Stationary gas turbine blades

Abstract

The present invention relates to a turbine blade having an airfoil portion (18), in which a first cooling passage (30) for a first cooling medium flow (M1) and a second cooling passage (50) for a second cooling medium flow (M2) are formed, the first cooling passage (30) including a first cooling medium flow path (32) for cyclone cooling of a leading edge (24) and a second cooling medium flow path (50) connected to the first cooling medium flow path (32) and extending below a blade tip (22). The second cooling passage (50) includes a second coolant flow path (34) extending from the edge (24) toward the trailing edge (26), the second cooling passage (50) including a serpentine coolant flow path (52) for cooling a central region (48) located chordwise aft of the leading edge region (39) of the airfoil (18), and a first trailing edge coolant flow path (54) for at least partially cooling a trailing edge region (59) chordwise aft of the central region (48) of the airfoil (18) and extending to the trailing edge.

Description

[Technical field]

[0001] The invention relates to a turbine blade according to the preamble of claim 1.

[0002] Turbine blades of gas turbines are subjected to the greatest thermal and mechanical loads during operation and are therefore today designed to be coolable and particularly robust by means of a complex, hollow internal geometry.

[0003] For example, a gas turbine blade corresponding to the preamble of claim 1 is known from US Pat. No. 5,999,333, in which the cooling air introduced tangentially into the leading edge cooling channel allows the leading edge to be cooled without the need for further film cooling holes therein, often called Showerhead-Holes in English, for cooling the leading edge. However, a significant part of the cooling air flowing in the leading edge cooling channel leaves the turbine blade via film cooling holes, also called Gill-Holes in English, arranged on the suction side near the leading edge, while the remaining part of this cooling air is led to the trailing edge at the bottom of the blade tip. On the other hand, the remaining part of the airfoil is cooled by trailing edge jets via serpentine cooling channels.

[0004] Furthermore, from the patent DE 10 200 03 133 A1 a so-called multi-wall turbine blade is known, which is also called in English "Multiwall-Turbine Blade". Inside it, two stoppers are provided, by means of which the cooling air flowing inside the turbine blade is forced in the immediate vicinity of the inner surface of the outer wall. Furthermore, the patent DE 10 200 03 133 A1 shows an alternative configuration of a multi-wall turbine blade. Furthermore, the patent DE 10 200 03 133 A1 shows a turbine blade with two adjacent serpentine cooling channels, based on the blade span direction, which are connected in series via a channel cooling the leading edge.

[0005] In the quest for ever more efficient turbines, there is a continuing need to conserve cooling air, since this can be used as primary air for the oxidation of fossil or synthetic fuels, thereby increasing efficiency. [Prior art documents] [Patent documents]

[0006] [Patent Document 1] International Publication No. 1996 / 15358A1 Brochure [Patent Document 2] International Publication No. 2017 / 039571A1 Brochure [Patent Document 3] European Patent Application Publication No. 1783327A2 [Patent Document 4] US Patent Application Publication No. 2010 / 0239431A1 Summary of the Invention [Problem to be solved by the invention]

[0007] SUMMARY OF THE PRESENT EMBODIMENT An object of the present invention is therefore to provide a turbine blade which consumes a further reduced amount of cooling medium and has a long service life. [Means for solving the problem]

[0008] This object is achieved according to the invention by a turbine blade as defined in claim 1. The invention proposes a turbine blade, in particular for an axial-flow stationary gas turbine, in particular for one of the high-pressure turbine stages, in which a cooling system is arranged, which comprises a first cooling passage for a first cooling medium flow and a second cooling passage for a second cooling medium flow, which is substantially, and preferably completely, separated from the first cooling passage, in which the first cooling passage comprises a first cooling medium passage provided for cyclonic cooling of the leading edge and a second cooling medium passage connected to the first cooling medium passage and running in the direction from the leading edge to the trailing edge under the blade tip, in which the second cooling passage comprises a serpentine cooling passage for cooling a central region arranged chordwise aft of the leading edge region of the airfoil. The cooling passage further comprises a cooling medium flow passage, a first trailing edge cooling medium flow passage arranged chordwise aft of the central region of the airfoil and for at least partially cooling the trailing edge region reaching the trailing edge, the first trailing edge cooling medium flow passage being fluidly connected to a plurality of first outlet holes arranged at the trailing edge, the first cooling medium flow passage being provided for cooling without outlet holes, i.e., for locally closed cooling, the first cooling passage further comprises a third cooling medium flow passage connected to the second cooling medium flow passage and extending mainly radially inward, and a second trailing edge cooling medium flow passage connected to the third cooling medium flow passage, the second trailing edge cooling medium flow passage being formed for cooling the blade tip region of the trailing edge region and being fluidly connected to a plurality of second outlet holes arranged at the trailing edge.

