Impact insert for reusing impact air in an airfoil, airfoil comprising an impact insert, turbomachinery component and gas turbine provided therewith

Abstract

Impact insert (80) for a turbomachinery component, the impact insert (80) comprising: - a double-walled section (1, 2) having an inner wall (81) and an outer wall (82) defining an inner channel (501) formed on an inner surface (81a) of the inner wall (81), an outer channel (503) formed on an outer surface (82b) of the outer wall (82), and a central channel (502) formed between the inner surface (82a) of the outer wall (82) and the outer surface (81b) of the inner wall (81); and - a plurality of impact cooling holes (85) formed in the outer wall (82) and configured to eject impact jets (86) into the outer channel (503), wherein the impact jets (86) are formed from the cooling air (5) of the central channel (502); characterized in that the impact insert (80) further - at least one suction line (9) extending between the outer wall (82) and the inner wall (81) over the central channel (502) and comprising an inlet (9a) on the outer channel (503) and an outlet (9b) on the inner channel (501) for sucking cooling air (5) from the outer channel (503) into the inner channel (501), and wherein the suction line (9) is aerodynamically shaped with respect to a flow of the cooling air (5) flowing through the central channel (502).

Description

[0001] The present invention relates to an impingement insert for reusing impingement air in an airfoil, an airfoil comprising an impingement insert, a turbomachinery component and a gas turbine having the same, and in particular to cooling a turbomachinery component or an airfoil comprising such an impingement insert.

[0002] Turbomachinery contains various turbomachinery components that benefit from cooling, resulting in increased component life. Cooling turbomachinery components allows for an increase in combustion temperature, which increases the overall power output of the engine.

[0003] Certain turbomachinery components have an airfoil, such as a blade or a vane. The airfoils enclose internal spaces and are cooled internally, or from within, by cooling air flowing through the interior of the airfoil or through one or more cooling channels formed within the airfoil.

[0004] The turbomachinery component—hereinafter also referred to as the blade or vane—generally comprises the airfoil (also referred to as the aerodynamic airfoil) having an airfoil wall and an interior defined by the airfoil wall. During operation of the gas turbine, the airfoils of the turbine section of the gas turbine are positioned in the hot gas path and are exposed to very high temperatures. To provide cooling to the airfoil, one or more cooling channels are defined in the interior of the airfoil. The entire interior of the airfoil can form a cooling channel that generally extends in the longitudinal direction of the airfoil.

[0005] Alternatively, the airfoil may contain one or more ribs on its inner surface, extending from a pressure side to a suction side of the airfoil, thereby mechanically strengthening the airfoil. Depending on the number of ribs, the rib divides the interior of the airfoil into one or more cooling channels extending along the longitudinal direction of the airfoil.

[0006] In such cooling channels, cooling air generally flows along the longitudinal direction of the airfoil after being introduced into the airfoil. Improving such internal airfoil cooling would have a beneficial effect on the efficiency of the gas turbine and / or the structural integrity of the airfoil.

[0007] It is generally known to use impingement cooling of an inner surface of the blade, for example by using impingement inserts in the cooling channels. Fig. 10 shows a conventional impingement insert 80'. The wall of the impingement insert 80' defines a flow channel in which the cooling air 5 flows. The wall of the impingement insert 80' contains a plurality of impingement cooling holes 85 facing an inner surface of the airfoil wall 101. The cooling air from the flow channel is directed out of the impingement cooling holes 85 in the form of impingement cooling jets 86 to impinge on the inner surface of the airfoil wall 101. The impinged air then flows into the space between the impingement insert 80' and the airfoil wall 101. This creates crossflows 5x for the impingement jets 86, which are located downstream in the flow direction of the impingement air flowing in the space between the impingement insert 80' and the airfoil wall 101. This reduces the cooling efficiency in such downstream sections or regions of the airfoil wall 101. Therefore, it is desirable to reduce such cross flows.

[0008] US 5,120,192 A shows a cooled turbine blade and a combined-cycle power plant with a gas turbine. The blade includes a blade body with a cavity, an insert inserted into the cavity, a plurality of impingement holes formed in a peripheral wall of the insert, and a recovery path for recovering a cooling medium sprayed from the impingement holes. The cooling medium can be recovered without being discharged into a main gas. In this system, steam is supplied to the blade as a cooling medium, and the steam recovered from the blade is fed to a steam turbine.

[0009] JP S56 72 201 A shows a cooling structure for a gas turbine blade. The impingement cooling insert inside the blade is designed as a double structure, and cooling air from a compressor is introduced into a chamber. The cooling air from the inner insert impinges on a blade wall through an impingement cooling opening, and interference with the impingement cooling air caused by crossflow is eliminated by a protruding portion of the outer insert. The cooling air exiting from a plurality of cooling air exhaust ports is guided to an exhaust duct and an opening at the trailing edge and discharged into a main gas flow through a blowout opening. This achieves better cooling performance, reduces the delay in the effectively cooled surface position, and achieves high cooling efficiency with a small cooling air flow rate.

[0010] DE 44 30 302 A1 shows an impact cooling system for wall parts, characterized by a plurality of impact tubes which are arranged with their inlet flat on a flat or curved support and with their mouth directed towards the wall part to be cooled, wherein the support is arranged at a distance from the wall part.

[0011] GB 849 255 A presents a method for cooling walls of combustion chambers or other thermally highly loaded spaces by means of a cooling medium which impinges on the outer surface of the wall in individual jets, wherein the cooling medium is introduced into the wall through nozzles arranged substantially at right angles thereto and the heated cooling medium is withdrawn directly from the region of the cooled wall.

[0012] US 2010 / 0 221 123 A1 describes a cooling structure for a gas turbine blade with several impingement cooling inserts.

[0013] Furthermore, to cool the components of the gas turbine, a portion of the air is sucked from the compressor section of the gas turbine and directed to various sections of the gas turbine to be used as cooling air. More cooling can be beneficial and can be achieved by sucking more air from the compressor. However, an increase in the amount of air sucked by the compressor for cooling inadvertently leads to a decrease in the amount of air available for combustion - this can adversely affect the efficiency of the gas turbine. Therefore, it would be advantageous if the cooling air that has been used once, e.g. for impingement cooling of a first surface, could be used to cool another surface, e.g.a second surface, by collecting or vacuuming it to be reused after it has been used on the first surface to form impingement jets that can impact the second surface.

