WO2017125289A1 - Agencement de profil aérodynamique - Google Patents

Agencement de profil aérodynamique Download PDF

Info

Publication number
WO2017125289A1
WO2017125289A1 PCT/EP2017/050449 EP2017050449W WO2017125289A1 WO 2017125289 A1 WO2017125289 A1 WO 2017125289A1 EP 2017050449 W EP2017050449 W EP 2017050449W WO 2017125289 A1 WO2017125289 A1 WO 2017125289A1
Authority
WO
WIPO (PCT)
Prior art keywords
aerofoil
platform
arrangement
cooling medium
deflector
Prior art date
Application number
PCT/EP2017/050449
Other languages
English (en)
Inventor
Anthony Davis
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 WO2017125289A1 publication Critical patent/WO2017125289A1/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
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • 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
    • 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
    • F05D2240/81—Cooled platforms
    • 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/23—Three-dimensional prismatic
    • 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/25—Three-dimensional helical
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to guide vanes for gas turbines and, more particularly, to guide vanes for gas turbines comprising aerofoils provided with internal cooling passages.
  • air is pressurized in a compressor and mixed with fuel in a combustor for generating hot
  • combustion gases are then channelled towards a turbine which transforms the energy from the hot gases into work for powering the compressor and other devices which converts power, for example an upstream fan in a typical aircraft turbofan engine application, or a generator in power generation application.
  • the turbine stages include stationary turbine nozzles having a row of vanes which channel the combustion gases into a corresponding row of rotor blades extending radially
  • the vanes and blades may have corresponding hollow aerofoils. Aerofoils may be designed and manufactured hollow in order to save weight, to change its eigenfrequency or to include a cooling circuit therein. In the latter case, the cooling gas which circulates inside the cooling circuits is typically bleed air from the compressor discharge.
  • the present invention relates to a turbine hollow aerofoil which can be used in a gas turbine guide vane and which include an internal passage for cooling the aerofoil.
  • upstream and downstream refer to the flow direction of the airflow and/or working hot gas flow through the gas turbine engine.
  • axial and radial are made with reference to a rotational axis of the gas turbine engine.
  • Each aerofoil includes a generally concave pressure sidewall and, an opposite, generally convex suction sidewall extending radially outwardly along a span from an aerofoil base to an aerofoil tip and axially in a chordwise direction between a leading and a trailing edge.
  • the aerofoil extends from a root base integral with a radially inner platform of a turbine stator to a radially outer tip integral with an outer platform of the outer casing of the stator.
  • a cooling passage provided in the aerofoil and oriented radially, for example from the outer tip towards the base and the inner platform.
  • the flow of the cooling gases is required to turn through 90 degrees within a short distance, in order to provide a flow of cooling air along the platform surface. This sharp turn generates a large pressure loss, high local turbulence, and poor control of the flow, leading to
  • an aerofoil for a guide vane of a gas turbine and a guide vane for a gas turbine are provided in accordance to the independent claims.
  • the dependent claims describe advantageous developments and modifications of the invention.
  • an aerofoil for a gas turbine comprises:
  • a guide vane for a gas turbine includes:
  • the design provided by the present invention reduces the pressure loss coefficient of the internal cooling passage with respect to the prior art solutions in which the cooling medium ejected radially from the aerofoil impinges onto a plate, so causing large dynamic pressure losses.
  • the first radial end of the aerofoil is a base of the aerofoil, connected to an inner platform of the stator of the turbine.
  • the second radial end of the aerofoil is a tip of the aerofoil, connected to an outer platform of the stator of the turbine.
  • the internal passage inside the aerofoil extends from the tip to the base for cooling the inner platform of the turbine stator.
  • the first radial end of the aerofoil is a tip of the aerofoil, connected to an outer platform of the stator of the turbine.
  • the second radial end of the aerofoil is a base of the aerofoil, connected to an inner platform of the stator of the turbine.
  • the internal passage inside the aerofoil extends from the base to the tip for cooling the outer platform of the turbine stator.
  • the aerofoil further comprises an inner insert or deflector provided in the internal passage and extending from the first radial end towards the second radial end, the inner insert being connected to the internal passage for receiving the cooling medium, the inner insert being shaped in such a way that at the first radial end the cooling medium exits the aerofoil according to a flow direction having at least a component orthogonal to the radial direction.
