Patent Publication Number: US-10309228-B2

Title: Impingement insert for a gas turbine engine

Description:
FIELD OF THE TECHNOLOGY 
     The present disclosure generally relates to a gas turbine engine. More particularly, the present disclosure relates to an impingement insert for a gas turbine engine. 
     BACKGROUND 
     A gas turbine engine generally includes a compressor section, a combustion section, a turbine section, and an exhaust section. The compressor section progressively increases the pressure of a working fluid entering the gas turbine engine and supplies this compressed working fluid to the combustion section. The compressed working fluid and a fuel (e.g., natural gas) mix within the combustion section and burn in a combustion chamber to generate high pressure and high temperature combustion gases. The combustion gases flow from the combustion section into the turbine section where they expand to produce work. For example, expansion of the combustion gases in the turbine section may rotate a rotor shaft connected, e.g., to a generator to produce electricity. The combustion gases then exit the gas turbine via the exhaust section. 
     The turbine section includes one or more turbine nozzles, which direct the flow of combustion gases onto one or more turbine rotor blades. The one or more turbine rotor blades, in turn, extract kinetic energy and/or thermal energy from the combustion gases, thereby driving the rotor shaft. In general, each turbine nozzle includes an inner side wall, an outer side wall, and one or more airfoils extending between the inner and the outer side walls. Since the one or more airfoils are in direct contact with the combustion gases, it may be necessary to cool the airfoils. 
     In certain configurations, cooling air is routed through one or more inner cavities defined by the airfoils. Typically, this cooling air is compressed air bled from compressor section. Bleeding air from the compressor section, however, reduces the volume of compressed air available for combustion, thereby reducing the efficiency of the gas turbine engine. 
     BRIEF DESCRIPTION OF THE TECHNOLOGY 
     Aspects and advantages of the technology will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the technology. 
     In one aspect, the present disclosure is directed to an impingement insert for a gas turbine engine. The impingement insert includes an insert wall having an inner surface and an outer surface spaced apart from the inner surface. A nozzle extends at least one of outwardly from the outer surface of the insert wall and inwardly from the inner surface of the insert wall. The nozzle includes an outer surface and a circumferential surface. The insert wall and the nozzle collectively define a cooling passage extending from the inner surface of the insert wall to the outer surface of the nozzle. The cooling passage includes an inlet portion, a throat portion, a converging portion extending from the inlet portion to the throat portion, an outlet portion, and a diverging portion extending from the throat portion to the outlet portion. The cooling passage further includes a cross-sectional shape having a semicircular portion and a non-circular portion. 
     A further aspect of the present disclosure is directed to a gas turbine engine having a compressor section, a combustion section, a turbine section, and a gas turbine engine component. An impingement insert is positioned within the gas turbine engine component. The impingement insert includes an insert wall having an inner surface and an outer surface spaced apart from the inner surface. A nozzle extends at least one of outwardly from the outer surface of the insert wall and inwardly from the inner surface of the insert wall. The nozzle includes an outer surface and a circumferential surface. The insert wall and the nozzle collectively define a cooling passage extending from the inner surface of the insert wall to the outer surface of the nozzle. The cooling passage includes an inlet portion, a throat portion, a converging portion extending from the inlet portion to the throat portion, an outlet portion, and a diverging portion extending from the throat portion to the outlet portion. The cooling passage further includes a cross-sectional shape having a semicircular portion and a non-circular portion. 
