EP2725194A1 - Aube de rotor d'une turbine à gaz - Google Patents

Aube de rotor d'une turbine à gaz Download PDF

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Publication number
EP2725194A1
EP2725194A1 EP13190022.7A EP13190022A EP2725194A1 EP 2725194 A1 EP2725194 A1 EP 2725194A1 EP 13190022 A EP13190022 A EP 13190022A EP 2725194 A1 EP2725194 A1 EP 2725194A1
Authority
EP
European Patent Office
Prior art keywords
blade
overhang
suction
edge
sealing 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.)
Granted
Application number
EP13190022.7A
Other languages
German (de)
English (en)
Other versions
EP2725194B1 (fr
Inventor
Knut Dr. Lehmann
Manuel Herm
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.)
Rolls Royce Deutschland Ltd and Co KG
Rolls Royce PLC
Original Assignee
Rolls Royce Deutschland Ltd and Co KG
Rolls Royce PLC
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
Priority claimed from GB201219267A external-priority patent/GB201219267D0/en
Priority claimed from DE201210021400 external-priority patent/DE102012021400A1/de
Application filed by Rolls Royce Deutschland Ltd and Co KG, Rolls Royce PLC filed Critical Rolls Royce Deutschland Ltd and Co KG
Publication of EP2725194A1 publication Critical patent/EP2725194A1/fr
Application granted granted Critical
Publication of EP2725194B1 publication Critical patent/EP2725194B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • 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
    • 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
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/307—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the tip of a rotor blade