  • the deflector may be manufactured separately to the aerofoil and inserted to the aerofoil after the aerofoil has been manufactured, for example the deflector might not be formed during a casting process of the aerofoil.
  • the use of an insert inside the aerofoil cooling internal passage permits to simplify the
  • the insert comprises at least a surface inclined with respect to the radial direction.
  • the insert may
  • the spiral surface may extend from to the other of the first radial end and the second radial end.
  • the insertion of a spiral surface inside the hollow aerofoil gives the cooling medium a swirling motion as it passes along the aerofoil cooling internal passage.
  • the flow has a velocity largely perpendicular to the radial axis of the turbine and so parallel to the platform surface. This allows an efficient cooling of that surface by convection .
  • the insert comprising the inclined surface has a triangular section in a sectional plane parallel to the radial
  • the triangular section of the insert may include a side lying along the first radial end of the aerofoil and a vertex oriented towards the second radial end of the aerofoil.
  • a simple flow deflector is used at the first radial end of the aerofoil for turning the flow of the cooling medium as it exits the aerofoil .
  • FIG. 1 shows part of a turbine engine in a sectional view and in which the present inventive aerofoil is incorporated
  • FIG. 2 shows an exploded view of a guide vane aerofoil for a gas turbine according to the present invention
  • FIG. 3 shows an isometric partial view of a guide vane for a gas turbine including the aerofoil of FIG. 2,
  • FIG. 4 shows a top view of the guide vane of Fig. 3,
  • FIG. 5 shows an isometric partially-sectioned view of the guide vane of Fig. 3,
  • FIG. 6 shows a front view of the guide vane of Fig. 3,
  • FIG. 7 shows an exploded view of another embodiment of a guide vane aerofoil for a gas turbine according to the present invention.
  • FIG. 1 shows an example of a gas turbine engine 10 in a sectional view.
  • the gas turbine engine 10 comprises, in flow series, an inlet 12, a compressor section 14, a combustor section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis 20.
  • the gas turbine engine 10 further comprises a shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine engine 10.
  • the shaft 22 drivingly connects the turbine section 18 to the
  • air 24 which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16.
  • the burner section 16 comprises a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28.
  • the combustion chambers 28 and the burners 30 are located inside the burner plenum 26.
  • the compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the burner 30 and is mixed with a gaseous or liquid fuel.
  • the air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.
  • This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment.
  • An annular array of transition duct outlets form an annulus for
  • the turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22.
  • two discs 36 each carry an annular array of turbine blades 38.
  • the number of blade carrying discs could be different, i.e. only one disc or more than two discs.
  • nozzle guiding vanes 40 which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and turn the flow of working gas onto the turbine blades 38.
  • the combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22.
  • the guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.
  • the turbine section 18 drives the compressor section 14.
  • the compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48.
  • the rotor blade stages 48 comprise a rotor disc supporting an annular array of blades.
  • the compressor section 14 also comprises a casing 50 that surrounds the rotor stages and supports the vane stages 48.
  • the guide vane stages include an annular array of
  • the vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point.
  • Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions.
  • the casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14.
  • a radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor which is partly defined by the annular array of blades 48.
  • the present invention is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine. However, it should be appreciated that the present invention is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications.
  • the terms upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine unless otherwise stated.
  • the terms forward and rearward refer to the general flow of gas through the engine.
  • the terms axial, radial and circumferential are made with reference to the rotational axis 20 of the engine.
  • Inventive embodiments of a gas turbine aerofoil arrangement 100 are shown in Figs. 2 to 7.
  • the turbine aerofoil arrangement 100 may be used in the turbine guiding vanes 40.
  • the aerofoil arrangement 100 may also be used in turbine blades 48 and other aerofoil