     These and other features, aspects and advantages of the present technology will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the technology and, together with the description, serve to explain the principles of the technology. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present technology, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended FIGS., in which: 
         FIG. 1  is a schematic view of an exemplary gas turbine engine that may incorporate various embodiments disclosed herein; 
         FIG. 2  is a cross-sectional view of an exemplary turbine section that may be incorporated in the gas turbine engine shown in  FIG. 1  and may incorporate various embodiments disclosed herein; 
         FIG. 3  is a perspective view of an exemplary nozzle that may be incorporated into the turbine section shown in  FIG. 2  and may incorporate various embodiments disclosed herein; 
         FIG. 4  is a cross-sectional view of the nozzle taken generally about line  4 - 4  in  FIG. 3 , further illustrating the features thereof; 
         FIG. 5  is a perspective view of a portion of the nozzle shown in  FIGS. 3 and 4 , illustrating an impingement insert positioned therein; 
         FIG. 6  is a perspective view of the impingement insert shown in  FIG. 5 , which may incorporate various embodiments disclosed herein; 
         FIG. 7  is a partial cross-sectional view of the impingement insert taken generally about line  7 - 7  in  FIG. 6 , illustrating a nozzle and a cooling passage; 
         FIG. 8A  is a front view of the nozzle shown in  FIG. 6 , illustrating one embodiment of a cross-sectional shape of the cooling passage; 
         FIG. 8B  is a front view of the nozzle shown in  FIG. 6 , illustrating another embodiment of a cross-sectional shape of the cooling passage; 
         FIG. 9  is a partial cross-sectional view of the impingement insert similar to  FIG. 7 , illustrating cooling air flowing through the cooling passage; and 
         FIG. 10  is a partial cross-sectional view of the impingement insert similar to  FIG. 7 , illustrating another embodiment of the a nozzle. 
     
    
    
     Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present technology. 
     DETAILED DESCRIPTION OF THE TECHNOLOGY 
     Reference will now be made in detail to present embodiments of the technology, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the technology. As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. 
     Each example is provided by way of explanation of the technology, not limitation of the technology. In fact, it will be apparent to those skilled in the art that modifications and variations can be made in the present technology without departing from the scope or spirit thereof. For instance, features illustrated or described as part of one embodiment may be used on another embodiment to yield a still further embodiment. Thus, it is intended that the present technology covers such modifications and variations as come within the scope of the appended claims and their equivalents. Although an industrial or land-based gas turbine is shown and described herein, the present technology as shown and described herein is not limited to a land-based and/or industrial gas turbine unless otherwise specified in the claims. For example, the technology as described herein may be used in any type of turbine including, but not limited to, aviation gas turbines (e.g., turbofans, etc.), steam turbines, and marine gas turbines. 
     Referring now to the drawings,  FIG. 1  is a schematic of an exemplary gas turbine engine  10  as may incorporate various embodiments disclosed herein. As shown, the gas turbine engine  10  generally includes a compressor section  12  having an inlet  14  disposed at an upstream end of an axial compressor  16 . The gas turbine engine  10  further includes a combustion section  18  having one or more combustors  20  positioned downstream from the compressor  16 . The gas turbine engine  10  also includes a turbine section  22  having a turbine  24  (e.g., an expansion turbine) disposed downstream from the combustion section  18 . A shaft  26  extends axially through the compressor  16  and the turbine  24  along an axial centerline  28  of the gas turbine engine  10 . 
     Referring now to the drawings,  FIG. 1  is a schematic view of an exemplary gas turbine engine  10  that may incorporate various embodiments disclosed herein. As shown, the gas turbine engine  10  generally includes a compressor section  12  having an inlet  14  disposed at an upstream end of a compressor  16  (e.g., an axial compressor). The gas turbine engine  10  also includes a combustion section  18  having one or more combustors  20  positioned downstream from the compressor  16 . The gas turbine engine  10  further includes a turbine section  22  having a turbine  24  (e.g., an expansion turbine) disposed downstream from the combustion section  18 . A rotor shaft  26  extends axially through the compressor  16  and the turbine  24  along an axial centerline  28  of the gas turbine engine  10 . 
       FIG. 2  is a cross-sectional side view of the turbine  24 , which may incorporate various embodiments disclosed herein. As shown in  FIG. 2 , the turbine  24  may include multiple turbine stages. For example, the turbine  24  may include a first stage  30 A, a second stage  30 B, and a third stage  30 C. Although, the turbine  24  may include more or less turbine stages as is necessary or desired. 
