Turbine blade for gas turbine

Abstract

Turbine blade (TV) for a gas turbine, comprising:-an airfoil (AF) having a suction side wall (SSW) and a pressure side wall (PSW) surrounding at least one central cavity (CC), the two walls (SSW, PSW) extending in an axial direction (X) of the gas turbine from a leading edge (LE) to a trailing edge (TE) when the turbine blade (TV) is assembled in the gas turbine, the invention relates to a gas turbine comprising:-an airfoil (AF) extending in a radial direction (Y) of said gas turbine from an outer end (OE) of the airfoil (AF) to an inner end (IE) of the airfoil (AF) for guiding hot gases of the gas turbine,-an outer platform (OP) and an inner platform (IP), each located at a respective end (OE, IE) of the airfoil (AF) and each having a hot gas surface (HGS) facing the airfoil (AF) and an inner cold gas surface (CGS) arranged opposite the hot gas surface (HGS), -a plurality of cooling channels (CMC) arranged in the suction side wall (SSW) and / or the pressure side wall (PSW), the cooling channels (CMC) extending substantially in the radial direction (X), each cooling channel (CMC) having at least one channel inlet (CI, FCI) and one channel outlet (CO) through which the coolant (CM) can enter and exit the respective cooling channel (CMC), respectively, the channel inlet (CI, CFI) is in flow connection with the at least one coolant supply chamber (CMSC) and the channel outlet (CO) is in flow connection with the at least one coolant discharge chamber (CMDC),-wherein, for a respective cooling channel (CMC), a first channel inlet (FCI) of the at least one channel inlet (CI, FCI) is in flow connection with one of the at least one coolant supply chamber (CMSC) and a second channel inlet (FCI) of the at least one coolant discharge chamber (CMDC) is in flow connection with the other of the at least one coolant discharge chamber (CMDC). The first channel inlets (CI, FCI) and the channel outlets (CO) of the plurality of cooling channels (CMC) are in flow connection with one of the at least one coolant discharge chamber (CMDS),-wherein the first channel inlets (CI, FCI) and the channel outlets (CO) of the plurality of cooling channels (CMC) are arranged such that the flow directions of directly adjacent cooling channels (CMC) are opposite for most of the cooling channels (CMC), preferably for all of the cooling channels (CMC), and-wherein the first channel inlets (CI, FCI) and the channel outlets (CO) of the plurality of cooling channels (CMC) are in flow connection with one of the at least one coolant discharge chamber (CMDS). And thereby a device for reducing the risk of particles clogging a cooling channel (CMC) is provided.

Description

Technical Field

[0001] The present invention relates to a turbine blade, comprising:

[0002] An airfoil, an outer platform and an inner platform for guiding hot gases of a gas turbine, and a plurality of cooling channels arranged in a suction side wall and / or a pressure side wall, wherein when the blade is installed in a gas turbine, the cooling channels extend in a radial direction of the blade, wherein each cooling channel has at least one channel inlet and one channel outlet, and coolant can enter and leave the corresponding cooling channel through at least one channel inlet and one channel outlet, respectively, the channel inlet is flow-connected to at least one coolant supply chamber, and the channel outlet is flow-connected to at least one coolant discharge chamber. Background Art

[0003] Turbine blades as described above are known from US 5,394,687 and EP 1101900 A1. According to EP1101900 A1, those radially extending cooling channels are cast or drilled. The use of casting cores for producing radial cooling channels is quite complicated and leads to particularly high scrap rates. Therefore, the implementation of such cooling designs in turbine components is limited by traditional methods (such as investment casting). The drilling of the cooling holes is limited to airfoils with a straight configuration in the radial direction, but this configuration is not conventional now, but only an exception. Modern airfoils for turbine blades include airfoil designs with 3D shapes such as bows, swept shapes, etc., in which drilling is not an option when producing so-called near-wall airfoil cooling holes.

[0004] Furthermore, it is also known to employ impingement cooling fins inside turbine blades to achieve a high cooling effect. However, those and other modular designs of turbine blades require the manufacture of additional components and subsequent joining processes, which reduces reliability and increases costs.