  • aerofoil arrangement is intended to mean a component having an aerofoil 101, at least one platform 124 and at least one retention feature 102 such as a (fir-tree) dovetail root 102 or pinned fixture to which the aerofoil arrangement 100 is secured within the gas turbine engine.
  • a first preferred embodiment of a gas turbine aerofoil arrangement 100 is shown, applied to a nozzle guiding vane 40.
  • the exemplary embodiments refer to a nozzle guide vane having a single and radially inner platform on to which the aerofoil extends radially outwardly therefrom.
  • present arrangement may be also applied to an aerofoil attached to a radially outer platform. In this case the terms radially inward / outward may be
  • the guiding vane 40 (partially represented in the attached figures) comprises:
  • the inner platform 124 extends circumferentially to abut platforms of adjacent guide vanes and axially to seal with other engine components such as casings and rotor stages.
  • the platform 124 has an aperture 125 through which an inner insert 170 extends. At least a part of the inner insert 170, 171 is radially inward of the platform 124.
  • a panel 126 may be attached to the radially innermost part of the inner insert 170, 171.
  • the panel 126 and the platform 124 are generally parallel to each other.
  • the platform 124 is more distant from the rotational axis 20 than the panel 126.
  • the inner platform 124 has a radially outer or gas-washed surface 123 and a radially inner or cooled surface 121.
  • the nozzle guide vane may further have an outer platform (not represented) on a radial outer side of the stator 42.
  • the outer platform is more distant from the rotational axis 20 than the inner platform 124.
  • the cooled surface 121 of the radially outer platform is the radially outer surface of the outer platform.
  • the hollow aerofoil 101 extending along the radial direction between a first radial end 134 and a second radial end 136.
  • the first radial end 134 of the aerofoil 101 is a base of the aerofoil 101, connected to the inner platform 124.
  • the second radial end 136 of the aerofoil 101 is a tip of the aerofoil 101, connected to the outer platform.
  • the guiding vanes 40 are arranged in a guiding vane sectors or an annular array of vanes, each comprising an inner platform, an outer platform and a
  • the aerofoil 101 has an external surface 128.
  • the external surface 128 axially extends between a leading edge 133 and a trailing edge 135 and radially extends between a base 134 and a tip 136.
  • the aerofoil 101 is hollow and further comprises an internal cooling passage 150 for channelling the cooling medium from the tip 136 to the base 134 and for finally delivering the cooling medium to the inner platform 124.
  • the cooling medium enters the cooling internal passage 150 at the tip 136 along an input radial flow direction Fl .
  • Other aerofoils may have different cooling passages, such as serpentine passages, and which can be fed via an aperture positioned radially inward of the aerofoil.
  • the cooling passage and therefore the cooling medium can make a number of radially inward and outward passes before being fed to the insert 170, 171 or aperture 125.
  • the internal passage 150 is shaped in such a way that at the first radial end 134 the cooling medium exits the aerofoil 101 according to an output flow direction F2 having at least a component orthogonal to the radial direction.
  • the cooling medium exits the aerofoil 101 according to a flow direction F2 largely perpendicular to the radial axis and so parallel to and over the cooled surface
  • the cooling medium flows according to a final flow direction F3 parallel to the cooled surface 121of the inner platform 124.
  • the smooth transition from the input radial flow direction Fl to the final flow direction F3, through the output flow direction F2 permits to achieve the advantages of the present
  • the cooling medium is directed by the insert 170, 171 to flow over the cooled surface 121 to remove heat in the platform 124.
  • the term , over' is intended to mean that the cooling medium is in contact with or flows across the cooled surface 121. Essentially, the cooling fluid or medium is intended to immediately adjacent the cooled surface 121.
  • this is obtained by providing an inner insert 170 having a spiral surface 162 in the internal passage 150.
  • the spiral surface 162 is arranged around a stem 175 and is inclined with respect to the radial direction of the gas turbine engine 10.
  • the insert 170 may have a single spiral surface or wall 162 or as shown two spiral walls 162 and 162A and which form two spiral channels 166 and 166A having two entries 163, 163A and two exits 164, 164A. Shown the two spiral walls 162, 162A are equally spaced from one another and approximately equal amounts of coolant pass along the two channels 166, 166A.
  • one channel it is possible for one channel to have a greater cross-sectional than the other such that preferential amounts of coolant is channelled to one side of the platform than the other e.g. a greater amount of coolant is channelled to the suction side than the pressure side or vice versa.
  • preferential cooling will depend on the particular thermal characteristics of each application of the present