[0009] The invention is based on the finding that a significant saving in cooling medium for cooling a turbine blade can only be achieved if the leading edge and / or the pressure and / or suction side walls of the airfoil are devoid of openings through which the cooling medium can exit and into which the hot gas flowing around the turbine blade can enter. To allow for this simple construction of the turbine blade, the cooling medium exits at least from the trailing edge and possibly also from the blade tip facing outward. In this respect, to provide locally closed cooling, only those passages and channels should be provided that are able to cool the leading edge and most of the pressure and suction sides of the airfoil. In other words, no showerhead holes, no gill holes or other film cooling holes branch off from the first and / or serpentine cooling medium passages, these passages have no exit holes. Exit holes are provided only at the trailing edge and possibly also at the blade tip. Locally closed cooling does not mean that the cooling medium does not exit at all from the airfoil into the hot gas.

[0010] Nevertheless, in order to achieve sufficient cooling, especially of the leading edge of a turbine blade, which is subject to extremely high thermal loads, the demand for cooling medium actually increases in the case of locally closed, i.e. without outflow holes, leading edge cooling. However, according to the invention, it is proposed for the first time that the first cooling medium flow used for cooling the leading edge is also used for cooling the radially outer part of the trailing edge region of the airfoil. Instead of discharging the cooling medium via the gill holes and directly from the trailing edge, as in the prior art, according to the invention, a rear separating rib is introduced into the system, which redirects the cooling medium coming from this system again inwards and then leads it to another trailing edge cooling medium channel. As a result, the first cooling medium flow is guided via a second cooling medium channel that runs directly below the blade tip to the rear end of the airfoil, and further via a third cooling medium channel that follows this, preferably to a height of about half the height of the trailing edge, and then effectively used in the trailing edge cooling medium channel arranged radially outwards. This solution allows a significant reduction in the demand for cooling air for the second fluid passage. The approach proposed here therefore provides for maximum utilization of the available cooling medium by a novel distribution and by using a cooling concept, namely cyclone cooling, which was previously considered totally inappropriate for the turbine blades of the first and / or second turbine stages of gas turbines with high compressor pressure ratios and high turbine inlet temperatures and was therefore not considered for these turbine blades.

[0011] Cyclonic cooling means cooling in which the main part of the cooling medium flowing in the cooling channel or in the cooling medium passage flows in a twisted manner from the main cooling medium inlet to the main outlet. Twisting means that a substantial part of the cooling medium flows in a spiral or helical manner along the channel or passage in question. This twisted flow should be distinguished from a turbulent flow. The latter is regularly caused by so-called turbulators and therefore occurs in a spatially very limited area, since only a very small proportion of the cooling medium is influenced and manipulated by the turbulators. After leaving the area in question, the turbulence breaks down again. Thus, the twisted main flow may locally have a secondary turbulent component in a very small area, but not vice versa.

[0012] According to the invention, the cooling medium consumption can be reduced to an unexpected extent, while at the same time the entire airfoil is sufficiently cooled. According to detailed simulations, this is true even for turbine blades in one of the two front turbine stages of stationary gas turbines with turbine inlet temperatures of 1300° C. or higher in normal operation according to ISO or compressor pressure ratios of 19:1 or higher. Even for such turbine blades, the amount of cooling medium can be reduced by about 30% compared to the conventional solution with multiple cooling holes arranged at the leading edge, while at the same time achieving a comparable service life.

[0013] According to another particularly preferred embodiment of the invention, one or more outlet holes for the cooling medium are arranged in the blade tip, which are fluidly connected to the second cooling medium channel, as a result of which the fatigue strength of the protruding edge of the blade tip, for example of the grazing edge, is improved.

[0014] In another preferred embodiment, the first cooling passage comprises a supply passage for the first cooling medium passage, which is arranged immediately next to the first cooling medium passage, extends over at least a large part of the span of the airfoil and is fluidly connected to the first cooling medium passage via a number of through-openings, which have means for imparting or enhancing the twist of the cooling medium flowing in the first cooling medium passage, by means of a special orientation of the through-openings. For example, if the through-openings are arranged tangentially, i.e. eccentrically, in the first cooling medium passage, in particular aligned with the inner surface of the suction or pressure side wall and / or radially, the twist required for cyclonic cooling can be imparted or enhanced by simple means to the cooling medium flowing in the first cooling medium passage. In this way, efficient cyclonic cooling of the leading edge can be achieved relatively easily.

[0015] According to another embodiment, the density of through openings that can be defined in the span direction is maximum at the root end and preferably decreases stepwise or continuously towards the tip, thereby adjusting and equalizing the cyclonic cooling of the leading edge over the height of the airfoil. In this way, the flow velocity in the first coolant passage can be kept approximately constant over the span of the airfoil, which can be achieved by the first coolant passage tapering in cross section also towards the tip.

[0016] According to another advantageous embodiment, in order to locally further increase the heat transfer to the first and / or second cooling medium and / or to support twisting, a plurality of preferably rib-shaped, in particular inclined turbulators are arranged on one or more inner surfaces of one or more cooling medium channels.

[0017] According to a further development of the invention, in each trailing edge coolant passage, a number of pedestals arranged in a patterned array are provided. In this way, the trailing edge regions of the suction and pressure sides of the airfoil, which continue from the central region of the airfoil to the trailing edge of the airfoil, can be cooled in a simple and efficient manner without outflow holes, i.e. locally closed. Furthermore, this allows the distribution of the coolant to the two cooling passages and the resulting pressure losses therein to be efficiently adjusted.