[0005] EP 1 101 900 A1 teaches having a plurality of axial zones with synchronized cooling flow directions, ie the flow direction in a plurality of adjacent cooling holes is the same and only changes from one zone to another, each zone comprising a plurality of cooling holes.

[0006] Therefore, there is a need for a turbine blade that overcomes the shortcomings of the prior art: the cooling scheme of the turbine blade must reliably achieve the maximum hot gas temperature without shortening the life of the turbine blade, also when manufactured by additive manufacturing methods.

[0007] It is an object of the present invention to provide an advantageous turbine blade which is easy to manufacture and which allows increasing the hot gas temperature and / or increasing the operating life of the turbine blade. Summary of the invention

[0008] This problem is solved by a turbine blade according to the features of the independent claim. Advantageous embodiments are the subject matter of the dependent claims and the following description. Their respective features can be combined arbitrarily if not stated otherwise.

[0009] In detail, the problem is solved by a turbine blade for a gas turbine, the turbine blade comprising an airfoil, an outer platform and an inner platform, and a plurality of cooling channels, the airfoil having a suction side wall and a pressure side wall surrounding at least one central cavity, the two walls extending from a leading edge to a trailing edge in an axial direction of the gas turbine and extending from an outer end of the airfoil to an inner end of the airfoil in a radial direction of the gas turbine for guiding hot gases of the gas turbine, the outer platform and the inner platform being each located at a respective end of the airfoil and each having a hot gas surface facing the airfoil and an inner cold gas surface arranged opposite to the hot gas surface, the plurality of cooling channels being arranged on the suction side wall and / or the pressure side wall. In the wall, cooling channels extend substantially in radial direction, wherein each cooling channel has at least one channel inlet and one channel outlet, through which coolant can enter and leave the corresponding cooling channel respectively, wherein, for the corresponding cooling channel, a first channel inlet of the at least one channel inlet of the corresponding cooling channel is flow-connected to one of the at least one coolant supply chambers, and the channel outlet of the corresponding cooling channel is flow-connected to one of the at least one coolant discharge chambers, wherein the first channel inlets and the channel outlets of the plurality of cooling channels are arranged such that for most of the cooling channels, preferably for all of the cooling channels, the flow directions of directly adjacent cooling channels are opposite.

[0010] In other words: for nearly all or all radially extending cooling channels in the airfoil wall, the flow direction of the immediately adjacent cooling channels is opposite and alternates from cooling channel to cooling channel.

[0011] The inventors have realised that the counterflow cooling scheme of the airfoil of the present invention is more advantageous than the cooling scheme known from EP 1101109 A1, in which a larger area with a plurality of cooling holes has the same flow direction. When flowing along the cooling holes, the coolant is heated and leaves the corresponding heated cooling holes. The cooling schemes of the prior art result in an unbalanced temperature distribution in the radial direction: the cooling effect at the supply end is greater than the cooling effect at the airfoil end, where the coolant (e.g. cooling air) leaves the airfoil. By repeatedly changing the flow direction of the coolant from cooling channel to cooling channel, it can be seen that the cooling effect in the radial direction is balanced, e.g. with lower temperature gradients in the airfoil wall, a reduced ratio between the maximum metal temperature and the average value. As a result, higher hot gas temperatures and / or a higher turbine blade life can be achieved. In other words: the advantage of the counterflow cooling scheme of the airfoil is the realization and result of a simple design of a more uniform airfoil metal temperature field.

[0012] Another advantage of the proposed airfoil counterflow cooling scheme is that the outer shape of the airfoil can be freely adapted to the aerodynamic requirements without any restrictions on heat transfer and cooling, especially when additively manufacturing turbine blades. The cooling channels mainly have a straight line configuration starting from the outer end and ending at the inner end (or vice versa), but in the lateral direction they can maintain a constant distance from the hot gas surface of the airfoil.