[0014] Therefore, it is advantageous to improve the internal cooling of the blade.

[0015] The above objects are achieved by the subject matter of the independent claims, in particular by an insert for a turbomachine component for a gas turbine. Advantageous embodiments are specified in the dependent claims.

[0016] Such turbomachinery components incorporating an airfoil are illustrated below by a blade, but the description is also applicable to other turbomachinery components incorporating an airfoil, such as a vane, unless otherwise specified.

[0017] According to a first aspect of the present technology, an impact insert for a turbomachinery component is provided.

[0018] The turbomachinery component may be a component having an airfoil, e.g., a blade or a vane of a turbine. One or more cooling channels may be formed in the airfoil of the turbomachinery component. The impingement insert may be inserted or installed in such a cooling channel to provide impingement jets on an inner surface of the cooling channel, e.g., the inner surface of the airfoil wall. Thus, the present technique also contemplates the turbomachinery component described above.

[0019] The impact insert, hereinafter also referred to as the insert, includes a double-walled structure or section having an outer wall and an inner wall.

[0020] The inner wall and the outer wall of the double-walled section define three spatial compartments - an inner channel formed on an inner surface of the inner wall, an outer channel formed on an outer surface of the outer wall, and a central channel formed between the inner surface of the outer wall and the outer surface of the inner wall.

[0021] The inner channel may be defined by a wall on the opposite side of the impact insert. In other words, if the double-walled section is present on a pressure side, then the opposite side would be the suction side, and vice versa. The wall on the opposite side of the impact insert may also include a double-walled section, similar to aspects of the present technique. Alternatively, the wall on the opposite side of the impact insert may be a single wall.

[0022] The impingement insert contains a plurality of impingement cooling holes formed in the outer wall and configured to expel the impingement jets into the outer channel. The impingement jets are formed from or by the cooling air from the central channel. In other words, the cooling air from the central channel is expelled as impingement jets through the impingement cooling holes into the outer channel.

[0023] The impact insert may contain at least one extraction duct. The extraction duct extends between the outer wall and the inner wall across the central channel. The at least one extraction duct thus guides the air through the central channel. Consequently, the cooling air entering the impact cooling holes from the central channel is not mixed with the air guided through the at least one extraction duct or is isolated from it.

[0024] The exhaust duct may have an inlet in the outer wall, preferably located in the outer surface of the outer wall, and may have an outlet in the inner wall of the insert, preferably located in the inner surface of the inner wall, so that the cooling air can flow from the outer channel through the exhaust duct into the inner channel. The exhaust duct thus operates to exhaust the cooling air from the outer channel into the inner channel.

[0025] Consequently, according to the present technique, the cooling air provided in the airfoil enters the impingement insert, particularly the double-walled portion of the impingement insert, flows into the central channel, and is then ejected as the impingement jets into the outer channel onto the inner surface of the airfoil to provide impingement cooling, and is then exhausted from the outer channel into the inner channel through the exhaust duct.

[0026] The extracted cooling air may have been used once in the outer duct to cool the inner surface of the blade wall facing the outer duct or adjacent to or facing the outer surface of the outer wall.

[0027] Preferably, this extracted cooling air can then be used for some further processes, such as providing impingement cooling to another part or section of the inner surface of the airfoil wall.

[0028] According to the present technique, the size of the inlet and / or outlet of the exhaust duct can be larger than the size of the impingement cooling holes. Consequently, a gentle flow of cooling air through the exhaust duct is provided, thus promoting the exhaust process.

[0029] The outer wall of the insert may have a corrugated shape.

[0030] The corrugated shape may comprise a plurality of recesses or troughs extending in a direction away from the inner wall, and one or more projections or one or more ribs located between the recesses or troughs, i.e., in an alternating manner. One or more impingement cooling holes may be provided in at least one of the recesses or troughs. Preferably, all of the recesses or troughs are provided with one or more impingement cooling holes.

[0031] The outlet of the suction line may be positioned on the one or more ribs or the one or more projections.

[0032] The double-walled section may include at least one main inlet for the cooling air. The at least one main inlet may be an inlet of the central duct.

[0033] The double-walled section may be configured so that the cooling air received via the main inlet in the central duct is expelled as impingement jets via the impingement cooling holes and then extracted from the outer duct into the inner duct via the extraction duct.

[0034] According to an exemplary embodiment, in a first sub-aspect of the first aspect, the main inlet can be arranged on a top and / or bottom of the central channel. The top and / or bottom can be understood as the sides or regions of the central channel that are spaced apart along a longitudinal direction of the impact insert. The cooling air can enter the central channel along the longitudinal direction.

[0035] According to a further exemplary embodiment, in a second sub-aspect of the first aspect, the at least one main inlet can be arranged on a lateral side of the central channel. The lateral side can be understood as running parallel to a longitudinal direction of the impact insert. The cooling air can enter the central channel perpendicular to the longitudinal direction.

[0036] According to another exemplary embodiment, the impact insert may have a first portion and a second portion. The first portion may be in accordance with the first sub-aspect of the first aspect described above, and the second portion may be in accordance with the second sub-aspect of the first aspect described above.

[0037] A main outlet of the first section may be fluidly connected to the main inlet of the second section.

[0038] The main outlet of the first section may be an outlet of the inner channel.

[0039] The cooling air can flow over or over the outer surfaces of the exhaust duct. The exhaust duct is therefore aerodynamically shaped with respect to the flow direction of the cooling air entering the central duct along the longitudinal direction or perpendicular to the longitudinal direction. The cross-section can be circular, oval, or elliptical.

[0040] According to a second aspect of the present technology, a turbomachinery component for a gas turbine is provided.

[0041] The turbomachinery component may include an airfoil having an airfoil wall defining an interior of the airfoil. At least one cooling channel may be formed in the interior of the airfoil. An impact insert may be inserted into the cooling channel. The impact insert may be in accordance with the first aspect of the present technique described above. The outer channel may be defined between the outer surface of the outer wall and an inner surface of the airfoil wall.

[0042] In the turbomachinery component, the double-walled section may include a main inlet formed on the central channel and a main outlet formed on the inner channel.