[0018] In another preferred embodiment, two cooling branch channels are provided that extend the second coolant flow passage, which branches radially inwards as they extend in the chord direction and connect to the third coolant flow passage. This measure reduces or compensates for the reduction in the flow cross section of the second coolant flow passage, which is caused by the drop shape of the blade profile tapering towards the trailing edge. This allows to obtain an almost constant cross section over the entire length of the second coolant flow passage, which allows the first coolant flow to flow through the second coolant flow passage at a constant speed. This allows to avoid flow separation while maintaining uniform cooling of this local area of ​​the blade tip and both side walls.

[0019] Furthermore, according to a further development of the above embodiment, a partition is arranged between the second cooling medium flow path and the serpentine cooling medium flow path, which partition connects the two side walls to each other and extends in the chord direction, in which case as it approaches the trailing edge it preferably forms a pointed stop wedge which cooperates with the inner faces of the two side walls to laterally separate the two cooling branch channels.

[0020] According to another particularly preferred embodiment of the invention, a spanwise extending aft separator rib is provided between the third coolant flow passage and the second trailing edge coolant flow passage, and if necessary, one or more holes may be provided in the aft separator rib to prevent localized dead water areas in the second trailing edge coolant flow passage.

[0021] According to an advantageous proposal of the invention, the trailing edge has a normalized height of 100%, starting at 0% at its root end and ending at 100% at the tip. The two trailing edge coolant channels are at least substantially separated from one another by a separating rib extending mainly in the chord direction, which is located at a height between 45% and 75% of the normalized height. In this way, in particular, a particularly efficient distribution of the total amount of available coolant can be achieved, so that, on the one hand, a uniform cooling of the airfoil is achieved and, on the other hand, a further reduction in the coolant consumption itself can be achieved. In order to be able to better mount the casting cores required for casting the turbine blade, which are left behind afterwards at the rear of the two trailing edge coolant channels, and to avoid core breakage, it is advantageous if these casting cores are directly connected to one another via a few supports. These supports leave a number of openings in the separating rib in the finished turbine blade, so that the two trailing edge coolant channels are not completely separated from one another, but the two trailing edge coolant channels are still substantially separated from one another.

[0022] In a further development of the invention, it is preferred if the serpentine coolant flow path has at least two channel sections extending in the blade span direction and at least two alternating redirections, in which the redirection located further downstream in the coolant flow is directly fluidically connected to the first trailing edge coolant flow path.

[0023] Particularly preferred and advantageous is a development of the above-mentioned embodiment, in which the two channel sections are formed in the cross section of the airfoil by the stopper and two side walls, each substantially C-shaped, by the suction side branch channel, the pressure side branch channel and the connecting channel connecting these two branch channels, which are arranged with respect to each other so as to almost completely surround the stopper. As a result, a turbine airfoil designed as a multi-wall structure can be provided. The multi-wall structure, on the one hand, makes it possible to manufacture an airfoil with a relatively small curvature at the leading edge even with a small amount of material used. Of course, this small curvature greatly promotes the occurrence of twisting in the first cooling medium flow passage. On the other hand, the multi-wall construction allows the cooling section to have a relatively small flow cross section. Thus, during operation, the second cooling medium flow flows through these channel sections or serpentine cooling medium flow passages at a sufficiently high speed and thus creates a sufficiently high heat transfer. This in particular reduces the amount of cooling medium required for efficient cooling of the central region between the leading edge region and the trailing edge region of the airfoil. This measure allows consumption to be reduced by a further 40 percent, thereby bringing the thermal efficiency of the turbine blades significantly closer to the theoretical maximum.

[0024] In this case, it has proven effective if the stop encompasses the cavity in cross section and is supported on the two side walls via bridges.

[0025] According to an advantageous development, in order to compensate the Coriolis forces acting on the second cooling medium during operation in the turbine blade, at least one, preferably both, of the support ribs connecting the pressure side wall and the suction side wall, which extend from the blade root side towards the blade tip, can be provided with a number of components, preferably turbulators, on the support rib or on the inner surface of the stopper separating the connecting channels, which can reduce the lateral flow of the cooling medium from the suction side branch channel through the connecting channel to the pressure side branch channel.

[0026] According to another preferred embodiment, the cavity does not have a cooling medium outlet opening, so that the cooling medium cannot flow through the cavity. This prevents undesired disturbances of the second cooling medium flow and allows the use of a particularly simple casting device, in which the casting core used can be attached to the other components of the casting device in a particularly simple and stable manner. The turbine blade according to the invention can therefore be advantageously cast, in which the openings present in the blade root after the turbine blade has been cast and which are directly connected to the cavity are closed by a separately manufactured cover plate. The same applies to the openings present in the blade root after the turbine blade has been cast and which are directly connected to the first trailing edge cooling medium channel. Preferably, such openings are closed by a separately manufactured cover plate which is fixed to the blade root in such a way that it completely covers the openings in question.

[0027] It is preferred that one or more inlets are provided for each cooling passage, which are directly fluidly connected to the first coolant flow path or to a supply flow path or to one of the serpentine coolant flow paths or to one of the channel sections thereof.