Definitions

  • the invention relates to a turbine rotor blade of a gas turbine with a blade profile formed in the radial direction (relative to an engine axis of the gas turbine) or in the longitudinal direction of the blade and with a blade tip.
  • a blade tip is referred to in the context of the present invention, the radially outer end of the turbine rotor blade.
  • the invention furthermore relates not only to rotor blades, but also to stator blades, wherein the blade tip in the case of stator blades is defined as a radially inner end of the blade.
  • the invention has for its object to provide a turbine rotor blade of the type mentioned, which allows for a simple structure and simple, inexpensive to manufacture an optimization of the leakage mass flow and has a good component strength.
  • the blade tip has at least on its suction side, starting from a stagnation point on the blade leading edge up to an intersection of the suction-side profile line of the blade with a trailing edge circle, an overhang (winglet).
  • the overhang at the stagnation point and at the intersection with the trailing edge circle is substantially zero and reaches its maximum value at about 40% of the run length of the suction side profile line.
  • the size of the suction-side overhang (vertical distance from the suction-side profile line) reaches approximately 45% of the diameter of the maximum circle T max which can be inscribed in the blade profile.
  • the blade tip on its pressure side starting from a stagnation point on the blade leading edge to an intersection of the pressure side profile line of the blade with the trailing edge circle, also has an overhang (winglet), which at the stagnation point and is substantially zero at the intersection and which has a maximum value at about a run length between 20% and 60% of the total run length of the pressure side profile line.
  • winglet overhang
  • a peripheral sealing edge is formed at the radially outer edge region of the blade (in the case of a rotor blade) or at the radially inner edge region in the case of a stator blade.
  • This can, for example, have a substantially rectangular cross-section, so that a depression / cavity is formed in the middle region of the blade tip.
  • the sealing edge can furthermore preferably have a region with a reduced height or a region with a height of zero, which is provided in the region of the suction-side overhang between a running length of the suction-side profile line of 10% to 30%.
  • an opening is formed through which an inflow of the boundary layer close to the housing can take place on the blade tip.
  • the radial height may be between half of the blade tip gap and the triple blade tip gap.
  • the width of the sealing edge this can be formed between the triple blade tip gap and the sixfold blade tip gap.
  • the height of the overhang (winglets) in the radial direction it can be particularly favorable if this height amounts to a maximum of 10% of the radial length of the blade profile.
  • a preferred value is 5%. This means that about 90% to 95% of the blade profile is formed unchanged and that only the outer 10 or 5% of the length of the blade profile is provided with the overhang or winglet according to the invention.
  • edge region of the overhang (winglets) at the radial end at an angle.
  • This angle is defined in a plane spanned by a radial vector from the sealing edge to the engine axis and a vector normal to the sealing edge. The angle then forms between a tangent to the outer sealing edge surface and the radial vector.
  • the gas turbine engine 10 is a generalized example of a turbomachine, in which the invention can be applied.
  • the engine 10 is formed in a conventional manner and comprises in succession an air inlet 11, a fan 12 circulating in a housing, a medium pressure compressor 13, a high pressure compressor 14, a combustion chamber 15, a high pressure turbine 16, a medium pressure turbine 17 and a low pressure turbine 18 and a Exhaust nozzle 19, which are all arranged around a central engine axis 1.
  • the intermediate pressure compressor 13 and the high pressure compressor 14 each include a plurality of stages, each of which includes a circumferentially extending array of fixed stationary vanes 20, commonly referred to as stator vanes, that radially inwardly from the engine casing 21 in an annular flow passage through the compressors 13, 14 protrude.
  • the compressors further include an array of compressor blades 22 projecting radially outwardly from a rotatable drum or disc 26 coupled to hubs 27 of high pressure turbine 16 and mid pressure turbine 17, respectively.
  • the turbine sections 16, 17, 18 have similar stages, comprising an array of fixed vanes 23 projecting radially inward from the housing 21 into the annular flow passage through the turbines 16, 17, 18, and a downstream array of turbine rotor blades 24 projecting outwardly from a rotatable hub 27.
  • the compressor drum or compressor disk 26 and the blades 22 disposed thereon and the turbine rotor hub 27 and the turbine rotor blades 24 disposed thereon rotate about the engine axis 1 during operation.
  • the Fig. 2 shows an end view of an embodiment of a turbine rotor blade according to the invention 24. It is understood that the frontal Surface is not flat, but part of a cylinder jacket around the engine axis 1. To simplify the illustration, the end surface is in each case flat in the following figures.
  • the Fig. 2 thus shows an inventive design of the rotor blade tip in plan view.
  • a feature of the invention the special shape of the suction-side overhang 30.
  • the inventive design of the suction-side overhang 30 is by means of Fig. 8 and 10 described in more detail.
  • the winglet overhang T w (s) is defined as a thickness distribution, ie as a vertical distance to the suction-side blade profile line.
  • the thickness distribution with the maximum profile thickness T max of the blade tip is made dimensionless.
  • the thickness distribution in Fig. 10 especially advantageous.
  • the thickness distribution is close to 0 (no significant overhang 30 is present).
  • the overhang 30 increases only very slowly along s.
  • two further thickness distributions (dashed lines shown), which thus narrow an area for the particularly advantageous design of the suction-side overhang 30.
  • a blade profile 29 is shown as a dashed line, this line corresponds to the blade profile under the winglet 30 at 90% of the blade height.
  • Line 38 shows the contour of the suction-side overhang ( Fig. 8 ) while the line 39 the contour of the pressure-side overhang ( Fig. 9 ) shows.
  • the reference numeral 31 indicates the circle which can be inscribed in the region of the maximum thickness of the cross section of the blade profile 29.
  • the reference numeral 32 shows the trailing edge circle.
  • the edge of the overhang 30 is formed in the form of a sealing edge 33, which is carried out substantially circumferentially. It has, as will be described below, an opening 34 ( FIGS. 12 and 13 ). While in Fig. 8 the suction-side overhang is shown and explained in detail, shows the Fig. 9 the pressure-side overhang with its contour 39.
  • the Fig. 4 to 7 each show sectional views along the in Fig. 3 shown cutting lines.
  • the thicknesses of the overhangs on the suction side and the pressure side are in the 10 and 11 shown.
  • the course is in each case applied over a dimensionless running length s, which extends from the stagnation point on the blade leading edge LE along the suction or pressure-side profile line to the intersection of the profile line with the trailing edge circle TE.
  • the size of the overhang T w (s) is normalized to the diameter of the maximum circle T max which can be inscribed in the blade profile. It follows at which points in each case the maximum values are provided particularly favorable.
  • the dashed lines in the 10 and 11 show a preferred design range, while the solid line represents an optimized solution.
  • FIG. 12 to 15 again gives a representation of the flow conditions.
  • the Fig. 13 shows in particular an inflow through the opening 34 and a flow through the blade tip gap 37. Accordingly, the FIGS. 14 and 15 for clarity, an example of a blade tip-splitting vortex 41 forming as well as a secondary flow vortex 42.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP13190022.7A 2012-10-26 2013-10-24 Aube de rotor d'une turbine à gaz Active EP2725194B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
GB201219267A GB201219267D0 (en) 2012-10-26 2012-10-26 Turbine blade
DE201210021400 DE102012021400A1 (de) 2012-10-31 2012-10-31 Turbinenrotorschaufel einer Gasturbine

Publications (2)

Publication Number Publication Date
EP2725194A1 true EP2725194A1 (fr) 2014-04-30
EP2725194B1 EP2725194B1 (fr) 2020-02-19

Family

ID=49448046

Family Applications (2)

Application Number Title Priority Date Filing Date
EP13190039.1A Active EP2725195B1 (fr) 2012-10-26 2013-10-24 Aube rotorique de turbine et étage rotorique associé
EP13190022.7A Active EP2725194B1 (fr) 2012-10-26 2013-10-24 Aube de rotor d'une turbine à gaz

Family Applications Before (1)

Application Number Title Priority Date Filing Date
EP13190039.1A Active EP2725195B1 (fr) 2012-10-26 2013-10-24 Aube rotorique de turbine et étage rotorique associé

Country Status (2)

Country Link
US (2) US10641107B2 (fr)
EP (2) EP2725195B1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3043715A1 (fr) * 2015-11-16 2017-05-19 Snecma Aube de turbine comprenant une pale avec baignoire comportant un intrados incurve dans la region du sommet de pale
EP3421725A1 (fr) * 2017-06-26 2019-01-02 Siemens Aktiengesellschaft Aube de compresseur
WO2019035800A1 (fr) * 2017-08-14 2019-02-21 Siemens Aktiengesellschaft Aubes de turbine

