EP3263838A1 - Pale de turbine avec canal de refroidissement interne - Google Patents

Pale de turbine avec canal de refroidissement interne Download PDF

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Publication number
EP3263838A1
EP3263838A1 EP16177523.4A EP16177523A EP3263838A1 EP 3263838 A1 EP3263838 A1 EP 3263838A1 EP 16177523 A EP16177523 A EP 16177523A EP 3263838 A1 EP3263838 A1 EP 3263838A1
Authority
EP
European Patent Office
Prior art keywords
cooling
turbine blade
airfoil
cooling fluid
wall
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Withdrawn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
EP16177523.4A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
Radan RADULOVIC
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 EP16177523.4A priority Critical patent/EP3263838A1/fr
Publication of EP3263838A1 publication Critical patent/EP3263838A1/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Withdrawn legal-status Critical Current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
    • 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

Definitions

  • the invention relates to a turbine blade with an airfoil, in particular for use in a gas turbine with at least one substantially radially extending cooling channel.
  • An often used in gas turbines cooling medium is air.
  • a portion of the compressed air is taken from the compressor and fed to the gas turbine blades, bypassing the combustion chamber.
  • cooling channels in the interior of the turbine blades can be supplied with cooling air in the turbine blades, which are reshaped to the aerodynamically adapted outer shape of the blade and provide by appropriate deflection for an impact of the cooling air to an inner wall of the airfoil, whereby it is cooled ,
  • the blade tip and the remainder of the blade are made separately, and at the blade tip cooling means are provided for redirecting a cooling flow to a blade inner wall. Due to the further increasing demands on the turbine blades from the high-temperature and high-pressure operation, however, such a measure is no longer sufficient in all applications for cooling, which allows a satisfactory creep life of the turbine blade.
  • the object of the invention is therefore to provide a turbine blade, which enables improved impact cooling for turbine blades with extending inside the blade cooling channels.
  • a turbine blade according to the invention with an airfoil has at least one substantially radially extending cooling channel, which has a plurality of cooling chambers arranged one after the other along a radial extent of the airfoil, wherein radially adjacent cooling spaces are each separated by a partition, which has one or more cooling fluid openings for the passage of Has cooling fluid.
  • the invention is based inter alia on the idea that by drawing a partition in the cooling channel and the simultaneous provision of a cooling fluid opening in the partition, which is formed with a substantially smaller cross-section than the cooling chambers, a significant acceleration of the cooling fluid entails in turn amplifies the impact effect for the impact cooling, and thus enables a higher cooling capacity.
  • a cooling channel according to the invention extends radially from a coolant fluid interface in a blade root of the turbine blade into a blade tip of the turbine blade.
  • the individual cooling chambers are preferably arranged distributed along the radial extension of the blade root, the blade and / or the blade tip.
  • the at least one cooling fluid opening (in particular each partition wall) has a main direction which has an inclined position towards an inner wall of the airfoil.
  • a main direction of the cooling fluid opening is understood in particular to be that axis along which cooling fluid which has passed through the opening emerges from this.
  • the longitudinal axis of the bore may be the main direction.
  • the impact effect of the cooling fluid on the inner wall of the airfoil can be enhanced.
  • the at least one partition wall preferably all partitions, has a first group and a second group of cooling fluid openings, this first group and this second group having different main directions.
  • the partition wall is designed in this embodiment so that the first group of cooling fluid leads to a different area for cooling an inner wall of the airfoil than the second group.
  • a main direction of the first group of cooling fluid openings is directed obliquely toward a front inner wall of the airfoil and a main direction of the second group of cooling fluid openings is directed obliquely toward a rear inner wall of the airfoil.
  • This makes it possible to cool both sidewalls of the airfoil, ie a front wall and a rear wall, with a single partition arranged at a specific position of a radial axis of the turbine blade. If, for example, the front wall requires more cooling, this can be accommodated by a different design and / or a different number of cooling fluid openings of the first and the second group.
  • the at least one partition (preferably several or all partitions) is a straight rib, or an oblique rib and / or as roof-shaped rib formed.
  • the differently shaped partitions can be used in different turbine blades according to the invention, depending on the application-related requirements of the impact cooling.
  • the rib extends substantially perpendicular to a radial axis of the turbine blade in the cooling channel.
  • oblique in the sense of this embodiment is preferably to be understood an inclined position towards an inner wall of the airfoil, analogous to the above-described inclined position of the main direction of the cooling fluid openings.
  • Room-shaped is preferably to be understood as meaning a double inclination in the sense that an inclined position is realized on the one hand towards a front inner wall of the airfoil and on the other hand towards a rear inner wall of the airfoil by means of the partition, which then forms a gable.
  • a straight partition with straight cooling fluid openings can be provided.
  • a straight partition may also be formed with cooling fluid openings with an inclined position of their main direction towards an inner wall of the airfoil.