[0043] The outer duct may be a closed chamber, except for, i.e., adjacent to, the impingement cooling holes of the outer wall and the inlet of the exhaust duct, and optionally, one or more film cooling holes that may be present in the airfoil wall. In other words, the outer duct may be a sealed space into which cooling air can only enter through the impingement cooling holes, i.e., there are no other inlets for air into the outer duct, and from which cooling air can only exit via the inlet of the exhaust duct or through one or more film cooling holes that may be optionally present, i.e., there are no other air outlets from the outer duct.

[0044] The central duct can be a closed chamber, except for, i.e., adjacent to, the impingement cooling holes of the outer wall and the main inlet of the double-walled section. In other words, the central duct can be a sealed space into which cooling air can only enter via the main inlet of the double-walled section, i.e., there are no other air inlets into the central duct, and which cooling air can only exit via the impingement cooling holes, i.e., there are no other air outlets from the central duct.

[0045] The inner duct may be a closed chamber except for, i.e., adjacent to, the outlet of the exhaust duct and the main outlet of the double-walled section. In other words, the inner duct may be a sealed space into which cooling air can only enter through the outlet of the exhaust duct, i.e., there are no other air inlets into the inner duct, and which cooling air can only exit via the main outlet of the double-walled section, i.e., there are no other air outlets from the inner duct.

[0046] The inner surface of the airfoil wall may include exhaust ducts that project from the inner surface of the airfoil wall to the outer surface of the outer wall. After impacting the inner surface of the airfoil wall, the cooling air is guided through the exhaust ducts toward the inlet of the exhaust duct.

[0047] According to a third aspect of the present technology, a gas turbine is presented. The gas turbine includes a turbomachine component according to the second aspect of the present technology.

[0048] The above-mentioned attributes and other features and advantages of the present technique and the manner of achieving them will become more apparent from the following description of embodiments of the present technique taken in conjunction with the accompanying drawings, in which the present technique itself will be better understood; in which: Fig. 1 shows a portion of an exemplary embodiment of a gas turbine in a sectional view incorporating a turbomachine component of the present technique; Fig. 2 is a perspective view illustrating an exemplary embodiment of a turbomachine assembly including an exemplary embodiment of a turbomachine component according to the present technique, illustrated by a blade in accordance with the present technique; Fig. 3 a cross-sectional view along the line YY in Fig. 2, which schematically illustrates an exemplary location of an impact insert of the present technique; Fig. 4 schematically illustrates an exemplary embodiment of the impact insert according to the present technique; Fig. 5 one in Fig. 4 shown section M of the impact insert according to Fig. 4 in a schematic way; Fig. 6 another in Fig. 4 shown section N of the impact insert according to Fig. 4 in a schematic way; Fig. 7 schematically illustrates a portion of another exemplary embodiment of the impact insert of the present technique; Fig. 8 shows a larger section of the exemplary embodiment of the impact insert of the present technique, including the section according to Fig. 7, in a schematic manner; Fig. 9 a relative size and / or orientation and / or distribution of the impingement holes and an inlet of the suction duct of the present technique in a schematic manner; and Fig. 10 a conventional impact insert for comparative understanding of the impact insert of the present technology.

[0049] The above-mentioned and other features of the present technique are described in detail below. Various embodiments are described with reference to the drawings, wherein like reference numerals are used to refer to like elements throughout. In the following description, for purposes of explanation, numerous specific details are set forth in order to provide a thorough understanding of one or more embodiments. It may be stated that the illustrated embodiments are intended to explain and not to limit the invention. It may be apparent that such embodiments may be practiced without these specific details.

[0050] Fig. 1 shows an example of a gas turbine 10 in a sectional view. The gas turbine 10 may include, in flow order, an inlet 12, a compressor or compressor section 14, a combustor section 16, and a turbine section 18, arranged generally in flow order and generally about and in the direction of a longitudinal or rotational axis 20. The gas turbine 10 may further include a shaft 22 rotatable about the rotational axis 20 and extending longitudinally through the gas turbine 10. The shaft 22 may driveably connect the turbine section 18 to the compressor section 14.

[0051] During operation of the gas turbine 10, the air 24 taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or combustor section 16. The combustor section 16 may include a burner plenum 26, one or more combustion chambers 28, and at least one burner 30 attached to each combustion chamber 28. The combustion chambers 28 and the burners 30 may be located within the burner plenum 26. The compressed air flowing through the compressor section 14 may enter a diffuser 32 and be expelled from the diffuser 32 into the burner plenum 26, from where a portion of the air may enter the burner 30 and be mixed with a gaseous or liquid fuel. The air / fuel mixture is then combusted, with the combustion gas 34 or working gas from the combustion being passed through the combustion chamber 28 via a transition line 17 to the turbine section 18.

[0052] This exemplary gas turbine 10 may include a tubular combustor section assembly 16 formed by an annular array of combustor sleeves 19, each containing the burner 30 and the combustion chamber 28, with the transition conduit 17 having a generally circular inlet connected to the combustion chamber 28 and an outlet in the form of a ring segment. An annular array of transition conduit outlets may form an annular space for directing the combustion gases to the turbine section 18.

[0053] The turbine section 18 may include a number of blade-carrying rotor disks 36 attached to the shaft 22. In the present example, two rotor disks 36 are shown, each supporting an annular array of turbine blades 38. However, the number of blade-carrying disks could vary, i.e., only one disk or more than two disks. Additionally, guide vanes 40, which are attached to a stator 42 of the gas turbine 10, may be arranged between the stages of the annular arrays of turbine blades 38. Guide vanes 44 may be provided between the exit of the combustion chamber 28 and the front turbine blades 38 and direct the flow of the working gas onto the turbine blades 38.

[0054] The combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38, which in turn rotate the shaft 22. The guide vanes 40, 44 serve to optimize the angle of the combustion or working gas on the turbine blades 38.