[0028] Preferably, the aspect ratio of the trailing edge span length to the trailing edge chord length measured at the root end is less than or equal to 3.0, since it has been found that the division of the available cooling medium into two preferably separated cooling medium streams and at the same time the proposed division of the cooling of the trailing edge region especially for this type of turbine blade allows for a considerable saving in the amount of cooling medium.

[0029] In essence, the turbine blades described above can be used both as moving blades attached to a rotor and as stationary vanes attached to a static support.

[0030] Surprisingly, the turbine blades described above can also be used in the first or second turbine stage of stationary gas turbines having a turbine inlet temperature of at least 1300° C. during normal operation according to ISO and / or a compression ratio of at least 19:1 during normal operation according to ISO. In the present application, so-called aircraft turbines do not fall within the definition of stationary gas turbines. The present invention is therefore not only suitable for stationary gas turbines, where the hot gas temperatures at the turbine inlet are considered relatively low by current standards.

[0031] The above description of advantageous configurations of the invention includes a number of features, which are partially grouped together in the individual dependent claims in the form of a unit. However, these features can also be considered individually and can be combined to form further combinations. In particular, these features can be combined with the method according to the invention and the device according to the invention, each individually and in any suitable combination. For example, features relating to the specifically described method can also be viewed as properties of the corresponding device unit and vice versa.

[0032] Although certain terms are used in the specification and claims in the singular or in combination with numerals, the scope of the invention is not limited to these terms in the singular or respective numerals. Furthermore, the words "ein" and "eine" are to be understood as indefinite articles and not as numerals. Similarly, the numerals "first", "second", "third", etc., serve only to distinguish features having substantially similar properties.

[0033] The above-mentioned characteristics, features and advantages of the present invention, as well as the manner in which they are achieved, will be explained in more detail below with reference to the following figures and with the description of the embodiments. [Brief description of the drawings]

[0034] [Figure 1] FIG. 1 is a side view of a turbine blade according to a first embodiment; [Diagram 2]Cooling scheme of turbine blades according to Fig. 1 [Diagram 3] FIG. 1 is a longitudinal sectional view of a turbine blade according to a first embodiment; [Figure 4] Cross-sectional view of the turbine blade taken along line AA in FIG. [Diagram 5] A longitudinal section of a turbine blade taken along line BB in FIG. [Figure 6] A longitudinal section of a turbine blade taken along line CC in FIG. [Figure 7] FIG. 3 shows a longitudinal section of a turbine blade taken along line DD. [Figure 8] Cross section of turbine blade taken along line EE in FIG. [Figure 9] Schematic diagram of a stationary gas turbine

[0035] In these drawings, all technical features which are provided with the same reference signs have the same technical effect. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0036] The invention will be explained below on the basis of a turbine blade 10 which is configured as a turbine rotor blade, nevertheless the invention can also be directed to turbine stator blades.

[0037] FIG. 1 shows a turbine blade 10 in a side view according to a first embodiment of the invention. The turbine blade 10, preferably manufactured by a precision casting process, comprises a blade root 12, only the tip of which is shown. The blade root 12 can be formed in a dovetail or Christmas tree shape in a known manner. This is followed by a platform 13 from which an airfoil 18 extends in the blade span direction R from a root end 20 to a blade tip 22. When the turbine blade 10 is installed in an axial gas turbine, the blade span direction and the radial direction of the gas turbine coincide. The airfoil 18 extends in a chord direction S transverse to the blade span direction R from a leading edge 24 to a trailing edge 26. At the trailing edge 26 a number of outlet holes 46, 56 are distributed along the blade span. The aspect ratio HSP / SL of the trailing edge blade span length HSP to the blade chord length SL measured at the root end is 1.9 in this embodiment and is preferably in the range of 1.5 to 3.

[0038] A plurality of outlet holes 28 similarly open into the side of the platform 13. These outlet holes 46, 56 and outlet holes 28 are in fluid communication with the internal cooling system of the turbine blade 10.

[0039] The cooling system of the turbine blade 10, in particular the airfoil 18, is shown diagrammatically as a cooling scheme in FIG. 2. The turbine blade 10 can be supplied with a first cooling medium flow M1 and a second cooling medium flow M2 separately. The first cooling medium flow M1 flows through a first cooling passage 30, which is made up of a number of cooling medium channels 31, 32, 33, 34, 36a, 36b, 38, 40, 44. Downstream of the inlet of the first cooling medium flow M1, not shown in FIG. 2, is the supply channel 31, which is fluidly connected via a number of through openings 33 to a first cooling medium channel 32. This first cooling medium channel 32 is used for cyclone cooling of the leading edge 24 of the airfoil 18 and the leading edge region 39 directly adjacent thereto. In the region of the blade tip 22, the first coolant passage 32 merges into a second coolant passage 34 which runs from the leading edge 24 over a relatively long chord length of the blade tip 22 in the direction of the trailing edge 26 in order to cool the blade tip 22. A number of third outlet holes 67 can be arranged in the blade tip for cooling the grazing edge, which will be described later. Furthermore, the second coolant passage 34 is divided into two cooling branch channels 36a, 36b starting from the rear half of the second coolant passage 34. b, which are connected to the third cooling medium passage 38, as well as to the downstream end of the second cooling medium passage 34. The latter is fluidically connected to the second trailing edge cooling medium passage 44 via a turning portion 40. The cooling medium flow M1 flowing through the first cooling passage 30 can then exit the trailing edge 26 of the turbine blade 10 via a number of second outlet holes 46. Arranged parallel to the first cooling passage 30, and preferably completely fluidically separated therefrom, is a second cooling passage 50, which has a serpentine cooling medium passage 52 downstream of an inlet, not shown in detail in FIG. 2. The serpentine cooling medium passage 52 comprises two channel sections 55a, 55b extending in the span direction in this embodiment for cooling the central region 48 (FIG. 1), which are connected to each other via a turning portion 57a ​​arranged between them. The downstream end of the second channel section 55b includes a second turn section 57b that fluidly connects the second channel section 55b with the first trailing edge coolant flow passage 54. The coolant flow M2 flowing through the second cooling passage 50 may then exit the trailing edge 26 of the turbine blade 10 via the plurality of first exit holes 46. Both trailing edge coolant flow passages 44, 54 serve to cool the trailing edge region 59 (FIG. 1).

