WO2013169471A1 - Fente de refroidissement de bord de fuite d'ailette de turbine - Google Patents

Fente de refroidissement de bord de fuite d'ailette de turbine Download PDF

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Publication number
WO2013169471A1
WO2013169471A1 PCT/US2013/037510 US2013037510W WO2013169471A1 WO 2013169471 A1 WO2013169471 A1 WO 2013169471A1 US 2013037510 W US2013037510 W US 2013037510W WO 2013169471 A1 WO2013169471 A1 WO 2013169471A1
Authority
WO
WIPO (PCT)
Prior art keywords
airfoil
spanwise
trailing edge
deck
section
Prior art date
Application number
PCT/US2013/037510
Other languages
English (en)
Inventor
Robert Frederick Bergholz Jr.
Daniel Lee Durstock
Original Assignee
General Electric Company
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 General Electric Company filed Critical General Electric Company
Priority to CA2872265A priority Critical patent/CA2872265A1/fr
Priority to JP2015511494A priority patent/JP2015516539A/ja
Priority to CN201380024101.6A priority patent/CN104285038A/zh
Priority to EP13720668.6A priority patent/EP2855850A1/fr
Priority to BR112014026785A priority patent/BR112014026785A2/pt
Publication of WO2013169471A1 publication Critical patent/WO2013169471A1/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
    • F01D5/186—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
    • F05D2240/20—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
    • F05D2240/301—Cross-sectional characteristics
    • 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 generally to gas turbine engine turbine airfoil cooling and, more
  • turbine airfoil trailing edge cooling slots specifically, to turbine airfoil trailing edge cooling slots .
  • the turbine stages include stationary turbine nozzles having a row of hollow vanes which channel the combustion gases into a corresponding row of rotor blades extending radially outwardly from a supporting rotor disk.
  • the vanes and blades have corresponding hollow airfoils with corresponding cooling circuits therein.
  • the cooling air is typically compressor
  • the airfoil span extends from a root at the radially inner platform to a radially outer tip spaced from a surrounding turbine shroud.
  • the airfoil extends from a root integral with a radially inner band to a radially outer tip integral with an outer band.
  • Each turbine airfoil also initially increases in thickness aft of the leading edge and then decreases in thickness to a relatively thin or sharp trailing edge where the pressure and suction sidewalls join together.
  • the wider portion of the airfoil has sufficient internal space for accommodating various forms of internal cooling circuits and turbulators for enhancing heat transfer cooling inside the airfoil, whereas, the relatively thin trailing edge has correspondingly limited internal cooling space.
  • Each airfoil typically includes various rows of film cooling holes extending through the sidewalls thereof which discharge the spent cooling air from the internal circuits.
  • the film cooling holes are typically inclined in the aft direction toward the trailing edge and create a thin film of cooling air over the external surface of the airfoil that provides a thermally
  • the thin trailing edge is typically protected by a row of trailing edge cooling slots or a single elongated slot which breach the pressure sidewall at a breakout immediately upstream of the trailing edge for discharging film cooling air thereover.
  • Each trailing edge cooling slot has an outlet aperture in the pressure side which begins at a breakout and may or may not be bounded in the radial direction by exposed lands at aft ends of axially extending partitions which define the cooling slots.
  • the axial partitions may be integrally formed with the pressure and suction sides of the airfoil and themselves must be cooled by the air discharged through the cooling slots defined thereby.
  • the partitions typically converge in the aft direction toward the trailing edge so that the cooling slots diverge toward the trailing edge with a shallow divergence angle that promotes diffusion of the discharged cooling air with little, if any, flow separation along the sides of the partitions .
  • Aerodynamic and cooling performance of the trailing edge cooling slots is directly related to the specific configuration of the cooling slots and the intervening partitions.
  • the flow area of the cooling slots regulates the flow of cooling air discharged through the cooling slots and the geometry of the cooling slots affects cooling performance thereof.
  • the lands are solid portions of the pressure sidewall integrally formed with the suction sidewall and must rely for cooling on the air discharged from the adjacent trailing edge cooling slots.
  • a gas turbine engine turbine airfoil includes widthwise spaced apart pressure and suction sidewalls extending outwardly along a span from an airfoil base to an airfoil tip and extending in a chordwise direction between opposite leading and trailing edges.
  • a spanwise row of spanwise spaced apart trailing edge cooling holes encased in the airfoil between the pressure and suction sidewalls end at a single spanwise extending trailing edge cooling slot extending chordally substantially to the trailing edge.
  • Each of the cooling holes includes in downstream serial cooling flow relationship, a curved inlet, a metering section with a constant area and constant width flow cross section, and a spanwise
  • Axial partitions extend chordally between and radially separate the cooling holes along the span. Aft ends of the partitions include swept boat tails.
  • the boat tails may be swept with each of the boat tails including a boat tail trailing edge having an apex spanwise located between the pressure and suction sidewalls.
  • the boat tail trailing edge sweeps aftwardly or downstream from the apex.