[0055] The turbine section 18 drives the compressor section 14. The compressor section 14 includes an axial array of vane stages 46 and rotor vane stages 48. The rotor vane stages 48 may comprise a rotor disk supporting an annular array of blades. The compressor section 14 may further include a casing 50 surrounding the rotor stages and supporting the vane stages 46. The vane stages may include an annular array of radially extending guide vanes attached to the casing 50. The guide vanes are provided to present the gas flow at an optimal angle for the blades at a given operating point of the gas turbine.Some of the vane stages may have variable vanes, where the angle of the vanes about their own longitudinal axis can be adjusted according to the airflow characteristics that may occur under various operating conditions of the gas turbine. The casing 50 may define a radially outer surface 52 of the passage 56 of the compressor section 14. A radially inner surface 54 of the passage 56 may be at least partially defined by a rotor drum 53 of the rotor, which may be partially defined by the annular array of rotor vane stages 48.

[0056] The present technique is described with respect to the above exemplary gas turbine, which has a single shaft or pulley connecting a single, multi-stage compressor and a single, single- or multi-stage turbine. However, it should be recognized that the present technique is equally applicable to two- or three-shaft gas turbines, which may be used for industrial, aviation, or marine applications.

[0057] The terms upstream and downstream refer to the flow direction of the air flow and / or working gas flow through the gas turbine, unless otherwise stated. The terms forward and reverse refer to the general flow of gas through the gas turbine, unless otherwise stated. The terms axial, radial, and circumferential are used with respect to the rotational axis 20 of the gas turbine, unless otherwise stated.

[0058] In the present technique, a turbomachinery component is presented that includes an airfoil 100, such as shown in the Fig. 2 and Fig. 3. The turbomachinery component of the present technique may be the turbine bucket 38 of the gas turbine 10 described above, unless otherwise specified. The turbomachinery component of the present technique may be the nozzle vane 40, 44 of the gas turbine 10 described above, unless otherwise specified. Hereinafter, for the sake of simplicity and brevity, and not as a limitation, unless otherwise specified, the turbomachinery component has been illustrated and also referred to as a bucket of the gas turbine, but it may be noted that the turbomachinery component according to the present technique may also be another turbomachinery component including an airfoil in accordance with the present technique.

[0059] The Fig. 2 and Fig. 3 schematically illustrate an example of a turbomachinery component, illustrated by the turbine blade 38 of the gas turbine 10. Fig. Figure 2 schematically illustrates an example of a turbomachinery assembly. The assembly may include turbine blades 38 as the turbomachinery component, disposed on the rotor disk 36. The turbine blade 38 may include a platform 200, an airfoil 100, and optionally a root 300. The turbine blade 38 may be attached or mounted to the rotor disk 36 via the root 300.

[0060] In the turbomachinery component, the airfoil 100 extends from the platform 200. The platform 200 may include a top surface 201 and a bottom surface 210. The airfoil 100 may extend from the top surface 201 of the platform 200. The top surface 201 may extend in the circumferential direction. Similarly, the bottom surface 210 may extend in the circumferential direction. The airfoil 100 extends radially outward from the top surface 201 of the platform 200. The root 300 may extend radially downward from the bottom surface 210 of the platform 200, opposite to the direction of extension of the airfoil.

[0061] The airfoil 100 includes an airfoil wall 101 that encloses an interior space 100s of the airfoil. The airfoil wall 101 may include a pressure side 102 (also referred to as a pressure surface or concave surface / side) and a suction side 104 (also referred to as a suction side or convex surface / side). The pressure side 102 and the suction side 104 meet at a leading edge 106 and a trailing edge 108 of the airfoil 100.

[0062] The airfoil 100 may have a base portion 100b adjacent to the platform 200 and a tip portion 100a spaced from the base portion 100b along a longitudinal direction A of the airfoil 100.

[0063] The interior space 100s of the blade 100 can form a cooling channel 70 which is delimited by the blade wall 101.

[0064] Alternatively, at least one rib 60 may be arranged within the interior space 100s of the airfoil 100. The rib 60 may extend between the pressure side 102 and the suction side 104. More specifically, each rib 60 may extend between an inner surface 101a of the airfoil wall 101 of the airfoil 100 on the pressure side 102 of the airfoil 100 and an inner surface 101a of the airfoil wall 101 of the airfoil 100 on the suction side 104 of the airfoil 100. It may be stated that, although the example according to Fig. 3 shows two such ribs 60 for exemplary purposes, the airfoil 100 may have one or three or more ribs 60. Each of the ribs 60 is connected to the pressure side 102 and the suction side 104. More specifically, each of the ribs 60 may be connected to an inner surface 101a of the airfoil wall 101 on the pressure side 102 and to the inner surface of the airfoil wall 101 on the suction side 104.

[0065] The wall, i.e., the airfoil wall 101, of the airfoil 100, which includes the pressure side 102 and the suction side 104 and defines the leading edge 106 and the trailing edge 108, in addition to being referred to as the airfoil wall 101, may also be referred to as an outer wall of the airfoil 100 or as a primary wall of the airfoil 100. The airfoil wall 101 defines the external appearance of the airfoil 100, or in other words, defines the airfoil shape.

[0066] Each of the ribs 60 may also be understood to be formed by a wall in the airfoil 100, however, the wall forming the rib 60 is different than the airfoil wall 101 and may be referred to as an inner wall or secondary wall of the airfoil 100.

[0067] As in the example after Fig. 3, the interior space 100s of the blade 100 may include at least one cooling channel 70 for the flow of the cooling air 5. The cooling channels 70 may be understood as a whole interior space 100s or as subdivisions of the interior space 100s of the blade 100 created by the ribs 60. It may be stated that, although the example according to Fig. 3 shows three such cooling channels 70 for example purposes, the airfoil 100 may have 1 or 2 or 4 or more such cooling channels 70.

[0068] The cooling air 5 can be provided into the cooling channel 70 from outside the airfoil 100, e.g., through cooling air flow paths (not shown) formed in the root 300 of the blade 1. Alternatively, or in addition to the above, the cooling air 5 can be provided into the cooling channel 70 from another, preferably adjacent, cooling channel 70 of the airfoil 100, causing the cooling air to make a U-turn at the tip portion 100a or the base portion 100b of the airfoil, thereby flowing out of a first cooling channel 70 and then flowing into a second cooling channel 70 from a top or bottom side relative to the direction A of the cooling channel.