[0040] FIG. 3 shows the internal structure of the turbine blade 10 according to FIG. 1 in a longitudinal section, which is arranged to correspond to the cooling scheme according to FIG. 2. For this purpose, the turbine blade 10 is provided with a series of variously arranged walls and ribs, which separate the individual cooling passages and the cooling medium channels from one another. Two inlets 80 for the two cooling medium flows M1 and M2, i.e. for the two cooling channels 30, 50, are provided in the blade root 12. Between the two inlets 80, a front support rib 66v is arranged, which connects the two side walls 14, 16 to one another and separates the first cooling channel 30 from the second cooling channel 50 for the first section. Furthermore, a front separating rib 49v separates the supply channel 31 from the first cooling medium channel 32, in which a number of through-openings 33 (detailed in FIG. 4) are arranged. However, in FIG. 3, only the connections of these through-openings are shown. As can be seen from FIG. 3, the density of through-openings 33 is greater in the region closer to the platform than in the region closer to the tip. The position and orientation of the through-openings 33 in the front separating rib 49v are selected in order to generate a more twisted coolant flow in the first coolant flow passage 32. By twisted coolant flow, we mean flows that develop in a cyclone-like or spiral-like manner, i.e. helical manner, from the root end 20 towards the tip 22. They are therefore arranged eccentrically in the front separating rib 49v, in particular in alignment with the inner wall of the suction side wall 16 (or the pressure side wall). If necessary, they may even be inclined towards the tip 22 in order to at least partially compensate for the weakening of the twist when flowing through the first coolant flow passage 32.

[0041] A second coolant passage 34 connects to the outer end of the first coolant passage 32 for cooling the bottom 37 of the blade tip 22 and is separated from the serpentine coolant passage 52 by a bulkhead 60. A third coolant passage 38 connects to the end of the second coolant passage 34 near the trailing edge and extends from the blade tip 22 to the root end. 203, but only extends to about half the height of the airfoil 18, the height of which is measured at the trailing edge 26. A further redirection 40 is connected thereto, by means of which the first coolant flow M1 can be fed into a second trailing edge coolant flow passage 44. This third coolant flow passage 38 is separated from the second trailing edge coolant flow passage 44 by a suitably configured aft separation rib 49h. 44 is largely separated from

[0042] In the second trailing edge coolant flow passage 44, a number of pedestals 53 are arranged one behind the other in a number of rows, through which the first coolant M1 can be bypassed. In the illustrated embodiment, the pedestals are formed in the shape of a racetrack with a relatively narrow passage in order to provide the lowest possible pressure loss. The first cooling passage 30 ends in a number of second outlet holes 46 at the trailing edge 26, through which at least a major part of the first coolant flow M1, which is fed through the associated inlet 80, can be discharged from the turbine rotor blade 10.

[0043] The second cooling passage 50 for guiding the second coolant flow M2 substantially comprises a serpentine coolant flow passage 52 and a first trailing edge coolant flow passage 44. The former can be divided into four successive sections, the first of which is called the first channel section 55a. This is followed by a first turn 57a, a second channel section 55b and a second turn 57b. The latter joins the serpentine coolant flow passage 52 with the second trailing edge coolant flow passage 54 and this first turn 57b. 1 The trailing edge coolant flow passage 54 is 2 Similar to the trailing edge coolant flow passage 44, the trailing edge coolant flow passage 44 is configured as multiple rows of pedestals 53 in a racetrack configuration.

[0044] The two channel sections 55a, 55b of the serpentine coolant flow passage 52 extend in the span direction R over most of the airfoil 18. Both the first channel section 55a and the second channel section 55b are substantially U-shaped, as additionally shown in FIG. 4, and each have branch channels 55as, 55bs arranged on the suction side and branch channels 55ad, 55bd arranged on the pressure side, as well as connecting channels 55av, 55bv connecting these branch channels. As a result, the first channel section 55a is surrounded by the pressure side wall 14, the front support rib 66v, the suction side wall 16, and an internally arranged stopper 70, which is shown in the cross section according to FIG. 4. The second channel section 55b is surrounded by the pressure side wall 14, the rear support rib 66h, the suction side wall 16, and an internally arranged stopper 70. The stopper 70 itself includes a cavity 72 and is supported on the pressure and suction side walls 14, 16 by a number of bridges 71 which extend over almost the entire height of the airfoil 18 and which, on the one hand, serve to monolithically fix the stopper 70 on the turbine blade 10 and, on the other hand, to separate the two channel sections 55, 57. With reference to FIG. 70 2, it can be seen that the radially outer end of the stopper 72 is truncated on the trailing edge side. This measure improves the mechanical integrity of the turbine blade 10 and in particular its vibration resistance. With reference to FIG. 2, it can be seen that the radially outer end of the stopper 72 is truncated on the trailing edge side. This measure improves the mechanical integrity of the turbine blade 10 and in particular its vibration resistance.