[0013] Furthermore, at least one device for reducing the risk of particles clogging the cooling channels is provided. Providing such a device in the turbine blade increases the life of the turbine blade and ensures safe operation of the gas turbine when equipped with such a turbine blade. By incorporating the device for reducing the risk of particles clogging the cooling channels directly into the component itself, an additional element for this purpose is introduced without changing other elements of the gas turbine. Existing structures for cleaning (cooling) air, such as particle filters located in the air intake system or inside the gas turbine, remain unchanged.

[0014] In a preferred embodiment of the invention, at least one coolant supply chamber is partially defined by the inner cold gas surface of the inner platform or by the inner cold gas surface of the outer platform, and at least one coolant discharge chamber is embodied as at least one central cavity. This design is simple, robust and enables a compact turbine blade.

[0015] In order to repeatedly change the flow direction of the coolant from the channel to the adjacent channel, the first channel inlet of each second cooling channel is arranged in the inner cold gas surface of the outer platform, and the first channel inlet of the alternating cooling channel located between two of each second cooling channel is arranged in the inner cold gas surface of the inner platform, and wherein the channel outlet of each second cooling channel is arranged at the inner end of the airfoil, and the channel outlet of the alternating cooling channel located between two of each second cooling channel is arranged at the outer end of the airfoil.

[0016] According to another preferred embodiment, nearly all or all cooling channels located along the suction sidewall and around the leading edge are elliptical or oval in cross section with a first average pitch between the cooling channels, and / or nearly all or all cooling channels located in the pressure sidewall are circular in cross section with a second average pitch between the cooling channels, wherein preferably the first average pitch is smaller than the second average pitch. Here, the shape of the cooling channels is adapted to local needs with regard to heat transfer, mechanical integrity and life of the turbine blade.

[0017] In order to enhance the trailing edge cooling of the airfoil and the effective discharge of coolant from the airfoil, the airfoil includes at least one discharge cooling hole located at or in the trailing edge, and wherein a cooling pin array and / or axially extending reinforcing ribs are arranged between at least one coolant discharge chamber and at least one discharge cooling hole.

[0018] Preferably, the axial length of the airfoil is determined between the leading edge and the trailing edge of the airfoil, and wherein the airfoil is free of film cooling holes over at least 85% of its axial length starting from its leading edge. Needless to say, by omitting the drilling of film cooling holes, the manufacturing process is significantly easier and no air mixing losses occur in the operation of the turbine blade of the invention. Furthermore, the amount of coolant required can be reduced by about 40% compared to turbine blades with leading edge, suction side and / or pressure side film cooling. In addition to this, in the absence of film cooling, the coolant supply pressure can be reduced, since the required discharge pressure of the coolant including the return flow margin is defined by the hot gas pressure at the trailing edge of the airfoil, rather than by the hot gas pressure at the leading edge of the airfoil when film cooling is employed. The savings in coolant pressure can be used to improve efficiency and / or for the introduction of additional anti-clogging systems, for example, dust collectors embedded in the turbine blades.

[0019] In order to provide a compact turbine blade, the inner platform and / or the outer platform each enclose a respective coolant supply chamber.Such a turbine blade structure may advantageously be manufactured by means of an additive manufacturing method, in particular by laser assisted powder bed fusion (LPDF).

[0020] Further preferably, the coolant discharge chamber is separated from the coolant supply chamber by a partition wall.Therefore, the rigidity of the region where the airfoil is joined to the platform can be improved.

[0021] In a very preferred embodiment, the device is embodied as a dust collector arranged in the inner platform and / or the outer platform. Analysis shows that the dust collector has a high efficiency and is able to remove about 90% of particles up to 1 mm in size. Before the cleaned coolant is directed to the cooling channel for cooling the airfoil, the dust collector removes at least most or almost all of the dust particles carried by the coolant. Supplying cleaned coolant allows the cooling channel to have a smaller cross-section and a lower risk of being blocked by particles adhering therein. As a result, the life of the turbine blades can be increased and / or further improved cooling by further reducing the pitch of the cooling channel can be achieved.