     Each stage  30 A- 30 C includes, in serial flow order, a corresponding row of turbine nozzles  32 A,  32 B, and  32 C and a corresponding row of turbine rotor blades  34 A,  34 B, and  34 C axially spaced apart along the rotor shaft  26  ( FIG. 1 ). Each of the turbine nozzles  32 A- 32 C remains stationary relative to the turbine rotor blades  34 A- 34 C during operation of the gas turbine  10 . Each of the rows of turbine nozzles  32 B,  32 C is respectively coupled to a corresponding diaphragm  42 B,  42 C. Although not shown in  FIG. 2 , the row of turbine nozzles  32 A may also couple to a corresponding diaphragm. A first turbine shroud  44 A, a second turbine shroud  44 B, and a third turbine shroud  44 C circumferentially enclose the corresponding row of turbine blades  34 A- 34 C. A casing or shell  36  circumferentially surrounds each stage  30 A- 30 C of the turbine nozzles  32 A- 32 C and the turbine rotor blades  34 A- 34 C. 
     As illustrated in  FIGS. 1 and 2 , the compressor  16  provides compressed air  38  to the combustors  20 . The compressed air  38  mixes with fuel (e.g., natural gas) in the combustors  20  and burns to create combustion gases  40 , which flow into the turbine  24 . The turbine nozzles  32 A- 32 C and turbine rotor blades  34 A- 34 C extract kinetic and/or thermal energy from the combustion gases  40 . This energy extraction drives the rotor shaft  26 . The combustion gases  40  then exit the turbine  24  and the gas turbine engine  10 . As will be discussed in greater detail below, a portion of the compressed air  38  may be used as a cooling medium for cooling the various components of the turbine  24  including, inter alia, the turbine nozzles  32 A- 32 C. 
       FIG. 3  is a perspective view of the turbine nozzle  32 B of the second stage  30 B, which may also be known in the industry as the stage two nozzle or S2N. The other turbine nozzles  32 A,  32 C include features similar to those of the turbine nozzle  32 B, which will be discussed in greater detail below. As shown in  FIG. 3 , the turbine nozzle  32 B includes an inner side wall  46  and an outer side wall  48  radially spaced apart from the inner side wall  46 . A pair of airfoils  50  extends in span from the inner side wall  46  to the outer side wall  48 . In this respect, the turbine nozzle  32 B illustrated in  FIG. 3  is referred to in the industry as a doublet. Nevertheless, the turbine nozzle  32 B may have only one airfoil  50  (i.e., a singlet), three airfoils  50  (i.e., a triplet), or more airfoils  50 . 
     As illustrated in  FIG. 3 , the inner and the outer side walls  46 ,  48  include various surfaces. More specifically, the inner side wall  46  includes a radially outer surface  52  and a radially inner surface  54  positioned radially inwardly from the radially outer surface  52 . Similarly, the outer side wall  48  includes a radially inner surface  56  and a radially outer surface  58  oriented radially outwardly from the radially inner surface  56 . As shown in  FIGS. 2 and 3 , the radially inner surface  56  of the outer side wall  48  and the radially outer surface  52  of the inner side wall  46  respectively define the inner and outer radial flow boundaries for the combustion gases  40  flowing through the turbine  24 . The inner side wall  46  also includes a forward surface  60  and an aft surface  62  positioned downstream from the forward surface  60 . The inner side wall  46  further includes a first circumferential surface  64  and a second circumferential surface  66  circumferentially spaced apart from the first circumferential surface  64 . Similarly, the outer side wall  48  includes a forward surface  68  and an aft surface  70  positioned downstream from the forward surface  68 . The outer side wall  48  also includes a first circumferential surface  72  and a second circumferential surface  74  spaced apart from the first circumferential surface  72 . The inner and the outer side walls  46 ,  48  are preferably constructed from a nickel-based superalloy or another suitable material capable of withstanding the combustion gases  40 . 
     As mentioned above, two airfoils  50  extend from the inner side wall  46  to the outer side wall  48 . As illustrated in  FIGS. 3 and 4 , each airfoil  50  includes a leading edge  76  disposed proximate to the forward surfaces  60 ,  68  of the inner and the outer side walls  46 ,  48 . Each airfoil  50  also includes a trailing edge  78  disposed proximate to the aft surfaces  62 ,  70  of the inner and the outer side walls  46 ,  48 . Furthermore, each airfoil  50  includes a pressure side wall  80  and an opposing suction side wall  82  extending from the leading edge  76  to the trailing edge  78 . The airfoils  50  are preferably constructed from a nickel-based superalloy or another suitable material capable of withstanding the combustion gases  40 . 