  • oblique or roof-shaped ribs in which the main directions of their cooling fluid openings are perpendicular to a surface of the rib, significantly cheaper and / or easier to manufacture.
  • the turbine blade has a cooling channel in the region of a blade front of the blade, that is to say, with respect to an axial direction of the turbine blade, facing the axial end of the blade which flows toward the working fluid.
  • the turbine blade has two or more cooling channels, which are arranged along an axial extension of the turbine blade, starting from the blade front of the airfoil, successively.
  • effective impingement cooling is possible along a larger portion of the axial extent of the turbine blade.
  • Such a turbine blade may be formed, for example, as a shrouded turbine blade, in which case preferably all cooling channels have the same radial flow direction, and in particular at the turbine base the cooling fluid is introduced into the cooling channels and, in particular, the cooling fluid is removed from the cooling channels on the blade shroud.
  • a turbine blade according to the invention may also be designed as a free turbine blade, in which case preferably adjacent cooling channels are connected to one another such that an opposite radial flow direction results.
  • the individual cooling channels are connected to each other at their ends to the coolant line, so that the coolant is both introduced at the blade root and discharged at a later arranged in the coolant flow cooling channel again.
  • the distance between adjacent partitions of a cooling channel can be varied as well as the inclination of the main directions of the coolant openings.
  • the turbine blade preferably has a plurality of cooling spaces in a cooling channel.
  • different distances between two partitions can be provided depending on a radial position of a cooling space on the blade be and / or different inclinations of the main directions of the coolant openings may be provided.
  • a smaller distance of the partition walls and / or a stronger inclined position towards the inner wall of the blade can be provided in the area of a central radial extent of the blade.
  • the quartz pins also provide the shape of the cooling fluid openings.
  • the cooling fluid openings are preferably formed with a circular cylindrical hole in a partition formed as a rib. A thickness of the ribs is preferably between 3 and 6 mm. If roof-shaped and / or inclined ribs are used, an angle of inclination to the associated inner wall of the airfoil of 45 ° to 75 ° is preferably provided.
  • FIG. 1 a turbine blade 1 is shown, comprising an airfoil 2 for the flow of a working fluid a gas turbine, in which the turbine blade is installed, and comprising a blade root 4, by means of which the turbine blade is fixed to the rotor shaft of the gas turbine.
  • a first cooling channel 6 is arranged, which - based on the cooling air flow 12 - is limited first by inner walls of the blade root 4 and then by a front inner wall 8 and a rear inner wall 10 of the turbine blade 2.
  • the cooling channel 6 extends substantially along a radial axis R of the turbine blade.
  • cooling channel 6 In the cooling channel 6 are - radially spaced from each other - two arranged as a roof-shaped ribs 14 partitions 16 are arranged, wherein in one embodiment, more than the two ribs shown are provided.
  • the partitions 16 delimit cooling chambers 28, 30 and 32 from each other, which are connected to each other with respect to the coolant flow 12 only via cooling fluid openings 18 and 20. At each of the partition walls 16, a plurality of cooling fluid openings 18, 20 is arranged.
  • the cooling fluid openings 18 have a main direction H 18 , which has an oblique position towards a front inner wall 8 of the airfoil at an angle ⁇ 18 .
  • the cooling fluid openings 20 have a main direction H 20 , which have an inclined position towards a rear inner wall 10 of the airfoil 2 at an angle ⁇ 20 .
  • cooling fluid openings 18 and three cooling fluid openings 20 are provided on the dividing wall 16.
  • the cooling fluid openings 18 have a corresponding main direction H 18 and are associated with a first group 118 of cooling fluid openings 18.
  • the cooling fluid openings 20 in turn have a corresponding main direction H 20 and are associated with a second group 120 of cooling fluid openings 20.
  • the working fluid of the gas turbine in which the turbine blade 1 is installed, this flows from a direction perpendicular to the representation towards the representation of FIG. 1 at.
  • a blade front 22 see. FIG. 2
  • the flow of working fluid is split by the turbine blade, wherein the energy of the working fluid flow is introduced into a front side 24 and a rear side 26 of the turbine blade 2 and heated.
  • this process takes place first on the partition 16.1 at a first radial position of the blade and then again on the partition 16.2 at a second radial position of the blade.
  • FIG. 2 It can be seen that not only the first cooling channel 6 is arranged in the axial direction along the longitudinal axis L of the turbine blade, but also further cooling channels 206 and 306 at positions further away from the blade front 22.
  • a plurality of partitions 16 are arranged, each one a plurality of roof-shaped ribs 14, in which in turn in each case a first group 118, 218, 318 of cooling fluid openings and a second group 120, 220, 320 of cooling fluid openings is arranged.
  • all three cooling channels 6, 206 and 306 are flowed through in the same direction from the blade root towards the blade casing or that the coolant flow in the first cooling channel 6 and the third cooling channel 306th flows radially outward, and is suitably connected to the second radial cooling passage 206, so that the radially outwardly promoted cooling fluid can be performed in this back to the blade root 4.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP16177523.4A 2016-07-01 2016-07-01 Pale de turbine avec canal de refroidissement interne Withdrawn EP3263838A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP16177523.4A EP3263838A1 (fr) 2016-07-01 2016-07-01 Pale de turbine avec canal de refroidissement interne