[0022] In a preferred embodiment of the dust collector, the dust inertial separator is a duct having a plurality of clean coolant outlet holes and extending in the same direction as the coolant leaves the air acceleration zone. No rework or additional components are required. Because the dust separation system is integrated into the platform, the size of the turbine blade of the invention does not exceed the size of the existing components. Therefore, a retrofittable turbine blade of the invention can be provided.

[0023] In addition to or as an alternative embodiment of the dust collector, the device can be embodied as a bypass channel that fluidly connects two cooling channels whose first channel inlets are directly adjacent to each other. Depending on the number of bypass channels that terminate at the respective cooling channels, the respective cooling channels include one or two second channel inlets for coolant. In the event that larger particles carried by the coolant (e.g., cooling air) may block the channel inlet, the coolant can be supplied to the respective cooling channels via the bypass channel. This improves the fault protection capability, durability and service life of the turbine blades.

[0024] Preferably, the turbine blade is one-piece and is manufactured by an additive manufacturing process, preferably a powder bed fusion process, in particular a laser assisted powder bed fusion process (LPBF).

[0025] The previous description of advantageous embodiments of the invention contains many features which are partially combined with one another in the dependent claims. Where appropriate, these features can also be considered individually, as well as combined with one another in further suitable combinations. Furthermore, method features which are expressed as device features can be considered as component features, and accordingly, component features which are expressed as process features can be considered as method features.

[0026] The above-mentioned characteristics, features and advantages of the present invention and their implementation methods can be more clearly understood in conjunction with the description of the exemplary embodiments to be explained below with reference to the accompanying drawings. The exemplary embodiments are intended to illustrate the present invention, but should not limit the scope of the present invention to the combination of features given therein, nor should the scope of the present invention be limited to functional features. In addition, the appropriate features of each exemplary embodiment may also be explicitly considered separately, removed from one exemplary embodiment, introduced into another exemplary embodiment and / or combined with any attached claims. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] The present invention will be described with reference to the accompanying drawings, in which:

[0028] Figure 1 shows a cross-sectional view of a turbine blade according to a first exemplary embodiment of the invention,

[0029] Figure 2 Shows Figure 1 First details about the outer end of the airfoil of the turbine blade,

[0030] Figure 3 Shows Figure 1 First details about the inner end of the airfoil of the turbine blade,

[0031] Figure 4 Shown throughout Figure 1 A cross-sectional view of the airfoil of a turbine blade shown in,

[0032] Figure 5 , Figure 6 and Figure 7 shows a number of cross-sectional views through an outer platform of a turbine blade according to a second exemplary embodiment of the invention, and

[0033] Figure 8 A cross-sectional view through a turbine blade according to another exemplary embodiment of the invention is shown. DETAILED DESCRIPTION

[0034] Figure 1 A turbine blade TV for a gas turbine is shown in a cross-sectional view. The turbine blade TV is manufactured by an additive manufacturing process such as laser-assisted powder bed fusion (LPBF) and is therefore embodied in one piece.

[0035] The turbine blade TV comprises as a main body an airfoil AF, the aerodynamic shape of which (see Figure 3) is an arcuate teardrop. When the turbine blade TV is assembled in a gas turbine, the airfoil AF (more precisely, the suction side wall SSW of the airfoil AF and the pressure side wall PSW of the airfoil AF) extends from the outer end OE to the inner end IE in the radial direction Y of the gas turbine. The two walls SSW and PSW also extend from the leading edge LE of the airfoil AF to the trailing edge TE in the axial direction of the gas turbine. During operation of the gas turbine, the airfoil AF is subjected to a hot gas medium to drive the gas turbine.

[0036] An outer platform OP and an inner platform IP are arranged at the outer end OE and the inner end IE of the airfoil AF, respectively. Each platform IP, OP is hollow and includes a coolant supply chamber CMSC. According to the exemplary embodiment shown, on the side of the platform opposite to the airfoil AF, each platform IP, OP includes a coolant blade inlet CMI that is fluidically connected to the coolant supply chamber CMSC. The turbine blade TV shown includes two coolant blade inlets CMI. Therefore, the outer platform OP and the inner platform IP have a double-walled configuration, wherein each platform includes a hot gas surface HGS facing the airfoil AF and an inner cold gas surface CGS arranged opposite to the hot gas surface HGS.