     Each airfoil  50  may define one or more inner cavities therein. An insert may be positioned in each of the inner cavities to provide the compressed air  38  (e.g., via impingement cooling) to the pressure-side and suction-side walls  80 ,  82  of the airfoil  50 . In the embodiment illustrated in  FIG. 4 , each airfoil  50  defines a forward inner cavity  86  having forward insert  90  positioned therein and an aft inner cavity  88  having an aft insert  92  positioned therein. A rib  94  ( FIG. 5 ) may separate the forward and aft inner cavities  86 ,  88 . Nevertheless, the airfoils  50  may define one inner cavity, three inner cavities, or four or more inner cavities in alternate embodiments. Furthermore, some or all of the inner cavities may not include inserts in certain embodiments as well. 
       FIGS. 5-8  illustrate embodiments of an impingement insert  100 , which may be incorporated into the gas turbine engine  10 . In particular, the impingement insert  100  may be positioned in the forward inner cavity  86  of one of the airfoils  50  in the nozzle  32 B in place of the forward insert  90  shown in  FIG. 4 . 
     As illustrated in  FIGS. 5-8 , the impingement insert  100  defines an axial direction A, a radial direction R, and a circumferential direction C. In general, the axial direction A extends parallel to the axial centerline  28 , the radial direction R extends orthogonally outward from the axial centerline  28 , and the circumferential direction C extends concentrically around the axial centerline  28 . 
     As illustrated in  FIGS. 5 and 6 , the impingement insert  100  includes a generally tubular insert wall  102  that defines an inner cavity  104  therein. In this respect, the insert wall  102  includes an inner surface  106 , which forms the outer boundary of the inner cavity  104 , and an outer surface  108  spaced apart from the inner surface  106 . In the embodiment illustrated in  FIG. 5 , the insert wall  102  generally has a D-shape. Although, the insert wall  102  may have any suitable shape (e.g., annular) in other embodiments as well. 
     Referring particularly to  FIG. 6 , the impingement insert  100  includes a plurality of nozzles  110  extending outwardly from the outer surface  108  of the insert wall  102 . In the embodiment shown in  FIG. 6 , the impingement insert  100  includes ten nozzles  110  positioned in two rows each having five nozzles  110 . The nozzles  110  are spaced apart within the rows in a manner that provides sufficient impingement cooling to the airfoil  50  as will be discussed in greater detail below. Preferably, the rows of nozzles  110  extend along substantially the entire radial length of the insert wall  102 . Although, the rows of nozzles  110  may extend along only a portion of the radial length of the insert wall  102  as well. Nevertheless, the plurality of nozzles  110  may be arranged in any suitable manner on the insert wall  102 . Furthermore, any number of nozzles  110  may extend outwardly from the outer surface  108  of the insert wall  102  so long as at least one nozzle  110  extends outwardly therefrom. 
     Referring again to  FIG. 5 , the impingement insert  100  is spaced apart from the pressure-side wall  80 , the suction-side wall  82 , and the rib  94  of the airfoil  50 . As illustrated therein, an inner surface  96  of the airfoil  50  (i.e., of the pressure-side wall  80 , the suction-side wall  82 , and the rib  94 ) forms the outer boundary of the forward inner cavity  86 . The impingement insert  100  is positioned within the forward inner cavity  86  in such a manner that the outer surface  108  of the insert wall  102  and the plurality of nozzles  110  are axially and/or circumferentially spaced apart from the inner surface  96  of the pressure-side wall  80 , the suction-side wall  82 , and the rib  94 . The spacing between the nozzles  110  and the inner surface  96  of the airfoil  50  should be sized to facilitate impingement cooling of the inner surface  96  as will be discussed in greater detail below. 
       FIGS. 7, 8A, and 8B  illustrate one of the nozzles  110  in greater detail. As depicted therein, the nozzle  110  has a frustoconical shape. More specifically, the nozzle  110  extends circumferentially outwardly from the outer surface  108  of the insert wall  102  and terminates at an outer surface  112  of the nozzle  110 . The outer surface  112  of the nozzle  110  is oriented parallel with and circumferentially spaced apart from the outer surface  108  of the insert wall  102 . Furthermore, the radial length of the nozzle  110  decreases from the outer surface  108  of the insert wall  102  to the outer surface of the nozzle  110 . The nozzle  110  also includes a circumferential surface  114 . 
