EP3832069A1 - Aube de turbine pour turbine à gaz fixe - Google Patents

Aube de turbine pour turbine à gaz fixe Download PDF

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Publication number
EP3832069A1
EP3832069A1 EP19214178.6A EP19214178A EP3832069A1 EP 3832069 A1 EP3832069 A1 EP 3832069A1 EP 19214178 A EP19214178 A EP 19214178A EP 3832069 A1 EP3832069 A1 EP 3832069A1
Authority
EP
European Patent Office
Prior art keywords
coolant passage
coolant
blade
trailing edge
turbine blade
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP19214178.6A
Other languages
German (de)
English (en)
Inventor
Philipp CAVADINI
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Siemens AG
Original Assignee
Siemens AG
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Siemens AG filed Critical Siemens AG
Priority to EP19214178.6A priority Critical patent/EP3832069A1/fr
Priority to JP2022532872A priority patent/JP2023505451A/ja
Priority to US17/780,670 priority patent/US12006838B2/en
Priority to CN202080084589.1A priority patent/CN114787482B/zh
Priority to EP20824139.8A priority patent/EP4048872B1/fr
Priority to KR1020227022611A priority patent/KR20220103799A/ko
Priority to PCT/EP2020/084603 priority patent/WO2021110899A1/fr
Publication of EP3832069A1 publication Critical patent/EP3832069A1/fr
Withdrawn legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/204—Heat transfer, e.g. cooling by the use of microcircuits
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes

Definitions

  • the invention relates to a turbine blade according to the preamble of claim 1.
  • Turbine blades of gas turbines are subject to the highest thermal and mechanical loads during operation, which is why they can now be cooled with the help of complex, hollow internal geometries and are designed to be particularly robust.
  • a gas turbine blade corresponding to the preamble of claim 1 is from FIG WO 1996/15358 A1 known, with the aid of cooling air introduced tangentially into a leading edge cooling channel, cooling of the leading edge is made possible without further film cooling holes, often also referred to as showerhead holes, being required for their cooling.
  • film cooling holes also known as gill holes, located in the suction side near the leading edge, whereas the remaining proportion of this cooling air is conducted below the blade tip to the trailing edge .
  • the remaining part of the airfoil is cooled via a serpentine cooling channel with subsequent rear edge blow-out.
  • multi-layer turbine blade which is also referred to in English as a "multi-wall turbine blade”.
  • two displacement bodies are provided with which the cooling air flowing inside the turbine blade is intended to be forced particularly close to the inner surfaces of the outer walls.
  • FIG EP 1 783 327 A2 A The alternative configuration of a multiwall turbine blade is also shown in FIG EP 1 783 327 A2 .
  • the object of the invention is consequently to provide a long-lasting turbine blade with a further reduced coolant consumption.
  • the present invention proposes a turbine blade for a stationary gas turbine, in particular for one of its high-pressure turbine stages, with a cooling system arranged in its interior, which has a first cooling path for a comprises first coolant flow and a second cooling path for a second coolant flow, in which the first cooling path comprises a first coolant passage, which is set up for cyclone cooling of the leading edge and a second coolant passage adjoining the first coolant passage, which extends below the blade tip from the leading edge in the direction the trailing edge, the second cooling path having a serpentine coolant passage for cooling a central area of the airfoil arranged in the chord direction behind the leading edge area and a first trailing edge coolant passage for at least partial cooling ng comprises a trailing edge area of the airfoil arranged in the chord direction behind the central area and extending to the trailing edge, the first trailing edge coolant passage having a plurality of first trailing edge Outlet holes is fluidically connected,
  • the invention is based on the knowledge that a significant saving in coolant for cooling the turbine blade can only be achieved if the leading edge and / or the pressure-side side wall and the suction-side side wall of the blade, ie the serpentine coolant passage, is set up for closed cooling .
  • This is understood here to mean that in particular in the leading edge and / or the flat areas of the blade - apart from the blade tip or its rubbing edges - there are no openings through which coolant can flow out and flow into a hot gas flowing around the turbine blade.
  • the first coolant flow is guided via a second coolant passage, which extends directly below the blade tip to the rear end of the airfoil, and via an adjoining third coolant passage at preferably approximately half the height of the trailing edge, and then in a radially outwardly arranged trailing edge coolant passage to be put to good use there. Because of this solution, the requirement for cooling air for the second flow path can be significantly reduced.
  • the approach proposed here therefore offers the maximum benefit of the available coolant due to a novel division and using a cooling concept, namely cyclone cooling, which was previously used for turbine blades of the first and / or second turbine stage of gas turbines with comparatively high compressor pressure ratios or high turbine inlet temperatures considered completely unsuitable and therefore not considered for their turbine blades.
  • the consumption of coolant can be reduced to an extent that cannot be expected in advance while at the same time providing adequate cooling of the entire airfoil. According to detailed simulations, this even applies to turbine blades in one of the two front turbine stages of a stationary gas turbine with a turbine inlet temperature of 1300 ° C and higher at ISO nominal operation or with a compressor pressure ratio of 19: 1 or higher. Even with such turbine blades, the The amount of coolant can be reduced by about 30% compared to a conventional one with cooling holes arranged in the front edge, while achieving the same service life.
  • one or more outlet holes for coolant which are fluidically connected to the second coolant passage, are arranged in the blade tip. This measure improves the fatigue strength of any rubbing edges protruding from the blade tip.
  • the first cooling path comprises a supply passage for the first coolant passage, which is arranged immediately next to the first coolant passage and extends at least over a large part of the span of the airfoil via a plurality of through openings with the first coolant passage in such a way that it is fluidically connected to the first coolant passage coolant flowing through the passage openings can impart or reinforce a swirl to the coolant flowing in the first coolant passage.
  • This can be achieved in that the passage openings open tangentially, i.e. eccentrically in the first coolant passage and in particular in alignment with the inner surface of the suction-side or pressure-side side wall and / or are positioned in the radial direction. Efficient cyclone cooling of the leading edge can thus be provided comparatively easily.
  • cyclone cooling of the leading edge that is adapted or homogenized over the height of the blade can be achieved in that a density of passage openings that can be determined in the spanwise direction is greatest at the base end, and preferably decreases gradually or continuously towards the blade tip.
  • the flow velocity in the first coolant passage can be kept almost constant over the span of the blade, which can also be achieved through a first coolant passage tapering in cross section to the blade tip.
  • a plurality of preferably rib-shaped, in particular inclined, turbulators are arranged on one or more inner surfaces of one or more coolant passages in order to further increase the heat transfer into the first and / or second coolant locally and / or to support the swirl .
  • a plurality of sockets arranged in a pattern, i.e. in several rows, are provided in each trailing edge coolant passage. This allows a trailing edge area of the airfoil, which adjoins the central area of the airfoil and extends to the rear edge of the airfoil, to be cooled in a closed manner in a simple and efficient manner.