  • the boat tail trailing edge sweeps from the apex spanwise or radially outwardly to the pressure sidewall and inwardly to the suction
  • the swept boat tails may further include rounded cross sections through the aft ends of the partitions between spanwise pairs of adjacent cooling holes .
  • the pressure and suction sidewalls may include pressure and suction sidewall surfaces respectively in the hole and the pressure sidewall surface may be planar through the entire metering and diverging sections.
  • the width may be constant through the metering and diverging sections of the hole.
  • the airfoil may include a deck in the slot extending chordwise or downstream from the diverging sections of the cooling holes substantially to the airfoil trailing edge and extending spanwise or radially outwardly from a bottommost one to a topmost one of the trailing edge cooling holes.
  • Upper and lower deck sidewalls spanwise bound the deck and extend from the deck to an external surface of the pressure sidewall. Fillets in slot corners between the upper and lower deck sidewalls and the deck have fillet radii substantially the same size as bottom corner radii of the flow cross section of the diverging sections adjacent the bottom corner radii.
  • FIG. 1 is a longitudinal, sectional view illustration of an exemplary embodiment of turbine vane and rotor blade airfoils having cooling holes culminating at a spanwise extending trailing edge cooling slot.
  • FIG. 2 is an enlarged view illustration of a blade illustrated in FIG. 1.
  • FIG. 3 is a pressure side sectional view illustration of the cooling holes with constant width metering and diffusing sections leading into the trailing edge cooling slot illustrated in FIG. 2.
  • FIG. 4 is a cross sectional schematical view illustration of the cooling holes with constant width metering and diffusing sections leading into the trailing edge cooling slot taken through 4-4 in FIG. 3.
  • FIG. 5 is an upstream looking perspective view illustration of the cooling holes and the trailing edge cooling slot illustrated in FIG. 3.
  • FIG. 6 is an enlarged perspective view
  • FIG. 7 is a cross sectional schematical view illustration of an elongated flow cross section in the constant width metering section taken through 7-7 in FIG.
  • FIG. 8 is a cross sectional schematical view illustration of an elongated flow cross section in the diffusing section taken through 8-8 in FIG. 3.
  • FIG. 9 is a cross sectional schematical view of a race track shaped flow cross section having four equal radii .
  • FIG. 10 is a cross sectional schematical view of an alternative race track shaped flow cross section with a larger width to height ratio than the race track shaped flow cross section illustrated in FIG. 9.
  • FIG. 11 is a cross sectional schematical view of an alternative flow cross section with unequal top and bottom corner radii.
  • FIG. 12 is a cross sectional schematical view of another alternative flow cross section with in
  • FIG. 1 Illustrated in FIG. 1 is an exemplary gas turbine engine high pressure turbine stage 10
  • the combustor 20 mixes fuel with pressurized air for generating hot combustion gases 19 which flows downstream D through the turbines.
  • the high pressure turbine stage 10 includes a turbine nozzle 28 upstream of a high pressure turbine (HPT) 22 through which the hot combustion gases 19 are discharged into from the combustor 20.
  • HPT high pressure turbine
  • the exemplary embodiment of the high pressure turbine 22 illustrated herein includes at least one row of circumferentially spaced apart high pressure turbine blades 32.
  • Each of the turbine blades 32 includes a turbine airfoil 12 integrally formed with a platform 14 and an axial entry dovetail 16 used to mount the turbine blade on a
  • the airfoil 12 extends radially outwardly along a span S from an airfoil base 34 on the blade platform 14 to an airfoil tip 36.
  • the hot combustion gases 19 are generated in the engine and flow downstream D over the turbine airfoil 12 which extracts energy therefrom for rotating the disk supporting the blade for powering the compressor (not shown) .
  • a portion of pressurized air 18 is suitably cooled and directed to the blade for cooling thereof during operation.
  • the airfoil 12 includes widthwise spaced apart generally concave pressure and convex suction sidewalls 42, 44.
  • the pressure and suction sidewalls 42, 44 extend longitudinally or radially outwardly along the span S from the airfoil base 34 to the airfoil tip 36.
  • the sidewalls also extend axially in a chordwise direction C between opposite airfoil leading and trailing edges LE,
  • the airfoil 12 is hollow with the pressure and suction sidewalls 42, 44 being spaced widthwise or laterally apart between the airfoil leading and trailing edges LE, TE to define an internal cooling cavity or circuit 54 therein for circulating pressurized cooling air or coolant flow 52 during operation.
  • the pressurized cooling air or coolant flow 52 is from the portion of pressurized air 18 diverted from the compressor.
  • the turbine airfoil 12 increases in width W or widthwise from the airfoil leading edge LE to a maximum width aft therefrom and then converges to a relatively thin or sharp airfoil trailing edge TE .
  • the size of the internal cooling circuit 54 therefore varies with the width W of the airfoil, and is relatively thin
  • a spanwise extending trailing edge cooling slot 66 is provided at or near this thin trailing edge portion 56 of the airfoil 12 to cool it.