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US10641107B2 (en) 2012-10-26 2020-05-05 Rolls-Royce Plc Turbine blade with tip overhang along suction side
US9797258B2 (en) 2013-10-23 2017-10-24 General Electric Company Turbine bucket including cooling passage with turn
US9551226B2 (en) 2013-10-23 2017-01-24 General Electric Company Turbine bucket with endwall contour and airfoil profile
US9528379B2 (en) 2013-10-23 2016-12-27 General Electric Company Turbine bucket having serpentine core
US9638041B2 (en) 2013-10-23 2017-05-02 General Electric Company Turbine bucket having non-axisymmetric base contour
US9670784B2 (en) 2013-10-23 2017-06-06 General Electric Company Turbine bucket base having serpentine cooling passage with leading edge cooling
US20150110617A1 (en) * 2013-10-23 2015-04-23 General Electric Company Turbine airfoil including tip fillet
US20150345301A1 (en) * 2014-05-29 2015-12-03 General Electric Company Rotor blade cooling flow
US10508549B2 (en) * 2014-06-06 2019-12-17 United Technologies Corporation Gas turbine engine airfoil with large thickness properties
EP2977549B1 (fr) * 2014-07-22 2017-05-31 Safran Aero Boosters SA Aubage de turbomachine axiale et turbomachine associée
EP2977548B1 (fr) * 2014-07-22 2021-03-10 Safran Aero Boosters SA Aube et turbomachine associée
EP2987956A1 (fr) * 2014-08-18 2016-02-24 Siemens Aktiengesellschaft Aube de compresseur
US9995166B2 (en) 2014-11-21 2018-06-12 General Electric Company Turbomachine including a vane and method of assembling such turbomachine
US20160245095A1 (en) * 2015-02-25 2016-08-25 General Electric Company Turbine rotor blade
US10107108B2 (en) 2015-04-29 2018-10-23 General Electric Company Rotor blade having a flared tip
US10253637B2 (en) * 2015-12-11 2019-04-09 General Electric Company Method and system for improving turbine blade performance
FR3055698B1 (fr) * 2016-09-08 2018-08-17 Safran Aircraft Engines Procede de controle de la conformite du profil d'une surface courbe d'un element d'une turbomachine
JP6871770B2 (ja) * 2017-03-17 2021-05-12 三菱重工業株式会社 タービン動翼、及びガスタービン
EP3477059A1 (fr) * 2017-10-26 2019-05-01 Siemens Aktiengesellschaft Surface portante de compresseur
GB201719538D0 (en) * 2017-11-24 2018-01-10 Rolls Royce Plc Gas turbine engine
US11118462B2 (en) * 2019-01-24 2021-09-14 Pratt & Whitney Canada Corp. Blade tip pocket rib
US11066935B1 (en) 2020-03-20 2021-07-20 General Electric Company Rotor blade airfoil
US11371359B2 (en) 2020-11-26 2022-06-28 Pratt & Whitney Canada Corp. Turbine blade for a gas turbine engine
US20230349299A1 (en) * 2022-04-28 2023-11-02 Hamilton Sundstrand Corporation Additively manufactures multi-metallic adaptive or abradable rotor tip seals

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Cited By (12)

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Publication number Priority date Publication date Assignee Title
FR3043715A1 (fr) * 2015-11-16 2017-05-19 Snecma Aube de turbine comprenant une pale avec baignoire comportant un intrados incurve dans la region du sommet de pale
WO2017085387A1 (fr) * 2015-11-16 2017-05-26 Safran Aircraft Engines Aube de turbine de turbomachine, turbine et turbomachine associées
GB2560124A (en) * 2015-11-16 2018-08-29 Safran Aircraft Engines Turbine engine turbine vane, and related turbine and turbine engine
US10753215B2 (en) 2015-11-16 2020-08-25 Safran Aircraft Engines Turbine vane comprising a blade with a tub including a curved pressure side in a blade apex region
GB2560124B (en) * 2015-11-16 2022-04-13 Safran Aircraft Engines Turbine vane comprising a blade with a tub including a curved pressure side in a blade apex region
EP3421725A1 (fr) * 2017-06-26 2019-01-02 Siemens Aktiengesellschaft Aube de compresseur
WO2019001979A1 (fr) * 2017-06-26 2019-01-03 Siemens Aktiengesellschaft Profil aérodynamique de compresseur
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CN110869584B (zh) * 2017-06-26 2022-10-11 西门子能源环球有限责任两合公司 压气机翼型
WO2019035800A1 (fr) * 2017-08-14 2019-02-21 Siemens Aktiengesellschaft Aubes de turbine

Also Published As

Publication number Publication date
EP2725194B1 (fr) 2020-02-19
EP2725195B1 (fr) 2019-09-25
US10641107B2 (en) 2020-05-05
EP2725195A1 (fr) 2014-04-30
US9593584B2 (en) 2017-03-14
US20140119920A1 (en) 2014-05-01
US20140119942A1 (en) 2014-05-01

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