     In the embodiment shown in  FIG. 7 , the impingement insert  100  includes a pedestal  116  that supports the nozzle  110 . As will be discussed in greater detail, the impingement insert  100  may formed via additive manufacturing methods. In this respect, the pedestal  116  provides the support necessary to form the nozzle  110  using additive manufacturing processes. As such, the pedestal  116  is positioned radially inward of the nozzle  110 . In particular, the pedestal  116  includes a pedestal surface  162  extends circumferentially and radially outward from the outer surface  108  of the insert body  102  and couples to a portion of the circumferential surface  114  of the nozzle  110 . In this respect, the pedestal  116  defines a pedestal angle  160  extending between the pedestal surface  162  and a circumferential line  164  extending circumferentially outward from the outer surface  108  of the insert wall  102 . The pedestal angle  160  may be between thirty degrees and ninety degrees. In the embodiment shown in  FIG. 7 , the pedestal  116  has a triangular cross-sectional shape. Nevertheless, the pedestal  116  may have any suitable cross-sectional shape as well. Some embodiments, however, may not include the pedestal  116 . 
       FIG. 10  illustrated an embodiment of the impingement insert  100  where the pedestal angle is ninety degrees. In this embodiment, the outlet portion  128  is flush with the outer surface  108  of the insert body  102  as illustrated in  FIG. 10 . In this respect, the nozzle  110  may extend circumferentially inwardly from the outer surface  108  of the insert wall  102 . 
     As illustrated in  FIG. 7 , the nozzle  110  and the insert wall  102  collectively define a cooling passage  118  extending therethrough. In particular, the cooling passage  118  extends from the inner surface  106  of the insert wall  102  to the outer surface  112  of the nozzle  110 . In this respect, the cooling passage  118  fluidly couples the inner cavity  104  of the impingement insert  100  and the forward inner cavity  86  of the airfoil  50 . As such, the cooling passage  118  provides impingement cooling to a portion of the inner surface  96  of the airfoil  50  as will be discussed in greater detail below. 
     The cooling passage  118  generally has a venturi-like configuration. More specifically, the cooling passage  118  includes an inlet portion  120 , a converging portion  122 , a throat portion  124 , a diverging portion  126 , and an outlet portion  128 . The inlet portion  120  occupies the circumferentially innermost position of the cooling passage  118 . In the embodiment illustrated in  FIG. 7 , the inlet portion  120  is entirely circumferentially aligned with the inner surface  106  of the insert wall  102 . Nevertheless, the inlet portion  120  may extend circumferentially outward from the inner surface  106  of the insert wall  102  (i.e., into the insert wall  102 ) as well. The converging portion  122  extends from the inlet portion  120  to the throat portion  124 . In particular, the diameter of the converging portion  122  narrows from the inlet portion  120  to the throat portion  124 . The throat portion  124  generally occupies the portion of the cooling passage  118  having the smallest diameter. In this respect, the throat portion  124  is positioned at a central position along the circumferential length of the cooling passage  118 . In the embodiment shown in  FIG. 7 , the throat portion  124  is circumferentially aligned with the outer surface  108  of the insert wall  102 . Although, the throat portion  124  may be positioned circumferentially inward or outward of the outer surface  108  as well. The diverging portion  126  extends from the throat portion  124  to the outlet portion  128 . The diameter of the diverging portion  126  expands from the throat portion  124  to outlet portion  128 . The outlet portion  128  occupies the circumferentially outermost position of the cooling passage  118 . In the embodiment illustrated in  FIG. 7 , the outlet portion  128  is entirely circumferentially aligned with the outer surface  112  of the nozzle  110 . Nevertheless, the outlet portion  128  may extend from circumferentially inward from the outer surface  112  of the nozzle  110  (i.e., into the nozzle  110 ) as well. 