[0037] A plurality of cooling channels CMC are arranged in the airfoil AF. They extend mainly in a straight line from the inner end IE to the outer end OE, or vice versa, and they are distributed along a large part of the periphery of the airfoil in the suction side wall SSW, at the leading edge and / or in the pressure side wall PSW of the airfoil (see Figure 3 ). Each of these cooling channels CMC has at least one channel inlet CI and one channel outlet CO. Figure 3 In the exemplary embodiment shown, each of the cooling channels CMC has only one channel inlet CI, namely the first channel inlet FCI.

[0038] The first channel inlet FCI is located in the inner cold gas surface CGS of the platform IP, OP, so that the coolant supply chamber CMSC is in flow connection with the first channel inlet FCI (see Figure 2 ). The channel outlet CO is arranged in the inner surface of the airfoil AF, near the outer platform OP or near the inner platform IP. All cooling channels CMC of the airfoil AF are arranged at a small distance of about 1 mm from the hot gas surface of the airfoil.

[0039] The first channel inlet FCI and the channel outlet CO of two directly adjacent cooling channels CMC are arranged so that the flow direction of the coolant in the two corresponding cooling channels is opposite. Therefore, each second cooling channel guides the coolant from the outer end OE of the airfoil AF to its inner end IE, and the remaining cooling channels CMC (i.e., the alternating cooling channels CMC) guide the coolant from the inner end IE of the airfoil AF to its outer end OE.

[0040] All channel outlets CO of the cooling channels CMC of the airfoil are flow-connected to at least one coolant discharge chamber CMDC. A series of radially distributed reinforcing ribs SR and a cooling pin array ACP are arranged between at least one central cavity CC and the trailing edge TE. The trailing edge comprises at least one coolant discharge cooling hole DCH, for example as a centerline discharge hole. Alternatively, the turbine blade TV (and accordingly, the airfoil AF of the turbine blade TV) may comprise a plurality of discharge cooling holes DCH in the trailing edge TE. When located near or at the trailing edge, the plurality of discharge cooling holes DCH may be embodied as cut-back openings (not shown). Each reinforcing rib SR connects the suction side wall SSW and the pressure side wall PSW, and extends in the axial direction X for guiding the coolant from the coolant discharge chamber CMDC to the cooling pin array ACP in an appropriate manner.

[0041] according to Figure 1 , Figure 2 , Figure 3 and Figure 4 In the first exemplary embodiment shown, the central chamber CC is a coolant discharge chamber CMDC. The central chamber CC is therefore separated from the cooling medium supply chamber CMSC by a partition wall SP of each respective platform IP, OP. In this case, the turbine blade comprises two coolant inlets CMI, one coolant inlet CMI at the outer platform OP and the other coolant inlet CMI at the inner platform IP.

[0042] The partition wall SP also separates a first channel inlet FCI of the first cooling channel CMC from a channel outlet CO of a second cooling channel CMC directly adjacent to the first cooling channel CMC.

[0043] According to an alternative embodiment of the invention (not shown), the turbine TV may include only one coolant vane inlet CMI at its inner platform IP or at the outer platform IP. In this case, the airfoil AF may include two central cavities, wherein the first central cavity is embodied as a cooling discharge chamber CMDC, and wherein the second central cavity is embodied as a radially extending tube for guiding unused coolant to the opposite end of the turbine blade TV, where there is no coolant vane inlet. Then, the airfoil cooling of the invention including the airfoil counterflow cooling scheme with cooling channels CMC can also be used for turbine blades TV having only one coolant vane inlet or coolant supply.