[0045] The two trailing edge coolant channels 44, 54 are at least largely, if not completely, separated from one another by a separating rib 64 that extends primarily in the chord direction S. According to this embodiment, the separating rib 64 terminates at a height that is 55% of the normalized airfoil height of the trailing edge 24. Preferably, the separating rib 64 is located at a height between 45% and 75% of the normalized height.

[0046] 5 to 7 show cross sections passing through the tip of the turbine rotor blade 10 along the three cross-sectional lines BB, CC and DD in FIG. 3. 78 indicates the blade tip. 22 1, the cavity 72 has no outlet opening, since the opening 74a necessary for the installation of the cavity 72 or the stopper 70 on the blade root 12 is closed by a cover plate 76a (FIG. 1) attached thereto after casting. As a result, no flow can be made through this cavity, which is formed as a dead space. It is therefore proposed to change its internal shape, if modal adjustment is required, already in the design phase by providing further structures such as ribs, braces, etc. The particular advantage is that only the natural frequencies of the turbine blades will be optimized, without affecting other characteristics such as aerodynamics or heat exchange.

[0047] 5 to 7 also show how the bulkhead 60 forms a pointed blocking wedge 62 as it approaches the trailing edge 24, which joins with the inner faces of the two side walls 14, 16 to laterally delimit the two cooling branch channels 36a and 36b, respectively. The truncation of the blocking body 70 can be compensated for with the help of the tapered blocking wedge 62, so that it is possible to continue to guide the cooling medium flow M2 near the side walls in the truncated area and thus to cool it sufficiently efficiently. If truncation of the blocking body is not necessary, the size of the blocking wedge can be reduced. In some cases, it can even be omitted completely.

[0048] FIG. 8 shows a cross section of the downstream half of the blade tip 22 through section line EE of FIG. 3, in a view looking towards the blade tip 22, ie, looking outward.

[0049] According to a second embodiment, not shown, instead of or in addition to the supply passage 31, a root channel section can be provided, which can be an extension of the first cooling medium passage 32 to the underside of the blade root part 12. Correspondingly, a number of suitable twist generators, for example helical ribs, can be provided in this root channel section, which twist the first cooling medium flow M1 in a cyclonic manner as it flows through the root channel section. In this case, the first cooling medium flow 32 would be separated from the connecting channel 55av by the front support rib 66v, so that the through-openings 33 arranged in the front support rib 66v could promote the renewal or intensification of the twist pulse. In this respect, it may be sensible not to guide the two cooling medium flows M1 and M2 completely separately from each other through the turbine blade 10, but to allow a very small extent of passage by connecting two separate cooling passages, which are otherwise fluidically separated, with very few individual holes, preferably with a small diameter.

[0050] FIG. 9 shows a gas turbine 100 with a compressor 110, a combustion chamber 120 and a turbine unit 130, which is only diagrammatically represented. According to this embodiment, a generator 150 for generating electricity is connected to the rotor 140 of the gas turbine. The compressor 110 is designed so that, during normal operation according to ISO, a pressure ratio of compressed ambient air VL to aspirated ambient air L is greater than or equal to 19:1. In the combustion chamber 120, the compressed air VL is then mixed with a fuel F and burned to generate hot gases HG. The combustion chamber 120 and the turbine unit 130 are designed so that the hot gases HG flowing at the outlet of the combustion chamber 120 or at the inlet of the turbine unit 130 have a temperature of at least 1300° C. during normal operation according to ISO, with the rotor blades and stator blades of the first or second turbine stage being configured as described above. The hot gases HG expanded in the turbine unit 130 are discharged as flue gas RG.

[0051] In summary, the invention proposes a turbine blade 10 with a blade root 12 and an airfoil 18, which extends in a blade span direction R from a root end 20 to a blade tip 22 and in a blade chord direction S transverse to the blade span direction R from a leading edge 24 to a trailing edge 26, inside of which a first cooling passage 30 for a first cooling medium flow M1 and a second cooling passage 50 for a second cooling medium flow M2 are formed, the first cooling passage 30 comprising a first cooling medium flow passage 32 provided for cyclone cooling of the leading edge 24, , a second cooling medium passage 34 extending from the leading edge 24 toward the trailing edge 26 beneath the blade tip 22 and continuing from the first cooling medium passage 32, the second cooling passage 50 including a serpentine coolant passage 52 for cooling a central region 48 chordwise aft of the leading edge region 39 of the airfoil 18, and a first trailing edge coolant passage 54 for at least partially cooling a trailing edge region 59 chordwise aft of the central region 48 of the airfoil 18 to the trailing edge, the first trailing edge coolant passage 54 being fluidly connected to a plurality of outlet holes 56 disposed at the trailing edge 26. In order to provide a turbine blade with an even lower cooling medium consumption, it is proposed to provide a first cooling medium passage 32 and / or a serpentine cooling medium passage 52 for localized closed cooling, the first cooling passage 30 including a third cooling medium passage 38 connected to the second cooling medium passage 34 and extending mainly radially inward, as well as a second trailing edge cooling medium passage 44 connected to the third cooling medium passage 38, the second trailing edge cooling medium passage 44 being designed to cool the blade tip region of the trailing edge region 59 and being fluidly connected to a plurality of second outflow holes 46 arranged at the trailing edge 26.