[0069] The cooling channel 70 can extend along the longitudinal direction A of the blade 100, as in the example according to the Fig. 2 and Fig. 3. As shown in the example according to Fig. 3, each cooling channel 70 of the airfoil may be defined by one or more of the ribs 60 and the pressure side 102 and the suction side 104. The example according to Fig. 3 shows a first cooling channel 70 defined by one of the ribs 60, a portion of the pressure side 102, a portion of the suction side 104, and the leading edge 106. The example according to Fig. 3 also shows a second cooling channel 70 defined by one of the ribs 60, a portion of the pressure side 102, a portion of the suction side 104 and the trailing edge 108. Furthermore, the example according to Fig. 3 shows a third cooling channel 70 defined by two adjacent fins 60 facing each other, a portion of the pressure side 102, and a portion of the suction side 104. The third cooling channel can be understood as the cooling channel between the first and second cooling channels and can also be present multiple times.

[0070] Fig. 3 also shows a schematic representation of one or more impact inserts 80 according to the present technique inserted, positioned, or formed in the cooling channel 70. The impact insert 80 according to the present technique will be described below with respect to the Fig. 4 to 9. A conventional impact insert 80' is shown for comparative understanding in Fig. 10 shown.

[0071] The impact inserts 80 (hereinafter also referred to as the insert) can generally be understood as a component inserted into the cooling channel 70 or as a component formed, for example, by additive manufacturing in the cooling channel 70, which includes one or more impact holes 85 for ejecting impact jets 86 of cooling air towards the inner surface 101a of the airfoil wall 101, preferably towards the pressure side 102 and / or the suction side 104 of the airfoil 100 and / or towards the leading edge 106 and / or towards the trailing edge 108 of the airfoil 100, for the purpose of impacting the inner surface 101a (hereinafter also referred to as a target surface) of the airfoil 100 to provide cooling of the target surface.

[0072] The impingement insert 80 may be inserted into the cooling channel 70 of the turbomachinery component, which may be the blade 38 or the guide vane 40, 44 of the gas turbine 10, to provide impingement cooling of the inner surface 101a of the airfoil wall 101 forming the cooling channel 70 in the airfoil 100 of the turbomachinery component of the gas turbine 10.

[0073] In the following, with reference to the Fig. 5 and Fig. 6 in combination with Fig. 4 explains an exemplary embodiment of the impact insert 80 of the present technique.

[0074] The impact insert 80, which is also referred to as the insert below, contains a double-walled structure or a double-walled section 1, 2. In Fig. 4, both section 1 and section 2 represent exemplary embodiments of such a double-walled section.

[0075] In general, the term 'double-walled' section or 'double-walled' structure can be understood as a section having two walls arranged substantially parallel to each other.

[0076] To further explain, when the impact insert 80 is positioned within the airfoil 100, a side or portion or region of the impact insert 80 may be disposed adjacent to the pressure side 102 of the airfoil 100, as shown in Fig. 3, which may be referred to as a pressure side of the impact insert 80. In other words, the pressure side of the impact insert 80 may also be understood as a side of the impact insert 80 for providing impact jets 86 toward the pressure side 102 of the airfoil 100.

[0077] Similarly, when the impact insert 80 is positioned within the airfoil 100, a side or portion or region of the impact insert 80 other than the pressure side of the impact insert 80 may be disposed adjacent the suction side 104 of the airfoil 100, as shown in Fig. 3, which may be referred to as a suction side of the impingement insert 80. In other words, the suction side of the impingement insert 80 may also be understood as a side of the impingement insert 80 for providing impingement jets 86 toward the suction side 104 of the airfoil 100.

[0078] The term 'double-walled' in the case of the impact insert 80 means that the suction side and / or the pressure side of the impact insert 80 each have two walls, namely an outer wall 82 and an inner wall 81, as in Fig. 4. Simply put, only the suction side or only the pressure side or both the suction side and the pressure side of the impact insert 80 according to the present technique has two walls - the inner wall 81 and the outer wall 82. The 'double-walled' section as used in the present technique may comprise a section, e.g., a section of the Fig. 10, which has only one wall on the suction side and only one wall on the pressure side, is not included.

[0079] To further explain, the pressure side of the impact insert 80 may comprise two walls—a pressure side inner wall 81 and a pressure side outer wall 82—thus constituting an example of the double-walled portion. Alternatively, or in addition to the above, the suction side of the impact insert 80 may comprise two walls—a suction side inner wall 81 and a suction side outer wall 82—thus constituting an example of the double-walled portion. In short, at least one of the pressure side and the suction side of the impact insert comprises the double-walled portion, while the other of the pressure side and the suction side of the impact insert may comprise a single wall or may also comprise the double-walled portion.

[0080] If both the pressure side and the suction side of the impact insert comprise a double-walled section, then the two double-walled sections can be symmetrical with respect to a chamber of the airfoil. If both the pressure side and the suction side of the impact insert comprise the double-walled section, then the two double-walled sections can be a mirror image of each other with respect to a chamber of the airfoil.

[0081] In the double-walled section of the present technique, the outer wall 82 has been referred to as 'outer' because it forms the external appearance of the impact insert 80. The inner wall 81 has been referred to as 'inner' because it is positioned within the outer wall 82 with respect to a center (not shown) of the impact insert 80 or with respect to a center (not shown) or central axis (not shown) of the cooling channel 70 defined in the airfoil 100, as in Fig. 3 is shown.

[0082] Alternatively, the terms 'inner' and 'outer' can be understood as follows: the outer wall 82 of the impact insert 80 is referred to as 'outer' because it is disposed toward the airfoil wall 101, i.e., near the pressure side 102 or the suction side 104 of the airfoil 100, when the impact insert 80 is located inside the airfoil 100. The outer wall 82 is located between the inner surface 101a of the airfoil wall 101 and the inner wall 81 of the double-walled section.

[0083] Simply put, when moving from an outer side of the impact insert 80 into the impact insert 80 from a lateral side of the impact insert 80, the outer wall 82 of the impact insert 80 appears first, followed by the inner wall 81 of the impact insert 80. Similarly, when the impact insert 80 is located within the airfoil 100, when moving from an outer side of the airfoil 100 into the airfoil 100 from a lateral side (e.g., the pressure side or suction side of the airfoil) of the airfoil 100, the airfoil wall 101 appears first, followed by the outer wall 82 of the impact insert 80, and then the inner wall 81 of the impact insert 80.