     The converging portion  122  and the diverging portion  126  define circumferential lengths. In particular, the converging portion  122  defines a converging portion length  130  extending circumferentially from the inlet portion  120  to the throat portion  124 . Similarly, the diverging portion  126  defines a diverging portion length  132  extending circumferentially from the throat portion  124  to the outlet portion  128 . In the embodiment shown in  FIG. 7 , the converging length  130  is the same as the diverging length  132 . Although, the converging length  130  and the diverging length  132  may be different in other embodiments. 
     The converging portion  122  and the diverging portion  128  may respectively define converging and diverging angles. As illustrated in  FIG. 7 , the cooling passage  118  defines a circumferential centerline  132  extending therethrough. In this respect the converging portion  122  defines a converging portion angle  136  at which the converging portion  122  expands radially outwardly from the throat portion  124  to inlet portion  120  relative to the circumferential centerline  132 . Similarly, the diverging portion  128  defines a diverging portion angle  138  at which the diverging portion  128  expands radially outwardly from the throat portion  124  to outlet portion  128  relative to the circumferential centerline  132 . In the embodiment shown in  FIG. 7 , the converging portion angle  136  is greater than the diverging portion angle  138 . The diverging portion angle  138  is preferably ten degrees, but may be as low as three degrees or high as fifteen degrees. The converging portion angle  136  is typically greater than fifteen degrees and may be as high as seventy-five degrees. Although, the converging portion angle  136  may the same as or smaller than the diverging portion angle  138  in other embodiments. 
       FIGS. 8A and 8B  illustrate different embodiments of a cross-sectional shape  140  of the cooling passage  118 . In particular, the cross-sectional shape  140  includes a semicircular portion  142  and a non-circular portion  144 . The semicircular portion  142  is positioned radially inwardly from the non-circular portion  144 . In the embodiments shown in  FIGS. 8A and 8B , the semicircular portion  142  forms the radially inner half of the cross-sectional shape  140 , while the non-circular portion  144  forms the radially outer half of the cross-sectional shape  140 . In this respect, the non-circular portion  144  of the cross-sectional shape  140  is directly coupled to the semicircular portion  142  of the cross-sectional shape  140 . Nevertheless, the semicircular and non-circular portions  142 ,  144  may occupy more or less than half of the cross-sectional shape  140  and may be spaced apart by other portions (not shown) of the cross-sectional shape  140 . 
       FIG. 8A  illustrates one embodiment of the non-circular portion  144  of the cross-sectional shape  140 . As illustrated therein, the non-circular portion  144  includes a first linear side  146  and a second linear side  148 . The first and the second linear sides  146 ,  148  extend radially outwardly and axially toward one another. In this respect, the first linear side  146  is oriented at an angle  158  relative to the second linear side  148 . The angle  158  is between 60 degrees and 120 degrees in some embodiments. In certain embodiments, angle  158  may be 90 degrees. A fillet  150  couples the first and the second linear sides  146 ,  148 . The non-circular portion  144 , and more particularly the first and the second linear sides  146 ,  148 , provide the support necessary to form the portions of the nozzle  110  circumferentially aligned with and positioned radially outwardly from the cooling passage  118  when using additive manufacturing processes. 
       FIG. 8B  illustrates another embodiment of the non-circular portion  144  of the cross-sectional shape  140 . The first and the second side linear sides  146 ,  148  extend radially outwardly and axially toward one another as with the embodiment shown in  FIG. 8A . As such, the first linear side  146  is oriented at an angle  158  relative to the second linear side  148 . The angle  158  is between 60 degrees and 120 degrees in some embodiments. In certain embodiments, angle  158  may be 90 degrees. In this embodiment shown in  FIG. 8B , however, the first linear side  146  couples to the second linear side  148 , thereby giving the non-circular portion  144  a triangular shape. Nevertheless, the non-circular portion  144  of the cross-sectional shape  140  may have any suitable non-circular shape. 
     The first and the second linear sides  146 ,  148  define lengths. In particular, the first linear side  146  defines a first linear side length  152 , and the second linear side  148  defines a second linear side length  154 . In the embodiment shown in  FIG. 8B , the first linear side length  152  is the same as the second linear side length  154 . In this respect, the non-circular portion  144  of the cross-sectional shape  140  is shaped like an isosceles triangle in the embodiment shown in  FIG. 8B . Although, the first linear side length  152  and the second linear side length  154  may be different in other embodiments. 