Claims

1. A turbine blade (10) for a gas turbine, the turbine blade comprising a blade root (12) and an airfoil portion (18) including a pressure sidewall (14) and a suction sidewall (16), the sidewalls (14, 16) (hereinafter referred to as "the two sidewalls (14, 16)") extending from a root end (20) to a tip end (22) along a blade span direction (R) and extending from a leading edge (24) to a trailing edge (26) along a chord direction (S) oriented transversely to the blade span direction (R); a first cooling passage (30) for a first cooling medium flow (M1) and a second cooling passage (50) for a second cooling medium flow (M2) substantially separated from the first cooling passage (30) for the first cooling medium flow (M1), the first cooling passage (30) includes a first cooling medium flow path (32) provided for cyclone cooling of the leading edge (24), and a second cooling medium flow path (34) connected to the first cooling medium flow path (32) and extending in a direction from the leading edge (24) to the trailing edge (26) below the blade tip (22); the second cooling passage (50) includes a serpentine coolant flow passage (52) for cooling a central region (48) chordwise aft of a leading edge region (39) of the airfoil (18), and a first trailing edge coolant flow passage (54) for at least partially cooling a trailing edge region (59) chordwise aft of the central region (48) of the airfoil (18) to the trailing edge, the first trailing edge coolant flow passage (54) is fluidly connected to a plurality of first outlet holes (46) disposed in the trailing edge (26); In turbine blades, the first coolant flow passage (32) and / or the serpentine coolant flow passage (52) are devoid of outlet holes; the first cooling passage (30) includes a third cooling medium flow passage (38) connected to the second cooling medium flow passage (34) and extending mainly radially inward, and a second trailing edge cooling medium flow passage (44) connected to the third cooling medium flow passage (38), the second trailing edge cooling medium flow passage (44) being formed for cooling a blade tip side region of the trailing edge region (59) and being fluidly connected to a plurality of second outflow holes (46) arranged in the trailing edge (26); The first cooling passage (30) includes a supply passage (31) for the first cooling medium flow passage (32), the supply passage comprising: located immediately adjacent to the first coolant flow passage (32); extending over at least a majority of the span length of said airfoil (18); a plurality of through openings (33) fluidly connected to the first coolant flow passage (32), the plurality of through openings (33) having means for imparting a twist to a coolant flow (M1) flowing in the first coolant flow passage (32); The density of the plurality of through openings (33) that can be defined in the blade span direction (R) is greatest at the root end (20); A turbine blade characterized by:

2. 2. The turbine blade (10) of claim 1, characterized in that one or more cooling medium outlet holes (67) are arranged in the blade tip (22) and are fluidly connected to the second cooling medium flow passage (34).

3. 3. The turbine blade (10) according to claim 1 or 2, characterized in that the density of the plurality of through openings (33) that can be defined in the blade span direction (R) decreases stepwise or continuously towards the blade tip (22).

4. 4. The turbine vane (10) of claim 1, wherein each of the trailing edge coolant flow passages (44, 54) includes a number of pedestals arranged in a patterned array of rows.

5. 5. The turbine blade (10) according to claim 1, characterized in that two cooling branch channels (36 a, 36 b) are provided extending the second cooling medium flow passage (34), the two cooling branch channels (36 a, 36 b) widening radially inward as they extend in the chord direction and connecting to the third cooling medium flow passage (38).

6. 6. The turbine blade (10) according to claim 5, characterized in that a partition (60) is arranged between the second cooling medium flow passage (34) and the serpentine cooling medium flow passage (52), the partition wall (60) connecting the two side walls (14, 16) to each other and extending in the chord direction (S), the partition wall (60) forming a pointed stop wedge (62) as it approaches the trailing edge (26), the stop wedge connecting with inner surfaces of the two side walls (14, 16) and laterally dividing the two cooling branch channels (36 a, 36 b).

7. 7. The turbine blade (10) according to claim 1, further comprising a rear separation rib (49h) extending in a blade span direction (S) between the third coolant flow passage (38) and the second trailing edge coolant flow passage (44).

8. 8. The turbine blade of claim 1, wherein the trailing edge (26) has a normalized height of 100% starting at 0% at its root end (20) and ending at 100% at the tip (22), and the first and second trailing edge coolant channels (44, 54) are at least substantially separated from one another by a primarily chordwise extending separation rib (64), the separation rib being located at a height between 45% and 75% of the normalized height.