[0084] As in the Fig. As shown in Figures 4 to 6, the inner wall 81 has an inner surface 81a and an outer surface 81b, while the outer wall 82 has an inner surface 82a and an outer surface 82b. The inner surface 82a of the outer wall 82 faces the outer surface 81b of the inner wall 81. The space between the inner and outer walls 81, 82 is referred to as a central channel 502. The central channel 502 is defined or present between the inner surface 82a of the outer wall 82 and the outer surface 81b of the inner wall 81.

[0085] The outer surface 82b of the outer wall 82 is configured to face the inner surface 101a of the airfoil wall 101 when the impact insert 80 is positioned within the airfoil 100.

[0086] As in the Fig. 4 to 6, in the impact insert 80, the inner wall 81 and the outer wall 82 of the double-walled section define three spatial compartments - an inner channel 501 formed on the inner surface 81a of the inner wall 81, an outer channel 503 formed on the outer surface 82b of the outer wall 82, and the central channel 502 formed between the inner surface 82a of the outer wall 82 and the outer surface 81b of the inner wall 81.

[0087] Simply put, the central channel 502 is defined between the outer and inner walls 81, 82, the inner channel 501 is located on the inner wall side of the central channel 502, and the outer channel 503 is located on the outer wall side of the central channel 502. The central channel 502 can be arranged between the inner and outer channels 501, 502.

[0088] As in the Fig. 4 through 6, the space between the airfoil wall 101 and the outer wall 82 may be referred to as the outer channel 503 when the impact insert 80 is positioned within the airfoil 100. More specifically, the space between the airfoil wall 101 and the outer surface 82b of the outer wall 82 may be referred to as the outer channel 503. Even more specifically, the space between the inner surface 101a of the airfoil wall 101 and the outer surface 82b of the outer wall 82 may be referred to as the outer channel 503.

[0089] To further explain, when moving from a center (not shown) of the impact insert 80 toward an outer side of the impact insert 80, first the inner channel 501 appears, then the inner surface 81a of the inner wall 81 of the double-walled section, then the outer surface 81b of the inner wall 81 of the double-walled section, then the center channel 502, then the inner surface 82a of the outer wall 82 of the double-walled section, then the outer surface 82b of the outer wall 82 of the double-walled section, as shown in the Fig. 4 to 6. Continuing further, the inner surface 101a of the airfoil wall 101 would eventually appear if the impact insert 80 were positioned or disposed within the airfoil 100.

[0090] As in the Fig. 4 to 6, the impingement insert 80 includes a plurality of impingement cooling holes 85 formed as through-holes in the outer wall 82 and configured to eject the impingement jets 86 into the outer channel 503. The impingement jets 86 are formed from or by the cooling air 5 of the center channel 502. In other words, the cooling air 5 of the center channel 502 is ejected as the impingement jets 86 through the impingement cooling openings 85 into the outer channel 503. The cooling air 5 is ejected via the impingement cooling holes 85 in the form of impingement jets 86 toward the inner surface 101a of the airfoil wall 101 if the impingement insert 80 were positioned or arranged within the airfoil 100.

[0091] As in the Fig. 4 to 6, the impact insert 80 contains at least one suction line 9. The suction line 9 can be understood as a pipe or hose that extends between the outer wall 82 and the inner wall 81 via the central channel 502, ie, from the outer wall 82 via the central channel 502 to the inner wall 81. A cross-section of the suction line 9 can be circular, oval, or polygonal. The cross-section of the suction line 9 can be aerodynamically shaped, which can be oriented according to any flow of the cooling air 5 that occurs through or past the suction line 9, which will later be described with respect to Fig. 9 is explained.

[0092] As in the Fig. 5 and Fig. As shown in Figure 6, the suction line 9 has an inlet 9a, which can be arranged on the outer surface 82b of the outer wall 82. The suction line 9 has an outlet 9b, which can be arranged on the inner surface 81a of the inner wall 81. In other words, the suction line 9 fluidly connects the outer channel 503 and the inner channel 501, so that the cooling air 5 can flow from the outer channel 503 through the suction line 9 into the inner channel 501. The cooling air 5 flows from the outer channel 503 through the intermediate central channel 502 to the inner channel 501, in a restricted manner, which is restricted in the suction line 9.

[0093] Consequently, the suction line 9 operates to suck the cooling air 5 from the outer channel 503 into the inner channel 501.

[0094] It may be noted that in the present technique, the terms 'inlet' and 'outlet' and similar terms have been used with reference to cooling air. In other words, an 'inlet' may mean an 'inlet for cooling air,' and similarly, an 'outlet' may mean an 'outlet for cooling air,' unless otherwise specified.

[0095] The inlet 9a of the suction line 9 can be flush with the outer surface 82b of the outer wall 82.

[0096] Similarly, the outlet 9b of the suction line 9 may be flush with the inner surface 81a of the inner wall 81.

[0097] As in the Fig. 4 to 6, in the present technique, the double-walled section 1, 2 structurally implements a flow pattern by which the cooling air 5 is ejected from the center channel 502 as the impingement jets 86 via the impingement cooling holes 85 into the outer channel 503 to impinge on the inner surface 101a of the airfoil wall 101 and then sucked out of the outer channel 503 via the suction line 9 into the inner channel 501.

[0098] As in Fig. As shown in Figure 4, the double-walled section 1, 2 may include a main inlet 5a for the cooling air 5. The main inlet 5a may be an inlet of the central channel 502. The main inlet 5a may be the only inlet of the double-walled section 1, 2.

[0099] The cooling air 5 circulating through the double-walled section 1, 2 can enter the double-walled section 1, 2 via the main inlet 5a. In other words, the cooling air 5 circulating through the double-walled section 1, 2 can first enter the central channel 502 via the main inlet 5a, then flow to the outer channel 503 via impingement cooling holes 85, and then flow to the inner channel 501 via the exhaust line 9.

[0100] Regarding the Fig. 5 shown section M and the section in Fig. 6, two alternative embodiments, e.g., a first embodiment and a second embodiment, of the impact insert 80 of the present technique will now be described.

