EP1857635A1 - Aube de turbine pour une turbine à gaz - Google Patents

Aube de turbine pour une turbine à gaz Download PDF

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Publication number
EP1857635A1
EP1857635A1 EP06010252A EP06010252A EP1857635A1 EP 1857635 A1 EP1857635 A1 EP 1857635A1 EP 06010252 A EP06010252 A EP 06010252A EP 06010252 A EP06010252 A EP 06010252A EP 1857635 A1 EP1857635 A1 EP 1857635A1
Authority
EP
European Patent Office
Prior art keywords
platform
turbine blade
blade
airfoil
edge
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Withdrawn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
EP06010252A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
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 EP06010252A priority Critical patent/EP1857635A1/fr
Priority to US11/803,495 priority patent/US20080107519A1/en
Publication of EP1857635A1 publication Critical patent/EP1857635A1/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Withdrawn legal-status Critical Current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
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/145—Means for influencing boundary layers or secondary circulations
    • 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
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • 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
    • F01D5/186—Film 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
    • 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
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • 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
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the invention relates to a turbine blade for a gas turbine, comprising a platform and a transversely extending, aerodynamically profiled airfoil comprising a suction side airfoil wall and a pressure side airfoil wall extending from an upstream side leading edge to a trailing side trailing edge, along the During operation, the airfoil blade walls can flow hot gas, wherein the platform and / or one of the two airfoil walls has or have at least two adjacent regions. Furthermore, the invention relates to a gas turbine with such a turbine blade.
  • a cooled blade of a gas turbine which has meandering running inside cooling channels.
  • delimiting inner walls of the airfoil turbulators are provided which improve the heat transfer from the blade material in the cavity flowing through the coolant. Due to the increased heat transfer, the turbine blade can permanently withstand higher operating temperatures.
  • cracks can occur in the region of the hollow-throat-like transition from platform to blade, which is also referred to as "fillet", and / or in the platform due to impermissibly large temperature gradients. If the resulting cracks exceed a critical crack length, safe operation of the gas turbine equipped with such a turbine blade is not ensured.
  • a particularly long service life of the turbine blade is a design goal with which the availability of a gas turbine equipped therewith can be further increased.
  • the object of the invention is accordingly to provide a turbine blade for a gas turbine whose life is extended.
  • the task directed to the turbine blade is achieved with a generic turbine blade, which is designed according to the features of claim 1.
  • the invention is based on the finding that the wear and the crack formation and the subsequent crack growth are thermally induced.
  • the material of the turbine blade is exposed to thermal stresses, since significantly different material temperatures occur in at least two adjoining regions of the turbine blade. Usually, these areas are influenced by suitable measures, for example by taking place in the interior of the turbine blade cooling such that they withstand the temperatures.
  • the cooling of the areas is often flat and thus can not be adjusted to local requirements, for example due to convective cooling, the thermally differently loaded areas are cooled evenly, resulting in particularly large temperature gradients in the blade material.
  • these high temperature gradients lead to life-shortening cracking and crack growth.
  • the two adjoining material regions of the turbine blade now have different hot gas side (ie the occurring between the blade material and the hot gas) heat transfer coefficients to match the thermal stresses occurring during operation of the areas. Due to the sometimes difficult to adjust cold side (ie occurring between the blade material and the coolant) heat transfer coefficient now the hot gas heat transfer coefficient is adjusted for equalization of the temperatures in the blade material for the first time. As a result, the local thermal stresses that occur between the two areas can be significantly reduced. To achieve this, it is provided in a generic turbine blade that a Means for adjusting the local hot gas heat transfer coefficient is arranged in one of the areas. Due to the equalized thermal stress between the two adjacent areas cracks occur less frequently than before.