     Preferably, the impingement insert  100  is integrally formed. In this respect, the insert wall  102 , the nozzles  110 , and the pedestals  116  are all formed as a single component. Nevertheless, the impingement insert  100  may be formed from two or more separate components as well. 
     As mentioned above, the impingement insert  100  is preferably formed via additive manufacturing. The term “additive manufacturing” as used herein refers to any process which results in a useful, three-dimensional object and includes a step of sequentially forming the shape of the object one layer at a time. Additive manufacturing processes include three-dimensional printing (3DP) processes, laser-net-shape manufacturing, direct metal laser sintering (DMLS), direct metal laser melting (DMLM), plasma transferred arc, freeform fabrication, etc. A particular type of additive manufacturing process uses an energy beam, for example, an electron beam or electromagnetic radiation such as a laser beam, to sinter or melt a powder material. Additive manufacturing processes typically employ metal powder materials or wire as a raw material. Nevertheless, the impingement insert  100  may be constructed using any suitable manufacturing process. 
     In operation, the impingement insert  100  provides cooling air  156  to the airfoils  50  of the nozzle  32 B. As illustrated in  FIG. 2 , a portion of the compressed air  38  bled from the compressor section  12  ( FIG. 1 ) is directed into the nozzle  32 B. In particular, this portion of the compressed air  38  flows through the inner cavity  104  defined by the impingement insert  100  positioned in the forward cavity  86  of the nozzle  32 B. In this respect, the compressed air  38  flows radially inwardly through the airfoils  50  of the nozzle  32 B (i.e., from the outer side wall  48  toward the inner side wall  46 ). As will be discussed in greater detail below, the impingement insert  100  directs at least a portion of the compressed air  38  flowing through the inner cavity  104  onto the inner surface  96  of the airfoil  50 . The portion of the compressed air  38  directed onto the inner surface  96  will hereinafter be referred to as the cooling air  156 . 
     As illustrated in  FIG. 9 , the cooling air  156  cools the inner surface  96  of the airfoil  50  via impingement cooling. More specifically, the cooling air  156  flows from the inner cavity  104  of the impingement insert  100  into inlet portion  120  of the cooling passage  118 . The cooling air  156  flows sequentially through the inlet portion  120 , the converging portion  122 , the throat portion  124 , the diverging portion  126 , and the outlet portion  128  of the cooling passage  118 . The venturi-like configuration of the cooling passage  118  increases the velocity of the cooling air  156  flowing therethrough. The cooling air  156  exits the cooling passage  118  and flows through the forward inner cavity  86  until striking the inner surface  96  of the airfoil  50 . As such, cooling passage  118  provides impingement cooling to airfoil  50 . In this respect, the nozzle  110  should have a circumferential length that permits impingement cooling of the airfoil  50 . Furthermore, the cooling passage  110  should be sized and arranged to provide impingement cooling of the airfoil  50  as well. 
     As discussed in greater detail above, the venturi-like configuration of the cooling passage  118  increases the velocity of the cooling air  156  flowing therethrough. In this respect, each cooling passage  110  provides greater impingement cooling to the inner surface  96  of the airfoil  50  than conventional impingement cooling passages. As such, the impingement insert  100  may define fewer cooling passages  110  extending therethrough than conventional inserts having conventional impingement cooling passages. Accordingly, the impingement insert  100  diverts less compressed air  38  from the compressor section  12  ( FIG. 1 ) than conventional inserts, thereby increasing the efficiency of the gas turbine engine  10 . 
     The impingement insert  100  was discussed above in the context of the forward insert  90  positioned in the forward cavity  86  of the second stage nozzle  32 B. Nevertheless, the impingement insert  100  may be any insert positioned in any cavity of any nozzle in the gas turbine engine  10 . In some embodiments, the impingement insert  100  may be incorporated into one or more of the turbine shrouds  44 A- 44 C or one or more of the rotor blades  32 A- 32 C. In fact, the impingement insert  100  may be incorporated into any suitable component in the gas turbine engine  10 . 
     This written description uses examples to disclose the technology, including the best mode, and also to enable any person skilled in the art to practice the technology, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the technology 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 include 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.