[0101] As stated in Section 1 of the Fig. 4 and also in Fig. 5, according to an exemplary embodiment, ie, the first embodiment, the main inlet 5a may be arranged at a top or at a bottom of the central channel 502. It may be possible for a main inlet to be present in both the bottom and the top of the central channel 502. The top and the bottom may be understood as the sides or regions of the central channel 502 that extend along the longitudinal direction A (which is also shown in the Fig. 2 and Fig. 3) of the impact insert 80. The top and bottom of the central channel 502 may correspond to the tip portion 100a and the base portion 100b of the Fig. 2. The top and bottom of the impact insert 80 may be spaced apart from each other along the longitudinal direction A, which may be understood as the same as a longitudinal direction of the impact insert 80. The cooling air 5 may enter the central channel 502 along the longitudinal direction A.

[0102] The longitudinal direction A can also be understood as the radial direction with respect to the rotation axis of the gas turbine.

[0103] Alternatively, according to another exemplary embodiment, ie, the second embodiment, the main inlet 5a may be arranged on a lateral side of the central channel 502, as shown in section 2 of the Fig. 4 and also in Fig. 6. The lateral side can be understood as extending parallel to the longitudinal direction A of the impact insert 80. The cooling air 5 can enter the central channel 502 perpendicular to the longitudinal direction A.

[0104] In accordance with aspects of the present technique, although not shown, the overall structure of the impact insert 80 may be as described for the first embodiment, e.g., as described in Section 1 of Fig. 4 and also in Fig. 5. Alternatively, in accordance with aspects of the present technique, although not shown, the overall structure of the impact insert 80 may be as described for the second embodiments, e.g., as shown in Section 2 of Fig. 4 and also in Fig. 6. Furthermore, in accordance with aspects of the present technique, as shown in Fig. 4, the overall structure of the impact insert 80 may comprise several sections - one of which, for example, may be a first section 1, as described above for the first embodiment and as shown in section 1 of Fig. 4, while another of them, for example a second section 2, may be as described above for the second embodiment and as shown in section 2 of Fig. 4 is shown.

[0105] Briefly, in an exemplary embodiment, the impact insert 80 may have a first portion 1 and a second portion 2, as shown in Fig. 4, as part of the same impact insert 80.

[0106] As further stated in Fig. 4, at the dashed line between the sections labeled '1' and '2', a main outlet 5b of the first section 1 may be fluidly connected to the main inlet 5a of the second section 2. In particular, the main outlet 5b of the first section 1 may be an outlet of the inner channel 501 of the first section 1. Consequently, the cooling air 5, as shown in Fig. 4, first enter the first section 1 and then flow from the first section 1 into the second section 2.

[0107] In particular, the cooling air 5, as in Fig. 4, first enters the central channel 502 of the first section 1 via the main inlet 5a, is then ejected as the impingement jets 86 via the impingement cooling holes 85 of the outer wall 82 of the first section 1 into the outer channel 503 of the first section 1 to impinge on the inner surface 101a of the blade wall 101 adjacent to the first section 1, and is then sucked out of the outer channel 503 of the first section 1 via the suction line 9 of the first section 1 into the inner channel 501 of the first section 1. Thereafter, the sucked-out cooling air 5 flows out of the main outlet 5b of the first section 1, which may simply be the outlet of the central channel 502 of the first section 1.

[0108] The outflowing extracted cooling air 5 then enters the second section 2 via the main inlet 5a of the second section 2. The cooling air 5 can enter the central channel 502 of the second section 2 via the main inlet 5a, is then expelled as the impingement jets 86 via the impingement cooling holes 85 of the outer wall 82 of the second section 2 into the outer channel 503 of the second section 2 to impinge on the inner surface 101a of the airfoil wall 101 adjacent to the second section 2, and is then extracted from the outer channel 503 of the second section 2 via the extraction line 9 of the second section 2 into the inner channel 501 of the second section 2. The extracted cooling air 5 can then flow out of the main outlet 5b of the second section 2, which can simply be the outlet of the central channel 502 of the second section 2.

[0109] As in Fig. 4, the impingement insert 80 may further include a third section 3, which may not be a double-walled section, but one or more walls of the third section 3 may have impingement cooling holes 85 formed therein and may form the impingement jets 86 ejected toward the inner surface 101a of the airfoil wall 101 positioned adjacent to the third section 3. The impingement jets 86 comprise the cooling air 5 flowing from the main outlet 5b of the second section 2 and into the third section 3.

[0110] Further aspects of the present technique are described below with regard to Fig. 7 and Fig. 8 has been discussed.

[0111] As in the Fig. 7 and Fig. 8, the outer wall 82 may have a corrugated shape. This may be the wall in any of the sections 1 and 2 according to Fig. 4. The corrugated shape includes a plurality of troughs 82t or notched portions extending in a direction away from the inner wall 81. One or more ribs 82r or one or more protruding portions may be located between the troughs 82t, i.e., in an alternating manner. One or more of the impingement cooling holes 85 may be housed or formed or located or arranged or provided in at least one of the troughs 82t. Preferably, all of the troughs 82t are provided with one or more of the impingement cooling holes 85.

[0112] As in the Fig. 7 and Fig. 8, the outlet 9b of the suction line 9 may be positioned on the one or more ribs 82r.

[0113] As in the Fig. 7 and Fig. 8, the inner surface 101a of the airfoil wall 101 may further include the extraction guides 99 that protrude from the inner surface 101a of the airfoil wall 101 toward the outer surface 82b of the outer wall 82 when the impact insert 80 is positioned in the airfoil 100. The extraction guides 99 may be configured, e.g., shaped and / or dimensioned, e.g., by having inclined surfaces, to guide the cooling air 5 from the outer duct 503 toward the inlet 9a of the extraction duct 9 or into the inlet 9a of the extraction duct 9.

[0114] Further aspects of the present technique are described below with regard to Fig. 9 has been discussed.

[0115] According to the present technique, a size of the inlet 9a and / or the outlet 9b of the exhaust duct 9 may be larger than a size of the impingement cooling holes 85. Here, a 'size' may be understood as the cross-sectional area. Consequently, a smooth flow of cooling air is provided through the exhaust duct 9, thus promoting the exhaust process.