[0044] The spacing between two directly adjacent cooling channels CMC is defined as the pitch. The average pitch of all cooling channels located in the leading edge LE and in the suction side wall SSW may be smaller than the average pitch of all cooling channels CMC located in the pressure side wall PSW. Moreover, the shape of the cooling channel depends on its position. In the suction side wall SSW and / or in the leading edge LE, the shape of the cooling channels CMC is elliptical or oval, while the cooling channels CMC located in the pressure side wall PSW are circular. Both the local pitch adjustment and the shape of the cooling channels arranged in this way enable local adaptive cooling of the walls SSW, PSW of the airfoil according to the typical temperatures distributed on the airfoil AF occurring during operation, without any overcooling of the walls and any undercooling.

[0045] Figures 5 to 7 The outer platform OP of a turbine blade TV according to a second exemplary embodiment of the present invention is shown in different cross-sectional views. In order to supply a cleaner coolant to the cooling channel CMC, that is, the coolant carries a smaller amount of dust or particles, for example, cooling air, a dust collector DP is arranged between the coolant blade inlet CMI and the first channel inlet FCI. Preferably, the dust collector DP includes an air acceleration area AAZ, a dust inertial separator DIS and a dust trap DT with a dust outlet hole DEH. The dust inertial separator DIS is embodied as a duct with a constant distribution of clean coolant outlet holes CCMHEH, and extends in the same direction as the coolant leaves the air acceleration area AAZ. When in operation, the accelerated coolant and its dust and particles flow into and through the duct. Due to mass and inertia, the coolant can change direction more easily than the dust and particles carried therein. Dust and particles continuously flow directly into the dust trap TD, while the cleaner coolant changes the flow direction and leaves the duct through the clean coolant outlet hole CCMHEH. Compared to turbine blades without dust collector DP, cleaner coolant flows to the cooling channels and the risk of clogging or plugging of the cooling channels CMC is reduced. This ensures that the expected life of the turbine blade can be achieved.

[0046] Figure 8 A perspective view of an inventive airfoil counterflow cooling scheme for a turbine blade TV according to another exemplary embodiment of the invention is shown. Here, a device for reducing the risk of particles DP blocking the cooling channel CMC is embedded in the airfoil AF. In detail, a plurality of bypass channels BP are located near the first channel inlet FCI of the cooling channel CMC. In the event that the first channel inlet FCI is blocked or clogged by particles DP, the coolant CM can be supplied to the corresponding cooling channel CMC via the two bypass channels shown. In this case, the cooling channel CMC comprises three channel inlets: the first channel inlet FCI represents the first channel inlet, and the two bypass channels BC represent the second channel inlet and the third channel inlet. Of course, it is possible to Figure 8In contrast to that shown in , only one second bypass channel BC is provided. In this case, the cooling channel (CMC) has only two channel inlets.

Claims

1. A turbine blade (TV) for a gas turbine, the turbine blade (TV) comprising: an airfoil (AF) having a suction side wall (SSW) and a pressure side wall (PSW) enclosing at least one central cavity (CC), the two walls (SSW, PSW) extending in an axial direction (X) of the gas turbine from a leading edge (LE) to a trailing edge (TE) and in a radial direction (Y) of the gas turbine from an outer end (OE) of the airfoil (AF) to an inner end (IE) of the airfoil (AF) for guiding hot gases of the gas turbine, when the turbine blade (TV) is assembled in a gas turbine, an outer platform (OP) and an inner platform (IP), each of which is located at a respective end (OE, IE) of the airfoil (AF) and each of which has a hot gas surface (HGS) facing the airfoil (AF) and an inner cold gas surface (CGS) arranged opposite the hot gas surface (HGS), a plurality of cooling channels (CMC), said plurality of cooling channels (CMC) being arranged in said suction side wall (SSW) and / or in said pressure side wall (PSW), said cooling channels (CMC) extending substantially in a radial direction (X), wherein each cooling channel (CMC) has at least one channel inlet (CI, FCI) and one channel outlet (CO), through which a coolant (CM) can enter and leave the respective cooling channel (CMC), respectively, - the channel inlet (CI, CFI) is in flow connection with at least one coolant supply chamber (CMSC) and the channel outlet (CO) is in flow connection with at least one coolant discharge chamber (CMDC), - wherein, for a respective cooling channel (CMC), a first channel inlet (FCI) of the at least one channel inlet (CI, FCI) of the cooling channel (CMC) is in flow connection with one of the at least one coolant supply chambers (CMSC), and a channel outlet (CO) of the cooling channel (CMC) is in flow connection with one of the at least one coolant discharge chambers (CMDS), wherein the first channel inlets (CI, FCI) and the channel outlets (CO) of the plurality of cooling channels (CMC) are arranged such that for most cooling channels (CMC), preferably for all cooling channels (CMC), the flow directions of directly adjacent cooling channels (CMC) are opposite, It is characterized in that Means are provided for reducing the risk of particles clogging the cooling channels (CMC).

