine axis i.e. the rotational axis  20 . “Inner” means radially inner, or closer to the rotational axis  20 , whereas “outer” means radially outer, or away from the rotational axis  20 . 
     While the present technique has been described in detail with reference to certain embodiments, it should be appreciated that the present technique is not limited to those precise embodiments. Rather, in view of the present disclosure which describes exemplary modes for practicing the invention, many modifications and variations would present themselves, to those skilled in the art without departing from the scope and spirit of this invention. The scope of the invention is, therefore, indicated by the following claims rather than by the foregoing description. All changes, modifications, and variations coming within the meaning and range of equivalency of the claims are to be considered within their scope.