9. A turbine blade (10) for a gas turbine, the turbine blade comprising a blade root (12) and an airfoil portion (18) including a pressure sidewall (14) and a suction sidewall (16), the sidewalls (14, 16) (hereinafter referred to as "the two sidewalls (14, 16)") extending from a root end (20) to a tip end (22) along a blade span direction (R) and extending from a leading edge (24) to a trailing edge (26) along a chord direction (S) oriented transversely to the blade span direction (R); a first cooling passage (30) for a first cooling medium flow (M1) and a second cooling passage (50) for a second cooling medium flow (M2) substantially separated from the first cooling passage (30) for the first cooling medium flow (M1), the first cooling passage (30) includes a first cooling medium flow path (32) provided for cyclone cooling of the leading edge (24), and a second cooling medium flow path (34) connected to the first cooling medium flow path (32) and extending in a direction from the leading edge (24) to the trailing edge (26) below the blade tip (22); the second cooling passage (50) includes a serpentine coolant flow passage (52) for cooling a central region (48) chordwise aft of a leading edge region (39) of the airfoil (18), and a first trailing edge coolant flow passage (54) for at least partially cooling a trailing edge region (59) chordwise aft of the central region (48) of the airfoil (18) to the trailing edge, the first trailing edge coolant flow passage (54) is fluidly connected to a plurality of first outlet holes (56) disposed in the trailing edge (26); In turbine blades, the first coolant flow passage (32) and / or the serpentine coolant flow passage (52) are devoid of outlet holes; the first cooling passage (30) includes a third cooling medium flow passage (38) connected to the second cooling medium flow passage (34) and extending mainly radially inward, and a second trailing edge cooling medium flow passage (44) connected to the third cooling medium flow passage (38), the second trailing edge cooling medium flow passage (44) being formed for cooling a blade tip side region of the trailing edge region (59) and being fluidly connected to a plurality of second outflow holes (46) arranged in the trailing edge (26); the serpentine coolant flow passage (52) includes at least two spanwise channel sections (55 a, 55 b) and at least two turn sections (57 a, 57 b), the turn section (57 b) being located further downstream in the coolant flow and directly fluidly connected to the first trailing edge coolant flow passage (54); the two channel sections (55a, 55b) are formed by a stopper (70) and by the two side walls (14, 16), each of which is substantially C-shaped in the cross section of the airfoil (18), and which are formed by a suction side branch channel (55as, 55bs), a pressure side branch channel (55ad, 55bd), and a connecting channel (55av, 55bv) connecting the suction side branch channel (55as, 55bs) and the pressure side branch channel (55ad, 55bd), and these channel sections are arranged relative to each other so as to almost completely surround the stopper (70); A turbine blade characterized by:

10. 10. A turbine blade according to claim 9, characterized in that the stop (70) includes a cavity (72) in cross section and is supported on the two side walls (14, 16) via bridges (71).

11. 11. The turbine blade (10) according to claim 9 or 10, characterized in that the serpentine coolant flow passage (52) is bounded by at least one, if not both, support ribs (66h, 66v) extending from the root end to the tip and connecting the pressure side wall (14) with the suction side wall (16), and a plurality of components are provided on the support ribs (66h, 66v) or on an inner surface of the stop (70) bounding the connecting channel (55av, 55bv), the plurality of components reducing lateral flow of coolant from the suction side branch channel (55as, 55bs) through the connecting channel (55av, 55bv) to the pressure side branch channel (55ad, 55bd).

12. A turbine blade (10) according to claim 10 or claim 11 which relies on claim 10, characterized in that the cavity (72) cannot allow the cooling medium flow (M) to flow and has no outlet opening for the cooling medium flow (M) (dead water cavity).

13. 13. A turbine blade (10) according to claim 10, claim 11 which depends on claim 10, or claim 12, characterized in that the turbine blade is cast and that an opening (74a) which is present in the blade root (12) after the turbine blade has been cast and which directly communicates with the cavity (72) is closed by a separately manufactured cover plate (76a).

14. A turbine blade (10) according to any one of the preceding claims, characterized in that it is cast.

15. 15. The turbine blade (10) according to claim 13 or 14, characterized in that an opening (74b) present in the blade root (12) after the turbine blade has been cast and which directly communicates with the first trailing edge coolant flow passage (54) is closed by a separately manufactured cover plate (76b).

16. 16. The turbine blade (10) according to any one of claims 1 to 8 or claim 14 or 15 directly or indirectly relying on claim 1, characterized in that for each cooling passage (30, 50) one or more inlets (80) are provided, which are fluidically connected directly to the first coolant flow channel (32) or to the supply flow channel (31) or to one of the serpentine coolant flow channels (52) or to one of its channel sections (55 a).

17. 17. The turbine blade (10) of claim 1, wherein an aspect ratio HSP / SL of a trailing edge span length (HSP) to a chord length (SL) measured at the root end is less than or equal to 3.

0.

18. 18. Use of a turbine blade (10) according to any one of claims 1 to 17 in the first or second turbine stage of a stationary gas turbine having a turbine inlet temperature of at least 1300°C during normal operation according to ISO and / or a compression ratio of 19:1 or greater during normal operation according to ISO.

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