  • a particularly durable turbine blade can be specified, with which also the disposal period of a gas turbine equipped with it is further increased.
  • the proposed measure extends the low cycle fatigue life (LCF) for the platform and its transition into the airfoil, ie in the fillet.
  • the proposed measure leads to increased heat input due to increased turbulence in the flow, whereby the difference in material temperatures of higher loaded area and lower loaded area is reduced. Due to the reduced difference in the material temperatures, a lower thermal stress occurs between the two areas, so that a homogenization of the material temperature is achieved, which has a life-extended effect on the turbine blade.
  • region of the platform has the swirling elements which adjoin the suction-side airfoil wall.
  • the swirl elements are provided in the middle region of the platform, which, viewed in the direction of flow of the hot gas, lies between the front edge and the rear edge.
  • Airfoil comparatively high temperature gradients along the platform, which favor the defects such as cracking and crack growth. Accordingly, the provision of the agents according to the invention in this area is of particular advantage. Since the cracking also occurs at the edge of the platform, the swirling elements can also be provided at this point.
  • Turbulators dimples, ribs or pins are used as swirling elements. These known configurations serve to increase the turbulence of the passing hot gas so as to enlarge the hot gas side heat transfer coefficient.
  • the platform can also have a corrugated surface as means for setting different heat transfer coefficients, the wavefront of the wave form being oriented transversely, preferably vertically, to the flow direction of the hot gas.
  • the wave crests of the waveform serve to slightly increase the turbulence in the hot gas flow, thereby slightly increasing the heat transfer coefficient on the hot gas side surface.
  • the wave troughs of the waveform occurs a slightly slowed flow, whereby the heat transfer coefficient is slightly lowered at this point.
  • the troughs and the wave crests are arranged in areas in which previously different thermal stresses occurred. Consequently, an approximation of the thermal stresses occurring during operation can also be achieved with a preferably slightly wave-like surface.
  • the platform preferably has a first transverse edge which can be anvoked by the hot gas and a second transverse edge which lies opposite the first transverse edge, the platform having only two concave troughs along the longitudinal extension of the platform between the first and second transverse edges.
  • the proposed measures have been found to be particularly efficient when the turbine blade has been produced in a casting process, that is cast, and in which the blade and / or the platform is coolable, preferably convective cooling is or is.
  • convectively cooled turbine blades experience a comparatively uniform cooling along the cooling channel, in which the coolant, usually cooling air, flows.
  • the adaptation of the hot gas heat transfer coefficient is particularly appropriate here.
  • the turbine blade is also heat insulation layer free and thus designed to withstand material temperatures of 850 ° C to 1000 ° C. These temperatures usually occur in the second or third turbine stage of a stationary, used for power generation gas turbine.
  • a designed as a blade turbine blade 10 is shown in a perspective view, which has a hammer-shaped in cross section mounting foot 12 for receiving in a groove, not shown, of the rotor disk of the rotor of a gas turbine.
  • a platform 14 connects, which radially the flow channel of the turbine, d. H. bounded transversely to the direction of the Z-axis.
  • an airfoil 18 extending transversely to the platform 14, which comprises a suction-side airfoil wall 20 and a pressure-side airfoil wall 22 which extend from an upstream-side leading edge 24 to a downstream trailing edge 26.
  • the hot gas 28 flowing through the turbine during operation which flows around the airfoil 18 essentially in the axial direction X, is referred to.
  • the turbine blade 10 is uncoated, i. it has no thermal barrier layer, and is intended for use in the second or third turbine stage of the stationary gas turbine
  • the blade 18 of the turbine blade 10 is partially hollow and has two separated by a support rib 30 cavities 32, which can be traversed in parallel by a fastening side supplied coolant, preferably cooling air 36 in the radial direction Z.
  • the airfoil 18 of a generic turbine blade heats up to about 850 ° C to 1000 ° C, and it is cooler, especially at its trailing edge 26 near the platform 14 in a pressure-side region 38 as opposed Suction-side region 40.
  • the region 38 is generally more than 130 ° C cooler than the region 40th
