WO2014009077A1 - 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
WO2014009077A1
WO2014009077A1 PCT/EP2013/061957 EP2013061957W WO2014009077A1 WO 2014009077 A1 WO2014009077 A1 WO 2014009077A1 EP 2013061957 W EP2013061957 W EP 2013061957W WO 2014009077 A1 WO2014009077 A1 WO 2014009077A1
Authority
WO
WIPO (PCT)
Prior art keywords
blade
rib
leading edge
turbine
turbine blade
Prior art date
Application number
PCT/EP2013/061957
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
Hans-Thomas Bolms
Original Assignee
Siemens Aktiengesellschaft
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 Aktiengesellschaft filed Critical Siemens Aktiengesellschaft
Publication of WO2014009077A1 publication Critical patent/WO2014009077A1/fr

Links

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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall

Definitions

  • the invention relates to a turbine blade for a gas turbine having an aerodynamically curved airfoil having a suction side sidewall and a pressure side sidewall extending along a chord from a common leading edge to a trailing edge and in a spanwise direction with a total span from a blade root end to a gas turbine engine blade extend the blade end, wherein the blade is designed to guide a coolant hollow and in the interior of a parallel to the front edge extending rib with openings suitable for impingement cooling of the inside of the leading edge is provided.
  • casting cores Lead the leading edge of the turbine blades.
  • such geometries lead to the required casting cores being comparatively complex.
  • the complexity of the casting cores with comparatively filigree designs requires a certain fragility, so that the handling of the casting cores to avoid fractures requires special care.
  • such casting cores can also be used when filling the Casting device with hot liquid metal break, which increases the reject rate of the manufacturing process.
  • the object of the invention is therefore to provide a turbine blade in which the disadvantages known from the prior art are reduced, if not avoided.
  • a turbine blade for a gas turbine engine comprising an aerodynamically curved airfoil having a suction side wall and a pressure side side wall extending along a chord from a leading edge to a trailing edge and in a spanwise direction with a total span from a shoe-tail end extend to a blade end, wherein the blade is designed to guide a coolant hollow and in the interior of which is provided parallel to the leading edge extending rib with openings suitable for impingement cooling the inside of the leading edge, it is proposed that the length of the rib in Spannweiteraum smaller is 80% of the total span and the rib is located approximately midway between the two rib ends.
  • the proposed solution reduces the complexity of the structure inside the turbine blade, with the result that the casting core required for this purpose is likewise less complex.
  • the reduced complexity then leads to increased stability of the casting core, so that it is less prone to breakage.
  • the reduced tendency to fracture improves handling and also the reject rate.
  • an adaptation of the cooling to the actually required cooling requirements is achieved with the embodiment according to the invention. It has been shown that in Spannweitecardi - ie in an axially flowed stationary gas turbine in the radial direction - the highest temperature load in the range of the central span occurs and to the edge sections, ie at an annular hot gas duct radially outside or radially inside, is less. In the course of this, a locally adapted cooling of the leading edge of the blade can be achieved with the proposed turbine blade.
  • no end of the two ends of the ribs, which face the blade root and the blade head, are integrally connected to the inside of the front edge by a transition region.
  • the unconnected ends lead to a comparatively large space between the rib ends and the inside of the front edge, which can be produced by a casting core which is thickened at this point. This measure increases the stability of the casting core and may ultimately lead to a lower reject rate in the manufacture of such turbine blades.
  • turbine blades have leading edge channels for coolant between the inside of the leading edge and the rib, which can be set up for different cooling concepts. Due to the existing free spaces between the ends of the rib and the inside of the leading edge, a purely convective cooling of the inside of the leading edge, possibly configured with turbulators, can be achieved.
  • the "crossover holes" present in the rib are then without function, but a turbine blade produced with the same casting device and the same casting cores can also be used for impact cooling of the billets. derkante be corrected. For this purpose, only after the casting of the turbine blade, the free space between one end of the rib and the inside of the front edge with the help of a plug to close later. The latter turbine blades are then capable of withstanding higher temperatures than the convectively cooled ones.
  • an end of the rib facing a blade root or an end of the rib facing a blade head forms with the inside of the leading edge an opening present after casting, which is closed in each case by means of a separately produced plug.
  • a purely convective cooling of the leading edge should be provided according to an advantageous development in the integral transition region openings for discharging a coolant from the space between the inside of the leading edge and the rib.
  • the turbine blade may be formed as a blade or as a vane.
  • FIG. 1 shows a schematic representation of a turbine blade in the form of a guide blade
  • FIG. 2 shows the longitudinal section through the blade of a turbine blade according to FIG. 1 in a first embodiment
  • FIG. 1 shows a side view of a turbine blade 10 designed as a guide vane of a first gas turbine stage.
  • the turbine blade 10 comprises a foot-side platform 12 for fastening the turbine blade 10 to a turbine guide vane carrier (not shown). Furthermore, it comprises an airfoil 14 which adjoins the platform 12 in spanwise direction. In the operating position of the turbine blade 10 in an axially flowing gas turbine, a head-side platform 15 adjoins the inner end of the blade 14.
  • the airfoil 14 includes a leading edge 16 from which a pressure-side sidewall 20 and a suction-side sidewall extend toward a trailing edge 22.
  • the airfoil 14 extends in a spanwise direction from a foot-side end 24 to a head-side end 26. Along this spanwise direction, the airfoil has a total span that is normalized to 100%, the origin being blade-sided - that is, at the transition point from platform 12 to shuffle - Foot end 24 - is arranged.
  • FIG. 2 shows the section through the front edge region of the turbine blade 10 according to FIG. 1.
  • the blade end-side blade end 24 is shown above the blade tip end 26, as in FIG.
  • the point provided with the reference numeral 21 is arranged centrally between the base foot-side end 24 and the blade end 26 of the airfoil 14. In this respect, the point 21 lies at 50% of the total span starting from the blade end 24.
  • a rib 30 extends parallel to the front edge 16. This rib 30 internally connects the pressure-side side wall 20 with the suction-side side wall, so that a front edge space 34 is formed between an inner side 32 of the front edge 16 and the rib 30.
  • the rib 30 also has an end 36 facing the blade root and an end 38 facing the blade head. The the scoop-head end 38 merges with the leading edge 16 via a transition region 40.
  • the rib 30 and the transition region 40 are integrally formed with the remainder of the airfoil 14, they are integral.
  • at least one opening, or as shown by way of example, two openings 42 for discharging a coolant 44 from the leading edge space 34 is provided.
  • the hollow space of the turbine blade 10 can be supplied with cooling air 44 from the base side. A portion of this cooling air 44 can flow into the leading edge space 34 and thereby cool the front edge 16 from the inside. The cooling air 44 flowing into the leading edge space 34 then exits via the openings 42.
  • turbulators 46 are provided on the inner sides 32 of the front edge 16 and / or inner sides of the side walls 20 in order to increase the heat transfer.
  • the opening 48 formed by the blade end 36 of the rib 30 and the leading edge 16 is closed by a separately produced plug 50 (FIG. 3).
  • exit holes 52 are introduced into the front edge 16.
  • a cooling air 44 flowing inside will flow into the leading edge space 34 through openings 54, 42 arranged in the rib, thereby preliminarily cooling the front edge 16 in the section of the rib 30.
  • the heated cooling air can escape through the subsequently produced outlet holes 52 from the turbine blade 10.
  • the openings 42 can also be closed by plugs.
  • the invention thus relates to a turbine blade 10 for a gas turbine, with an aerodynamically curved blade 14, which has a suction-side side wall and a pressure-side side wall 20 extending along a chord from a common leading edge 16 to a trailing edge 22 and in a Spannweiteraum with a Overall span ranging from a blade end 24 to a blade end 26, wherein the blade 14 is designed to guide a coolant hollow and in the interior of a parallel to the front edge 16 extending rib 30 with openings 54 suitable for impingement cooling the inside 32 of the leading edge 16 is provided is.
  • Spanweitecardi is less than 80% of the total span and the rib is approximately centrally located between its two ends 24, 26.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