  • the distribution of the coolant for the two cooling paths and the pressure losses occurring therein can also be efficiently adjusted.
  • two cooling channel arms which widen the second coolant passage are provided, which widen radially inward with increasing extension in the direction of the chord and open into the third coolant passage.
  • This measure reduces or compensates for the reduction in the flow cross-section of the second coolant passage, which results from the teardrop-shaped shape of the blade profile which tapers to a point towards the rear edge.
  • an approximately constant cross-sectional area can be achieved for the entire length of the second coolant passage, whereby the first coolant flow can flow through the second coolant passage at a constant speed. Flow separation can thus be avoided while maintaining uniform cooling of the blade tip and the local areas of the side walls.
  • a partition is arranged between the second coolant passage and the serpentine coolant passage, which connects the two side walls to one another and extends in the direction of the chord, the partition forming a preferably tapering displacement wedge as it approaches the rear edge in connection with the inner surfaces of the two side walls, the two cooling duct arms are laterally bounded.
  • a rear separating rib extending in the spanwise direction is provided between the third coolant passage and the second rear edge coolant passage. If necessary, one or more holes can also be present in the rear separating rib in order to prevent local dead water areas in the second rear edge coolant passage.
  • the trailing edge has a normalized height of 100%, starting at its root end at 0% and ending at the blade tip at 100%, the two trailing edge coolant passages at least from a separating rib that extends mainly in the direction of the chord are essentially separated from one another, which is arranged at a height between 45% and 75% of the normalized height.
  • this allows a particularly efficient division of the total available Achieve the amount of coolant with which, on the one hand, homogeneous cooling of the airfoil and, on the other hand, a further reduced coolant consumption per se can be achieved.
  • the serpentine coolant passage comprises at least two channel sections extending in the spanwise direction and at least two reversal sections which alternate with one another, the reversal section further downstream in the coolant flow being directly fluidically connected to the first trailing edge coolant passage is.
  • the two channel sections by means of a displacement body and by means of the two side walls in a cross-sectional view of the blade are each essentially C-shaped with a suction-side channel arm, a pressure-side channel arm and one connecting the two channel arms Connecting arms are designed and arranged to one another in such a way that they almost completely surround the displacement body.
  • a turbine blade configured as a multiwall can be provided.
  • it is designed as a multiwall possible to produce an airfoil that has a relatively small curvature at the leading edge even with low consumption of resources. This slight curvature is of course very beneficial to the generation of swirl in the first coolant passage.
  • the cooling sections can have comparatively small flow cross-sections.
  • the second coolant stream then flows through the channel sections or through the serpentine coolant passage at a sufficiently high speed and thus with the formation of a sufficiently high heat transfer.
  • This in particular reduces the amount of coolant required for efficient cooling of the central area of the airfoil between the leading edge and trailing edge area.
  • the consumption can be reduced by about a further 40%, whereby the thermal efficiency of the turbine blade can then be brought comparatively close to the theoretical maximum.
  • the displacement body viewed in cross section, encompasses a cavity and is supported on the two side walls via webs.
  • At least one, preferably both, support ribs connecting the pressure side wall to the suction side wall, which extend from the base end to the blade tip, can be used on the support rib or on the turbine rotor blade to compensate for Coriolis forces occurring on the second coolant during operation the connecting arms delimiting inner surfaces of the displacement body elements, preferably turbulators, may be provided. This can cause a cross flow of coolant can be reduced from the suction-side channel arm through the connecting arm to the pressure-side channel arm.
  • the turbine blade according to the invention is preferably cast, with an opening which is present in the blade root after the casting of the turbine blade and which is in direct, i.e. immediate connection with the cavity, being closed by a separately produced cover plate.
  • Such is preferably also closed in that a separately produced cover plate is fastened to the blade root so that it completely covers the relevant opening.
  • one or more inlets are provided for each cooling path, which are directly fluidically connected to the first coolant passage or the supply passage or to the serpentine coolant passage or one of its channel sections.
  • the turbine blade preferably has an aspect ratio of a trailing edge span based on a chord length to be detected at the root end, which is 3.0 or less, since it has been found that the proposed distribution of the available coolant in two coolant flows, which are preferably separate from one another, and the simultaneously proposed division of the cooling of the trailing edge region, in particular for turbine blades of this type, enables a considerable saving in the amount of coolant.
  • the turbine blade described above can be used both as a rotor blade attached to a rotor or as a guide blade attached to a static carrier.
  • the turbine blade described above can also be used in a first or second turbine stage of a stationary gas turbine that has a turbine inlet temperature of at least 1300 ° C. in ISO nominal operation and / or a compression ratio of 19: 1 or greater in ISO nominal operation .
  • a turbine inlet temperature of at least 1300 ° C. in ISO nominal operation and / or a compression ratio of 19: 1 or greater in ISO nominal operation .
  • aero derivatives do not fall under the definition of stationary gas turbines.
  • the invention is therefore not only suitable for those stationary gas turbines whose hot gas temperatures at the turbine inlet are considered to be comparatively low by today's standards.
  • FIG. 1 a turbine blade 10 in a side view.