2. Turbine blade (TV) according to claim 1, in, The at least one coolant supply chamber (CMSC) is partially defined by an inner cold gas surface (CGS) of the inner platform (IP) or the outer platform (OP), and the at least one coolant discharge chamber (CMDC) is embodied as the at least one central cavity (CC).

3. Turbine blade (TV) according to one of the preceding claims, in, Along the axial direction (X), a first channel inlet (FCI) of each second cooling channel is arranged in an inner cold gas surface (CGS) of the outer platform (OP), and a first channel inlet (FCI) of an alternating cooling channel located between two of the second cooling channels is arranged in an inner cold gas surface (CGS) of the inner platform (IP), and wherein a channel outlet (CO) of each second cooling channel is arranged at an inner end (IE) of the airfoil (AF), and a channel outlet (CO) of an alternating cooling channel located between two of the second cooling channels is arranged at an outer end (OE) of the airfoil (AF).

4. Turbine blade (TV) according to one of the preceding claims, in, Almost all or all of the cooling channels (CMCs) located in the suction side wall (SSW) and around the leading edge (LE) are elliptical or oval in cross section, with a first average pitch between the cooling channels (CMCs), and / or almost all or all of the cooling channels (CMCs) located in the pressure side wall (PSW) are circular in cross section, with a second average pitch between the cooling channels (CMCs), wherein preferably, the first average pitch is smaller than the second average pitch.

5. Turbine blade (TV) according to one of the preceding claims, in, The airfoil (AF) comprises at least one discharge cooling hole (DCH) at or in the trailing edge (TE), and Therein, a cooling pin array (ACP) and / or an axially extending reinforcement rib (SR) are arranged between the at least one coolant discharge cavity (CMDC) and the at least one discharge cooling hole (DCH).

6. Turbine blade (TV) according to one of the preceding claims, in, An axial length (AL) of the airfoil (AF) is determined between a leading edge (LE) and a trailing edge (TE) of the airfoil (AF), and Wherein the airfoil (AF) is free of film-cooling holes over at least 85% of an axial length from a leading edge (LE) of the airfoil (AF).

7. Turbine blade (TV) according to one of the preceding claims, in, The inner platform (IP) and / or the outer platform (OP) each encloses a respective coolant supply chamber (CMSC).

8. Turbine blade (TV) according to one of the preceding claims, in, A partition wall (SW) partitions the coolant discharge chamber (CMDC) from the coolant supply chamber (CMSC).

9. Turbine blade (TV) according to claim 8, in, The device is embodied as a dust collector (DP), which is arranged in the inner platform (IP) and / or the outer platform (OP).

10. Turbine blade (TV) according to claim 9, in, The dust collector (DP) comprises an air acceleration zone (AAZ), a dust inertial separator (DIS) and a dust trap (DT) having a dust exit hole (DEH).

11. Turbine blade (TV) according to claim 10, in, The dust inertial separator (DIS) is embodied as a duct having a plurality of clean coolant outlet holes (CCMEH) and extending in the same direction as the coolant leaves the air acceleration zone (AAZ).

12. Turbine blade (TV) according to one of the preceding claims, in, The turbine blade (TV) is in one piece.

13. A method for additive manufacturing of a turbine blade (TV) according to one of the preceding claims, The method is a powder bed fusion process, in particular laser assisted powder bed fusion (LPBF).

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