La présente invention concerne une aube de turbine (10) pour une turbine à gaz, comportant une aube (14) aérodynamiquement bombée présentant une paroi latérale côté aspiration et une paroi latérale côté pression (20) qui s'étendent le long d'une corde de profil, d'un bord d'attaque (16) à un bord de fuite (22), et dans la direction de l'envergure, avec une envergure commune, d'une extrémité côté pied d'aube (24) à une extrémité côté tête d'aube (26). L'aube (14) est creuse afin de pouvoir guider un liquide de refroidissement, et à l'intérieur de ladite aube, une nervure (30) présentant des ouvertures (54) s'étend parallèlement au bord d'attaque (16) pour obtenir un refroidissement par impact du côté intérieur (32) du bord d'attaque (16). Selon l'invention, afin d'obtenir une aube de turbine (10) permettant différentes configurations de refroidissement, et pour laquelle le taux de rebuts est réduit en raison d'un noyau de coulée plus stable, la longueur de la nervure (30) dans la direction de l'envergure est inférieure à 80% de l'envergure totale, et la nervure est située sensiblement au milieu, entre les deux extrémités (24, 26).
PCT/EP2013/061957 2012-07-13 2013-06-11 Aube de turbine pour une turbine à gaz WO2014009077A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE201210212289 DE102012212289A1 (de) 2012-07-13 2012-07-13 Turbinenschaufel für eine Gasturbine
DE102012212289.3 2012-07-13

Publications (1)

Publication Number Publication Date
WO2014009077A1 true WO2014009077A1 (fr) 2014-01-16

Family

ID=48652032

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2013/061957 WO2014009077A1 (fr) 2012-07-13 2013-06-11 Aube de turbine pour une turbine à gaz

Country Status (2)

Country Link
DE (1) DE102012212289A1 (fr)
WO (1) WO2014009077A1 (fr)

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10480327B2 (en) 2017-01-03 2019-11-19 General Electric Company Components having channels for impingement cooling

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19801804A1 (de) * 1998-01-19 1999-07-22 Siemens Ag Turbinenschaufel sowie Verfahren zur Inspektion und/oder Reinigung einer Turbinenschaufel
EP1154124A1 (fr) * 2000-05-10 2001-11-14 General Electric Company Aube refroidie par impact
US20080019841A1 (en) * 2006-07-21 2008-01-24 United Technologies Corporation Integrated platform, tip, and main body microcircuits for turbine blades
EP1882817A2 (fr) * 2006-07-27 2008-01-30 General Electric Company Aube de dôme de trou à poussière

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5931638A (en) * 1997-08-07 1999-08-03 United Technologies Corporation Turbomachinery airfoil with optimized heat transfer

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19801804A1 (de) * 1998-01-19 1999-07-22 Siemens Ag Turbinenschaufel sowie Verfahren zur Inspektion und/oder Reinigung einer Turbinenschaufel
EP1154124A1 (fr) * 2000-05-10 2001-11-14 General Electric Company Aube refroidie par impact
US20080019841A1 (en) * 2006-07-21 2008-01-24 United Technologies Corporation Integrated platform, tip, and main body microcircuits for turbine blades
EP1882817A2 (fr) * 2006-07-27 2008-01-30 General Electric Company Aube de dôme de trou à poussière

Also Published As

Publication number Publication date
DE102012212289A1 (de) 2014-01-16

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