  • the turbine blade 10 which is preferably produced in an investment casting process, comprises a blade root 12, which is only shown in the approach.
  • the blade root 12 can be designed in a known manner in a dovetail shape or a Christmas tree shape. This is followed by a platform 13, from which a blade 18 extends in the spanwise direction R from a foot-side end 20 to a blade tip 22. If the turbine rotor blade 10 is installed in a gas turbine through which there is an axial flow, the span direction and the radial direction of the gas turbine coincide. In a chord direction S oriented transversely to the span direction R, the blade 18 extends from a leading edge 24 to a trailing edge 26.
  • exit holes 46, 56 are distributed along the span direction.
  • an aspect ratio HSP / SL of a trailing edge span HSP based on a chord length SL to be detected at the foot end is 1.9 and is preferably in the range between 1.5 and 3.
  • Outlet openings 28 likewise open out on a lateral surface of the platform 13.
  • the outlet holes 46, 56 and the outlet openings 28 are in flow connection with an inner cooling system of the turbine rotor blade 10.
  • FIG Figure 2 The cooling system of the turbine blade 10 and in particular of the airfoil 18 is shown in FIG Figure 2 shown schematically as cooling schemes.
  • a first coolant flow M1 and a second coolant flow M2 can be fed separately to the turbine rotor blade 10.
  • the first coolant flow M1 flows through a first cooling path 30, which is composed of a plurality of coolant passages 31, 32, 33, 34, 36a, 36b, 38, 40, 44. Downstream one in the Figure 2
  • An inlet (not shown) for the coolant flow M1 is followed by a supply passage 31 which is in flow connection with a first coolant passage 32 via a multiplicity of passage openings 33.
  • the first coolant passage 32 is used for cyclone cooling of the leading edge 24 of the airfoil 18 and the immediately adjoining leading edge area 39.
  • the first coolant passage 32 merges into a second coolant passage 34, which extends from the leading edge 24 to cool the blade tip 22
  • the blade tip 22 extends over a comparatively large chord length in the direction of the rear edge 26.
  • Third outlet holes 67 for cooling rubbing edges explained later can be arranged in the blade tip.
  • the second coolant passage 34 further comprises two cooling channel arms 36a, 36 which only begin in the second half of the second coolant passage 34 and which, like the downstream end of the second coolant passage 34, are connected to a third coolant passage 38.
  • the latter is via a reversal section 40 fluidically connected to a second trailing edge coolant passage 44.
  • the coolant flow M1 flowing through the first cooling path 30 can then leave the turbine rotor blade 10 at its trailing edge 26 via a multiplicity of second outlet holes 46.
  • the second cooling path 50 is arranged, which is downstream of an in Figure 2 has a serpentine coolant passage 52 inlet, not shown further.
  • the serpentine coolant passage 52 includes a central region 48 ( Figure 1 ) According to this exemplary embodiment, two channel sections 55a, 55b which extend in the spanwise direction and which are connected to one another via a reversing section 57a arranged between them.
  • first trailing edge coolant passage 54 At the downstream end of the second channel section 55b, there is a second reversal section 57b, which fluidically connects the second channel section 55b to a first trailing edge coolant passage 54.
  • the coolant flow M2 flowing through the second cooling path 50 can then leave the turbine rotor blade 10 at its rear edge 26 via a multiplicity of first outlet holes 46.
  • Both trailing edge coolant passages 44, 54 serve to cool a trailing edge area 59 ( Figure 1 ).
  • FIG. 3 shows, as a longitudinal section, an inner structure of the turbine rotor blade 10 according to FIG Figure 1 , which according to the cooling scheme Figure 2 is designed accordingly.
  • the turbine rotor blade 10 comprises a number of differently arranged walls and ribs which separate the individual cooling paths and coolant passages from one another.
  • Two inlets 80 for the two coolant flows M1 and M2 or for the two cooling paths 30, 50 are provided in the blade root 12. Between the two inlets 80 there is a front support rib 66v which connects the two side walls 14, 16 to one another and which, for a first section, forms the first Separates cooling path 30 from second cooling path 50.
  • a front separating rib 49v also separates the supply passage 31 from the first coolant passage 32, a plurality of passage openings 33 (detail to FIG. 4) being arranged in the front separating rib 49v. In Figure 3 of these, however, only the mouths of the passage openings are shown. How out Figure 3 As can be seen, a greater density of passage openings 33 is provided in the area close to the platform than in the area close to the tip. The position and the orientation of the passage openings 33 in the front separating rib 49v is selected such that a comparatively strongly twisted coolant flow can arise in the first coolant passage 32.
  • a swirled coolant flow is to be understood as one which can form cyclone-like or analogous to a helical line or a helix from the foot-side end 20 to the blade tip 22. They are therefore arranged eccentrically in the front separating rib 49v and in particular aligned with the inner walls of the suction side wall 16 (or pressure side wall), possibly even at an incline towards the blade tip 22 in order to at least partially compensate for the weakening of the swirl when flowing through the first coolant passage 32.
  • the second coolant passage 34 connects to cool a bottom 37 of the blade tip 22, the second coolant passage 34 being separated from the serpentine coolant passage 52 by a partition 60.
  • the third coolant passage 38 connects, which extends from the blade tip 22 in the direction of the root end 22, but only up to approximately half the height of the blade 18, the height of the blade 18 at the trailing edge 26 is to be recorded. This is followed by a further reversal section 40, by means of which the first coolant flow M1 can be supplied to the second trailing edge coolant passage 44.
  • the third coolant passage 38 is largely separated from the second rear edge coolant passage 54 by a correspondingly configured rear partition rib 49h.
  • the bases are designed more like a race track with comparatively narrow passages in order to bring about the highest possible pressure loss.
  • the first cooling path 30 ends in second outlet holes 46 provided in the rear edge 26, through which at least a large part of the coolant flow M1 supplied through the associated inlet 80 can be released from the turbine rotor blade 10.