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP16177523.4A EP3263838A1 (fr) 2016-07-01 2016-07-01 Pale de turbine avec canal de refroidissement interne

Publications (1)

Publication Number Publication Date
EP3263838A1 true EP3263838A1 (fr) 2018-01-03

Family

ID=56292600

Family Applications (1)

Application Number Title Priority Date Filing Date
EP16177523.4A Withdrawn EP3263838A1 (fr) 2016-07-01 2016-07-01 Pale de turbine avec canal de refroidissement interne

Country Status (1)

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EP (1) EP3263838A1 (fr)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10626733B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10626734B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10633980B2 (en) 2017-10-03 2020-04-28 United Technologies Coproration Airfoil having internal hybrid cooling cavities
US10704398B2 (en) 2017-10-03 2020-07-07 Raytheon Technologies Corporation Airfoil having internal hybrid cooling cavities

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2260166A (en) * 1985-10-18 1993-04-07 Rolls Royce Cooled aerofoil blade or vane for a gas turbine engine
EP1154124A1 (fr) * 2000-05-10 2001-11-14 General Electric Company Aube refroidie par impact
EP2977557A1 (fr) * 2014-07-24 2016-01-27 United Technologies Corporation Structure d'aube refroidie et procédé de refroidissement associé
DE102014220787A1 (de) * 2014-10-14 2016-04-14 Siemens Aktiengesellschaft Gasturbinenbauteil mit Innenmodul und Verfahren zu seiner Herstellung unter Verwendung von Selektivem Laserschmelzen

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2260166A (en) * 1985-10-18 1993-04-07 Rolls Royce Cooled aerofoil blade or vane for a gas turbine engine
EP1154124A1 (fr) * 2000-05-10 2001-11-14 General Electric Company Aube refroidie par impact
EP2977557A1 (fr) * 2014-07-24 2016-01-27 United Technologies Corporation Structure d'aube refroidie et procédé de refroidissement associé
DE102014220787A1 (de) * 2014-10-14 2016-04-14 Siemens Aktiengesellschaft Gasturbinenbauteil mit Innenmodul und Verfahren zu seiner Herstellung unter Verwendung von Selektivem Laserschmelzen

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10626733B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10626734B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10633980B2 (en) 2017-10-03 2020-04-28 United Technologies Coproration Airfoil having internal hybrid cooling cavities
US10704398B2 (en) 2017-10-03 2020-07-07 Raytheon Technologies Corporation Airfoil having internal hybrid cooling cavities
US11649731B2 (en) 2017-10-03 2023-05-16 Raytheon Technologies Corporation Airfoil having internal hybrid cooling cavities

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