[0116] Because in the present technique the cooling air 5 flows into the central channel 502 and because the suction lines 9 are arranged transversely to the central channel 502, the cooling air 5 continues to flow over or past the outer surfaces of the suction line 9, as in Fig. 9. Consequently, the suction line 9 can be aerodynamically shaped with respect to a direction of the cooling air 5 upon entering the central channel 502 and / or while flowing through the central channel 502 - whether along the longitudinal direction A (ie, the flow direction coincides with the direction A), as shown in Fig. 5 or in Section 1 after Fig. 4, or perpendicular to the longitudinal direction A (ie, the flow direction is transverse or normal to the direction A), as shown in Fig. 6 or in Section 2 after Fig. 4 is shown.

[0117] As in Fig. As shown in Figure 9, the cross-section of the exhaust duct 9 may have an oval or elliptical shape. Preferably, the long axis or longer axis of the shape is aligned with or parallel to the flow direction of the cooling air as it flows through the central channel 502.

[0118] Furthermore, as in Fig. 9, a plurality of suction lines 9 may be present, wherein the suction lines 9 may preferably be distributed evenly with respect to a distribution of the impingement cooling holes 85 in the outer wall 82. In other words, the inlets 9a of the suction lines 9 may be distributed on the outer surface 82b of the outer wall 82, preferably evenly between the impingement cooling holes 85 of the outer wall 82. As in the example according to Fig. 9, each inlet 9a of the suction lines 9 can be surrounded by several impingement cooling holes 85, for example, in Fig. 9 4 impact cooling holes 85 shown.

Claims

[1] Impact insert (80) for a turbomachinery component, the impact insert (80) comprising: - a double-walled section (1, 2) having an inner wall (81) and an outer wall (82) defining an inner channel (501) formed on an inner surface (81a) of the inner wall (81), an outer channel (503) formed on an outer surface (82b) of the outer wall (82), and a central channel (502) formed between the inner surface (82a) of the outer wall (82) and the outer surface (81b) of the inner wall (81); and - a plurality of impingement cooling holes (85) formed in the outer wall (82) and configured to eject impingement jets (86) into the outer channel (503), the impingement jets (86) being formed from the cooling air (5) of the central channel (502); characterized by that the impact insert (80) continues - at least one suction line (9) extending between the outer wall (82) and the inner wall (81) over the central channel (502) and comprising an inlet (9a) on the outer channel (503) and an outlet (9b) on the inner channel (501) for sucking cooling air (5) from the outer channel (503) into the inner channel (501), and wherein the suction line (9) is aerodynamically shaped with respect to a flow of the cooling air (5) flowing through the central channel (502). [2] Impact insert (80) according to claim 1, wherein a size of the inlet (9a) and / or the outlet (9b) of the suction line (9) is larger than a size of the impact cooling holes (85). [3] The impact insert (80) according to claim 1 or 2, wherein the outer wall (82) has a corrugated shape including a plurality of troughs (82t) extending in a direction away from the inner wall (81) and one or more ribs (82r) located between the troughs (82t); wherein one or more of the impingement cooling holes (85) are provided in at least one of the troughs (82t). [4] Impact insert (80) according to claim 3, wherein the inlet (9a) of the suction line (9) is positioned on one of the one or more ribs (82r). [5] Impact insert (80) according to one of claims 1 to 4, wherein the double-walled section (1, 2) comprises a main outlet (5b) for the cooling air (5) and wherein the main outlet (5b) is an outlet of the inner channel (501). [6] Impact insert (80) according to one of claims 1 to 5, wherein the double-walled section (1, 2) comprises at least one main inlet (5a) for the cooling air (5) and wherein the at least one main inlet (5a) is an inlet of the central channel (502). [7] Impact insert (80) according to claim 6, wherein the double-walled section (1, 2) is configured such that the cooling air (5) received via the at least one main inlet (5a) in the central channel (502) is ejected as impact jets (86) via the impact cooling holes (85) and is then sucked out of the outer channel (503) via the suction line (9) into the inner channel (501). [8] Impact insert (80) according to claim 6 or 7, wherein the main inlet (5a) is arranged on an upper side and / or a lower side of the central channel (502), the upper side and the lower side being spaced apart from one another along a longitudinal direction (A) of the impact insert (80) such that the cooling air (5) flows through the central channel (502) along the longitudinal direction (A). [9] Impact insert (80) according to claim 6, 7 or 8, wherein the main inlet (5a) is arranged on a lateral side of the central channel (502), the lateral side extending parallel to a longitudinal direction (A) of the impact insert (80) so that the cooling air (5) flows through the central channel (502) perpendicular to the longitudinal direction (A). [10] Impact insert (80) according to one of the preceding claims, wherein a cross section of the suction line (9) has one of a round shape, an oval shape and an elliptical shape. [11] Impact insert (80) according to one of claims 6 to 10, wherein the double-walled section comprises a first double-walled section (1) and a second double-walled section (2), wherein the first double-walled section (1) has the main inlet (5a) on a top side and / or a bottom side of the central channel (502) and the second double-walled section (2) has the main inlet (5a) on the lateral side of the central channel (502), or vice versa. [12] Impact insert (80) according to claim 11, wherein the first double-walled section (1) is fluidly connected to the second double-walled section (2). [13] Turbomachinery component (38, 40, 44) for a gas turbine (10), the turbomachinery component (38, 40, 44) comprising: - an airfoil (100) having an airfoil wall (101) defining an interior space (100s) of the airfoil (100); - at least one cooling channel (70) formed in the interior (100s) of the blade (100); and - an impact insert (80) inserted into the cooling channel (70), wherein the impact insert (80) is designed according to one of claims 1 to 12, and wherein the outer channel (503) is defined between the outer surface (82b) of the outer wall (82) and an inner surface (101a) of the airfoil wall (101). [14] Turbomachinery component (38, 40, 44) according to claim 13, wherein the inner surface (101a) of the airfoil wall (101) has suction guides (99) which protrude from the inner surface (101a) of the airfoil wall (101) towards the outer surface (82b) of the outer wall (82) and are configured to guide the cooling air (5), after it has impacted the inner surface (101a) of the airfoil wall (101), towards the inlet (9a) of the suction line (9). [15] A gas turbine (10) comprising a turbomachine component, wherein the turbomachine component (38, 40, 44) is configured according to claim 13 or 14.

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