  • 26 turbulators 44 are provided according to the invention to equalize the thermal stress on the suction side region 38 at the trailing edge 26, the heat transfer coefficient ⁇ , based on the other of the two areas 40, increased, which also increases the heat input from the hot gas 28 into the blade material, contrary to the otherwise usual efforts. Due to the larger heat input 38, the blade material is warmer in this area than without the arrangement of turbulators 44. However, the permissible material temperature is not exceeded. Since the opposite region 40 is much more heavily loaded in operation on the suction side airfoil wall 20, ie generally warmer than 130.degree.
  • the swirling elements 42 may be designed as turbulators 44, dimples, as ribs or pins and be made directly during the casting of the turbine blade 10 with.
  • Turbulators 44 can be used both as rib-shaped, d. H. be substantially along its longitudinal extension straight ribs or be designed as crescent-shaped ribs.
  • FIG. 2 shows a cross section through the blade 18 of the turbine blade 10 as a plan view, wherein the blade 18 now has four cavities 32, which are sequentially flowed through by cooling air 36.
  • thermal stresses occur due to the hot gas flowing along 28, which depends on the suction-side width of the platform 14 seen in the circumferential direction Y - between the edge 50 of the platform and the suction-side airfoil wall 20.
  • the suction-side width of the platform 14 is greater than in a second region B.
  • the suction-side width between the platform edge 50 and the blade 18 increases again. In these areas occur due to the cooling of the airfoil 18 different thermal gradients that could previously lead to defects.
  • area B was previously affected by crack growth.
  • the surface 16 of the platform 14 is provided with locally swirling elements 42 in the form of turbulators 44, which increase the heat input of the hot gas 28 flowing past it into the turbine blade material.
  • the temperature difference between the first region A and the third region C, which were hitherto more strongly temperature-stressed, and the region B, which was previously less thermally stressed, can be significantly reduced, as a result of which the thermal stress between the regions A, B, C is evened out. Cracking and crack growth can be effectively avoided, resulting in a longer life for the turbine blade 10.
  • FIG. 3 shows the temperature profile T in the platform material 14 along the axial direction X.
  • the temperature T A of the region A of the inflow-side transverse edge 52 is comparatively high, for example 850 ° C., and decreases in the direction of the hot gas 28 flowing along up to a temperature minimum T B , which is to be found in area B. From there, the material temperature again increases to a mean temperature value T c , which occurs in operation in the area of the downstream transverse edge 54 of the platform 14.
  • the temperatures can be measured with a suitable measuring method or can be determined simulatively with the aid of a finite element calculation program. So far, the temperature difference between the region A and the region B in the order of magnitude greater than 130 ° C.
  • the heat transfer coefficient ⁇ as a function of the X-axis.
  • the hot gas side heat transfer coefficient ⁇ is greater than in the regions A and C, which are to be found at the inlet end 52 of the platform 14 and at the outlet end 54 of the platform 14.
  • This wave-like characteristic of the heat transfer coefficient ⁇ along the platform 14 goes back to the swirling elements 42, which are provided in the area B for the adjustment of the material temperatures of the platform 14.
  • FIG 5 shows in an alternative embodiment, the surface 16 of the platform 14 along the section line VV of FIG 2.
  • the surface 16 is wavy, so that they height of the platform, viewed in the radial direction Z, in the area B. is increased compared to the areas A and C.
  • a maximum B is provided between the troughs A and C, which also leads to a demand-adapted heat transfer coefficient ⁇ .
  • the wavefront of the undulating surface 16 of the platform 14 may be transverse to the flow direction of the hot gas 28 or perpendicular to the platform edge 50th
  • the lifetime of the turbine blade 10 according to the invention compared to a generic turbine blade can be significantly extended, since in one of at least two previously different thermal load areas an adaptation of the hot gas heat transfer coefficient ⁇ to equalize the thermal stresses and material temperatures is proceeding.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP06010252A 2006-05-18 2006-05-18 Aube de turbine pour une turbine à gaz Withdrawn EP1857635A1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP06010252A EP1857635A1 (fr) 2006-05-18 2006-05-18 Aube de turbine pour une turbine à gaz
US11/803,495 US20080107519A1 (en) 2006-05-18 2007-05-15 Turbine blade for a gas turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP06010252A EP1857635A1 (fr) 2006-05-18 2006-05-18 Aube de turbine pour une turbine à gaz