  • the second cooling path 50 for guiding the second coolant flow M2 and essentially comprises the serpentine coolant passage 52 and the first trailing edge coolant passage 44.
  • the former can be divided into four successive sections, the first of which is referred to as the first channel section 55a. This is followed by a first reversal section 57a, a second channel section 55b and a second reversal section 57b.
  • the latter connects the serpentine coolant passage 52 with the second trailing edge coolant passage 54, which is designed analogously to the first trailing edge coolant passage 44 with racetrack-shaped sockets 53 arranged in several rows.
  • the two channel sections 55a, 55b of the serpentine coolant passage 52 extend along the span direction R over a large part of the airfoil 18 first channel section 55a and second channel section 55b are, as in FIG Figure 4 additionally shown, essentially U-shaped, each with a channel arm 55as, 55bs arranged on the suction side, a channel arm 55ad, 55bd arranged on the pressure side and a connecting arm 55av, 55bv connecting the respective channel arms.
  • the first channel section 55a of the pressure-side side wall 14, of the front support rib 66v, of the suction-side side wall 16 and a displacement body 70 arranged in the interior is shown in cross section according to Figure 4 - surround.
  • the second channel section 55b is surrounded by the pressure-side side wall 14, by a rear support rib 66h, by the suction-side side wall 16 and the displacement body 70 arranged in the interior.
  • the displacement body 70 itself engages around a cavity 72 and is supported via webs 71 on the pressure-side side wall 14 or the suction-side side wall 16.
  • the webs 71 extend approximately over the entire height of the blade 18 and serve on the one hand for monolithic fastening of the displacement body 70 in the turbine rotor blade 10 and on the other hand to separate the two channel sections 55, 57 Figure 2 it can be seen that the displacement body 72 is trimmed at its radially outer end on the rear edge side. This measure improves the mechanical integrity of the turbine rotor blade 10 and, in particular, its vibration resistance.
  • the two trailing edge coolant passages 44, 54 are separated from one another by a separating rib 64 that extends mainly in the chordal direction S, at least for the most part, if not completely.
  • the separating rib 64 ends at a height of 55% of a standardized blade height of the trailing edge 24.
  • the separating rib 64 is preferably arranged at a height between 45% and 75% of the standardized height.
  • FIGS 5 to 7 show sections through the tip of the turbine rotor blade 10 according to the three section lines BB, CC and DD Figure 3 .
  • rubbing edges 78 are provided both on the suction side and on the pressure side.
  • the displacement body 70 is not closed at its radially outer end, but rather is open towards the first reversing section 57a. In this respect, it would be possible for the second coolant flow M2 to flow in.
  • an opening 74a on the blade root 12, which is required for creating the cavity 72 or the displacement body 70 is replaced by a cover plate 76a attached there after casting ( Figure 1 ) is closed, the cavity 72 lacks outlet openings.
  • Figure 8 shows in a view directed towards the blade tip 22 - that is to say outward - a cross section of the downstream half of the blade tip 22 according to section line EE Figure 3 .
  • a channel section on the blade root side can be provided, which can represent an extension of the first coolant passage 32 to the underside of the blade root 12.
  • suitable swirl generators for example spiral ribs, can be provided which twist the coolant flow M1 in a cyclonic manner when it flows through the channel section on the blade root side.
  • the first coolant passage 32 would be separated from the connecting channel 55av by the front support rib 66v, so that passage opening 33 arranged in the front support rib 66v could promote a refreshment or amplification of the swirl pulse.
  • it can possibly even make sense not to completely separate the two coolant flows M1 and M2 from one another, but rather to support them to a small extent.
  • Figure 9 shows a gas turbine 100 with a compressor 110, a combustion chamber 120 and a turbine unit 130 only schematically.
  • a generator 150 for generating electricity is coupled to a rotor 140 of the gas turbine.
  • the compressor 110 is designed in such a way that, during operation under ISO standard conditions, it has a pressure ratio of compressed ambient air VL to sucked-in ambient air L of 19: 1 or greater.
  • the compressed air VL is then mixed with a fuel F and burned to form a hot gas HG.
  • Combustion chamber 120 and turbine unit 130 are designed in such a way that the hot gas HG flowing at the exit of the combustion chamber 120 or at the inlet of the turbine unit 130 has a temperature of at least 1300 ° C. under ISO standard conditions, the rotor blades and guide vanes of the first turbine stage or the second turbine stage are designed in the manner described here.
  • the hot gas HG expanded in the turbine unit 130 leaves it as flue gas RG.
  • the invention proposes a turbine blade 10 with a blade root 12 and a blade 18, which extends along a span direction R from a root end 20 to a blade tip 22 and along a chord direction S arranged transversely to the span direction R from a leading edge 24 to a trailing edge 26, wherein a first cooling path 30 for a first coolant flow M1 and a second cooling path 50 for a second coolant flow M2 are configured in the interior of the airfoil 18, the first cooling path 30 being a first coolant passage 32 which is configured for cyclone cooling of the leading edge 24 and a second coolant passage 34 adjoining the first coolant passage 32 and extending below the blade tip 22 from the leading edge 24 towards the trailing edge 26, wherein the second cooling path 50 comprises a serpentine coolant passage 52 for cooling a chordwise trailing edge edge area 39 of the airfoil 18 and a first trailing edge coolant passage 54 for at least partial cooling of a central area arranged in the direction of the chord behind the central area 48 and reaching to the trailing edge
  • the first coolant passage 32 and / or the serpentine coolant passage 52 is set up for closed cooling and the first cooling path 30 is a third coolant passage 38 that adjoins the second coolant passage 34 and that is extends mainly radially inward and comprises a second trailing edge coolant passage 44 adjoining the third coolant passage 38, which is designed for cooling a region of the trailing edge region 59 on the blade tip side and is fluidically connected to a plurality of second outlet holes 46 arranged in the trailing edge 26 .