  • the exit 164 is arranged to direct the cooling medium to a preferred location, for example a hot spot, on the platform to preferentially cool that location. This is beneficial because cooling a hot spot reduces the temperature gradient across the platform and help prevent or minimise thermal gradients which can be
  • the insert 170 is fixed to the inner platform 124 and extends from one to the other of the base 134 and the tip 136.
  • the insert 170 may be fixed to other parts of the aerofoil arrangement. In other embodiments the insert 170 only extends part of the span of the aerofoil from the inner platform 124 towards the tip of the aerofoil 136 or other platform.
  • the inner insert 170 receives the cooling medium from the tip 136 and guides it along the spiral surface 162, 162A (as schematically represented by the arrows F4 of Figs. 4 and 5) to the base 134 and to the platform 124, where the cooling medium exits the aerofoil 101 according to the output flow direction F2.
  • the deflector 171 has two surfaces 173, 174 both
  • the two surfaces 173, 174 are inclined towards opposite sides of the aerofoil 101, e.g. towards the pressure side 137 and the suction side 138 of the aerofoil.
  • the deflector 171 has a triangular section in a sectional plane parallel to the radial direction, i.e. a sectional plane orthogonal to rotational axis 20.
  • the triangular section of the deflector 171 includes a base side 177 parallel to the base 134 of the aerofoil 101 and a vertex 178 oriented towards the tip 136.
  • the deflector 171 constitutes a flow deflector, used at the first radial end 134 of the aerofoil 101 for turning the flow of the cooling medium as it exits the aerofoil, from the radial direction Fl to the output direction F2.
  • He deflector may be trapezoidal prism (i.e. a truncated
  • triangular prism having a relatively small flat surface instead of an apex. Otherwise the trapezoidal prism operates and can have the same function as the triangular prism.
  • the triangular prism 171 is curved in the plane of the platform 124 and generally conforms to the curvature of the aerofoil 101.
  • the triangular prism 171 is generally positioned central to the aperture 125 formed in the platform 124 such that an approximately equal amount of cooling medium is directed towards the pressure and suction sides of the platform and over the cooled surface 121 on each side of the platform.
  • the heat incurred by the platform can be greater on one side of the platform than the other and therefore the triangular prism 171 may be offset to direct more cooling medium to the hotter side of the platform than the other side of the platform. For example, if the suction side 124S of the platform is hotter than the pressure side 124P, the
  • triangular prism 171 is located nearer the pressure side so that more cooling medium is directed over the cooled surface of the platform's suction side 124S. This is by virtue of a greater area of the opening 125 being on the suction side of the apex 178 of the triangular prism 171.
  • the base side 177 of the triangular prism 171 is attached to or integral to a lower panel or plate 126.
  • the lower panel 126 helps to direct the flow F2 of cooling medium over the cooled surface 121 of the platform 124.
  • the lower plate 126 extends part of the circumferential and/or axial extent of the platform, but may extend to overlap all the platform.
  • the first radial end of the aerofoil is a tip of the aerofoil, connected to an outer platform of the stator of the turbine.
  • the second radial end of the aerofoil is a base of the aerofoil, connected to an inner platform of the stator of the turbine.
  • the internal passage inside the aerofoil extends from the base to the tip for cooling the outer platform of the turbine stator.
  • the outer platform has a similar structure of the inner platform of the embodiments in Figs. 2 to 7, i.e.
  • Cooling of the cooled surface 121 may be further enhanced by the addition of well known surface features 141 (see dashed lines in Fig.3) such as ribs, fins, pins, pedestals or turbulators. These cooling features may be located on part or all of the cooled surface 121 both to reduce overall temperature and/or the temperature gradient of the platform.
  • the cooling features 14 may be located only on the pressure side or the suction side of the platform. In particular, there surface features may be placed where there is a hot- spot on the platform to enhance cooling in the region of the hot-spot.
  • the cooling features can function in two ways: firstly, to increase surface area exposed to the cooling medium and therefore increase heat transfer to the cooling medium; and secondly, to create turbulence in the cooling medium to maximise convection in the bulk cooling medium.
  • the cooling medium within the aerofoil generally passes in a radial direction and that the deflector 170, 171 directs the radial flow to a plane generally perpendicular to the radial direction. This generally perpendicular plane is also defined by the coled surface 121 of the platform.
  • the deflector 170, 171 can be arranged to direct cooling medium in the circumferential direction or the axial direction or in a combination of circumferential and axial directions as the particular application requires.