Publications (1)

Publication Number Publication Date
EP1857635A1 true EP1857635A1 (fr) 2007-11-21

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ID=37101833

Family Applications (1)

Application Number Title Priority Date Filing Date
EP06010252A Withdrawn EP1857635A1 (fr) 2006-05-18 2006-05-18 Aube de turbine pour une turbine à gaz

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US (1) US20080107519A1 (fr)
EP (1) EP1857635A1 (fr)

Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8157527B2 (en) * 2008-07-03 2012-04-17 United Technologies Corporation Airfoil with tapered radial cooling passage
JP2012233406A (ja) 2011-04-28 2012-11-29 Hitachi Ltd ガスタービン静翼
WO2014175937A2 (fr) 2013-02-05 2014-10-30 United Technologies Corporation Pièce de turbine à gaz comportant un turbulateur incurvé
US10358978B2 (en) 2013-03-15 2019-07-23 United Technologies Corporation Gas turbine engine component having shaped pedestals
US10196903B2 (en) * 2016-01-15 2019-02-05 General Electric Company Rotor blade cooling circuit

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE681479C (de) * 1937-03-04 1939-09-23 Escher Wyss Maschinenfabrik G Leitschaufel fuer Dampf- oder Gasturbinen, insbesondere fuer den Niederdruckteil solcher Turbinen
GB944166A (en) * 1960-03-02 1963-12-11 Werner Hausammann Rotor for turbines or compressors
US4023350A (en) 1975-11-10 1977-05-17 United Technologies Corporation Exhaust case for a turbine machine
US4822249A (en) * 1983-07-15 1989-04-18 Mtu Motoren-Und Turbinen-Union Munich Gmbh Axial flow blade wheel of a gas or steam driven turbine
EP0976928A2 (fr) * 1998-07-31 2000-02-02 DLR Deutsches Zentrum für Luft- und Raumfahrt e.V. Ensemble de pale pour turbomachine
US6213711B1 (en) * 1997-04-01 2001-04-10 Siemens Aktiengesellschaft Steam turbine and blade or vane for a steam turbine
EP1178181A2 (fr) 2000-07-31 2002-02-06 General Electric Company Refroidissement en série pour aubes de turbine
EP1270872A1 (fr) * 2000-03-27 2003-01-02 Honda Giken Kogyo Kabushiki Kaisha Turbine a gaz
EP1469163A2 (fr) 2000-02-23 2004-10-20 Mitsubishi Heavy Industries, Ltd. Aube rotorique de turbine a gaz

Family Cites Families (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2281356B (en) * 1993-08-20 1997-01-29 Rolls Royce Plc Gas turbine engine turbine
GB9823840D0 (en) * 1998-10-30 1998-12-23 Rolls Royce Plc Bladed ducting for turbomachinery
US6183197B1 (en) * 1999-02-22 2001-02-06 General Electric Company Airfoil with reduced heat load
US6722134B2 (en) * 2002-09-18 2004-04-20 General Electric Company Linear surface concavity enhancement
US7484935B2 (en) * 2005-06-02 2009-02-03 Honeywell International Inc. Turbine rotor hub contour

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE681479C (de) * 1937-03-04 1939-09-23 Escher Wyss Maschinenfabrik G Leitschaufel fuer Dampf- oder Gasturbinen, insbesondere fuer den Niederdruckteil solcher Turbinen
GB944166A (en) * 1960-03-02 1963-12-11 Werner Hausammann Rotor for turbines or compressors
US4023350A (en) 1975-11-10 1977-05-17 United Technologies Corporation Exhaust case for a turbine machine
US4822249A (en) * 1983-07-15 1989-04-18 Mtu Motoren-Und Turbinen-Union Munich Gmbh Axial flow blade wheel of a gas or steam driven turbine
US6213711B1 (en) * 1997-04-01 2001-04-10 Siemens Aktiengesellschaft Steam turbine and blade or vane for a steam turbine
EP0976928A2 (fr) * 1998-07-31 2000-02-02 DLR Deutsches Zentrum für Luft- und Raumfahrt e.V. Ensemble de pale pour turbomachine
EP1469163A2 (fr) 2000-02-23 2004-10-20 Mitsubishi Heavy Industries, Ltd. Aube rotorique de turbine a gaz
EP1270872A1 (fr) * 2000-03-27 2003-01-02 Honda Giken Kogyo Kabushiki Kaisha Turbine a gaz
EP1178181A2 (fr) 2000-07-31 2002-02-06 General Electric Company Refroidissement en série pour aubes de turbine

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Publication number Publication date
US20080107519A1 (en) 2008-05-08

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