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP19214178.6A 2019-12-06 2019-12-06 Aube de turbine pour turbine à gaz fixe Withdrawn EP3832069A1 (fr)

Priority Applications (7)

Application Number Priority Date Filing Date Title
EP19214178.6A EP3832069A1 (fr) 2019-12-06 2019-12-06 Aube de turbine pour turbine à gaz fixe
JP2022532872A JP2023505451A (ja) 2019-12-06 2020-12-04 定置形ガスタービンのタービン翼
US17/780,670 US12006838B2 (en) 2019-12-06 2020-12-04 Turbine blade for a stationary gas turbine
CN202080084589.1A CN114787482B (zh) 2019-12-06 2020-12-04 用于固定式的燃气轮机的涡轮叶片
EP20824139.8A EP4048872B1 (fr) 2019-12-06 2020-12-04 Aube de turbine pour turbine à gaz fixe
KR1020227022611A KR20220103799A (ko) 2019-12-06 2020-12-04 고정식 가스 터빈용 터빈 블레이드
PCT/EP2020/084603 WO2021110899A1 (fr) 2019-12-06 2020-12-04 Aube de turbine pour turbine à gaz stationnaire

Applications Claiming Priority (1)

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US20050084370A1 (en) * 2003-07-29 2005-04-21 Heinz-Jurgen Gross Cooled turbine blade
EP1783327A2 (fr) 2005-11-08 2007-05-09 United Technologies Corporation Refroidissement avec microcanaux périphériques en serpentin pour aube de turbine
US20100239431A1 (en) * 2009-03-20 2010-09-23 Siemens Energy, Inc. Turbine Airfoil Cooling System with Dual Serpentine Cooling Chambers
US8628298B1 (en) * 2011-07-22 2014-01-14 Florida Turbine Technologies, Inc. Turbine rotor blade with serpentine cooling
WO2016043742A1 (fr) * 2014-09-18 2016-03-24 Siemens Aktiengesellschaft Profil aérodynamique de turbine à gaz comprenant un passage de fluide de refroidissement intégré de bord d'attaque et d'extrémité et structure d'âme utilisée pour former un tel profil aérodynamique
WO2016076834A1 (fr) * 2014-11-11 2016-05-19 Siemens Aktiengesellschaft Aube de turbine muni de circuit de refroidissement de pointe axiale
WO2017039571A1 (fr) 2015-08-28 2017-03-09 Siemens Aktiengesellschaft Profil aérodynamique de turbine à refroidissement interne doté d'élément de déplacement d'écoulement

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Publication number Publication date
EP4048872B1 (fr) 2024-01-31
JP2023505451A (ja) 2023-02-09
WO2021110899A1 (fr) 2021-06-10
EP4048872A1 (fr) 2022-08-31
CN114787482A (zh) 2022-07-22
CN114787482B (zh) 2024-04-09
US12006838B2 (en) 2024-06-11
US20230358142A1 (en) 2023-11-09
KR20220103799A (ko) 2022-07-22

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