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne un profil aérodynamique (100) pour une turbine à gaz (10), comprenant : – une surface externe (128) s'étendant le long d'une première direction axiale entre un bord d'attaque (133) et un bord de fuite (135) et s'étendant le long d'une deuxième direction radiale entre une base (134) et une pointe (136) du profil aérodynamique (100) ; – un passage interne (150) pour canaliser radialement un fluide de refroidissement allant de la pointe (136) vers la base (134), le passage interne (150) étant formé de telle sorte qu'à la base (134), le fluide de refroidissement sorte du profil aérodynamique (100) suivant une direction d'écoulement (F2) ayant au moins une composante perpendiculaire à la direction radiale.
PCT/EP2017/050449 2016-01-19 2017-01-11 Agencement de profil aérodynamique WO2017125289A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP16151804.8 2016-01-19
EP16151804 2016-01-19

Publications (1)

Publication Number Publication Date
WO2017125289A1 true WO2017125289A1 (fr) 2017-07-27

Family

ID=55177813

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2017/050449 WO2017125289A1 (fr) 2016-01-19 2017-01-11 Agencement de profil aérodynamique

Country Status (1)

Country Link
WO (1) WO2017125289A1 (fr)

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0974733A2 (fr) * 1998-07-22 2000-01-26 General Electric Company Aubes de guidage pour une turbine ayant un système de transfert de l'air de refroidissement
EP1288442A1 (fr) * 2001-08-27 2003-03-05 General Electric Company Méthode et dispositif de contrôle du flux de refroidissement dans une aube de turbine
EP1452690A2 (fr) * 2003-02-27 2004-09-01 General Electric Company Refroidissement par convection forcée d'une tuyère de guidage pour turbines à gaz
EP1571296A1 (fr) * 2004-03-01 2005-09-07 Alstom Technology Ltd Aube refroidie de turbomachine et procédé de refroidissement

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0974733A2 (fr) * 1998-07-22 2000-01-26 General Electric Company Aubes de guidage pour une turbine ayant un système de transfert de l'air de refroidissement
EP1288442A1 (fr) * 2001-08-27 2003-03-05 General Electric Company Méthode et dispositif de contrôle du flux de refroidissement dans une aube de turbine
EP1452690A2 (fr) * 2003-02-27 2004-09-01 General Electric Company Refroidissement par convection forcée d'une tuyère de guidage pour turbines à gaz
EP1571296A1 (fr) * 2004-03-01 2005-09-07 Alstom Technology Ltd Aube refroidie de turbomachine et procédé de refroidissement

Similar Documents

Publication Publication Date Title
US11230935B2 (en) Stator component cooling
JP5503140B2 (ja) 発散型タービンノズル
US11655718B2 (en) Blade with tip rail, cooling
US6422819B1 (en) Cooled airfoil for gas turbine engine and method of making the same
CN106988789B (zh) 具有膜冷却的发动机构件
CN108868898B (zh) 用于冷却涡轮发动机的翼型件顶端的设备和方法
EP3485147B1 (fr) Refroidissement par impact d'une plate-forme d'aube
US20170211393A1 (en) Gas turbine aerofoil trailing edge
US10605170B2 (en) Engine component with film cooling
US10830057B2 (en) Airfoil with tip rail cooling
US11624286B2 (en) Insert for re-using impingement air in an airfoil, airfoil comprising an impingement insert, turbomachine component and a gas turbine having the same
CA2956979A1 (fr) Trous d'impacteur de composant de turbine
US11396818B2 (en) Triple-walled impingement insert for re-using impingement air in an airfoil, airfoil comprising the impingement insert, turbomachine component and a gas turbine having the same
CN107091122B (zh) 具有冷却的涡轮发动机翼型件
EP3460190A1 (fr) Structures d'amélioration de transfert de chaleur sur des nervures en ligne d'une cavité de surface portante d'une turbine à gaz
EP3205824A1 (fr) Insert d'accélérateur pour un profil aérodynamique de turbine à gaz
US20180156041A1 (en) Engine with chevron pin bank
GB2054046A (en) Cooling turbine rotors
EP3252270B1 (fr) Système de refroidissement de buse pour moteur à turbine à gaz
US20200325780A1 (en) A turbomachine blade or vane having a vortex generating element
WO2017125289A1 (fr) Agencement de profil aérodynamique
EP3241991A1 (fr) Ensemble de turbine
JP2021156284A (ja) ターボ機械構成要素用の冷却回路
EP3279432A1 (fr) Aube avec un ou plusieurs socles ayant une surface alvéolée destinée à refroidir
CA3202558A1 (fr) Refroidissement accru pour l'optimisation du degagement des pointes

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 17700641

Country of ref document: EP

Kind code of ref document: A1

NENP Non-entry into the national phase

Ref country code: DE

122 Ep: pct application non-entry in european phase

Ref document number: 17700641

Country of ref document: EP

Kind code of ref document: A1