  • the trailing edge cooling slot 66 extends chordally substantially to the trailing edge TE .
  • the trailing edge cooling holes 30 are disposed along the span S of the trailing edge TE in flow
  • a floor or deck 130 of the slot 66 extends chordwise or downstream from the diverging sections 102 of the cooling holes 30 substantially to the airfoil trailing edge TE .
  • the deck 130 extends spanwise or radially outwardly from a bottommost one 132 to a topmost one 134 of the trailing edge cooling holes 30.
  • the deck 130 is spanwise bound by upper and lower deck sidewalls 136, 138 which extend from the deck 130 to an external surface 43 of the pressure sidewall 42.
  • a slot surface 60 extends widthwise between the upper and lower deck sidewalls 136, 138 along the deck 130. Fillets 62 in slot corners 64 between the upper and lower deck
  • sidewalls 136, 138 and the deck 130 have fillet radii RF that may be substantially the same size as bottom corner radii RT of the flow cross section 74 of the diverging sections 102 adjacent the bottom corner radii RT
  • the fillet radii RF helps with castability of the trailing edge cooling slot 66.
  • Each cooling hole 30 includes, in downstream serial cooling flow relationship, a downstream converging or bellmouth shaped curved inlet 70, a constant area and constant width flow cross section metering section 100, and a spanwise diverging section 102 which leads into the trailing edge cooling slot 66 and supplies the slot with cooling air or coolant flow 52.
  • the trailing edge cooling slot 66 begins at a breakout 58 located at downstream ends 69 of the diverging sections 102. The diverging sections 102 of the cooling holes 30 lead into the trailing edge cooling slot 66 which
  • the cooling holes 30 are separated radially along the span S from each other by corresponding axial partitions 68 which extend downstream toward the trailing edge TE .
  • the curved inlet 70 is illustrated herein as downstream converging or, more particularly, a bellmouth inlet.
  • the inlet 70 is defined at and between forward ends 72 of the partitions 68.
  • the partitions 68 include semi -circular forward ends 72 having diameters 73 that define the bellmouth inlet 70.
  • Each of the cooling holes 30 includes spanwise spaced apart upper and lower hole surfaces 46, 48 along a corresponding adjacent pair of upper and lower ones 25, 26 of the axial partitions 68.
  • a spanwise height H of the hole 30 is defined between the upper and lower hole surfaces 46, 48 of the upper and lower ones 25, 26 of the axial partitions 68 as illustrated in FIG. 3.
  • the metering section 100, the diverging section 102, and the trailing edge cooling slot 66 have downstream extending first, second, and third lengths LI, L2 , and L3
  • aft ends 86 of the partitions 68 have aerodynamically- shaped swept boat tails 88 design and shaped to reduce aerodynamic losses due to flow separation wakes at the aft ends 86.
  • the swept boat tails 88 are also designed to facilitate flow spreading past the slot breakout 58 at the downstream end
  • Each of the swept boat tails 88 include a boat tail trailing edge 90 extending spanwise between the pressure and suction sidewalls 42, 44 and having an apex 92 spanwise located between the breakout lip 49 or the pressure sidewall 42 and the suction sidewall 44 as illustrated in FIG. 6.
  • the boat tail trailing edge 90 sweeps aftwardly or downstream from the apex 92.
  • the boat tail trailing edge 90 sweeps from the apex 92 spanwise or radially outwardly to the
  • Each of the swept boat tails 88 includes rounded cross sections 96 through the aft ends 86 of the partitions 68 between spanwise pairs 94 of adjacent cooling holes 30.
  • the boat tail trailing edges 90 provide additional strength at the breakout lip 49 and merges the flows of the different cooling holes 30 at the floor or deck 130 upstream of the breakout 58 which is an exit of the cooling holes 30.
  • the cooling holes 30, the trailing edge cooling slot 66, and the swept boat tails 88 are designed to provide a spanwise deck 130 film effectiveness over the entire slot deck 130 all the way downstream or aft to the terminus of the deck 130 the airfoil trailing edge TE, even at significantly reduced cooling flow.
  • Airfoil cooling design studies have shown a potential cooling flow reduction of about 10 percent of stage 1 blade flow.
  • trailing edge temperatures are still 80 to 90 degrees F lower that more conventional slot designs, so further flow reductions are possible. This is a significant benefit to engine performance .
  • a hole width W of the hole 30 is defined between pressure and suction sidewall surfaces 39, 40 of the pressure and suction sidewalls 42, 44 respectively as illustrated in FIG. 4.
  • the trailing edge cooling slot 66 and the deck 130 are open and exposed to the hot combustion gases 19 that pass through the high pressure turbine 22.
  • the deck 130 extends for the entire third length L3 along the suction sidewall 44.
  • the cooling hole 30 has a generally spanwise elongated flow cross section 74 and the spanwise height H is substantially greater than the hole width W.
  • the cooling hole 30 has a height to width ratio H/W in a range of about 2:1 to 10:1 (see FIGS. 7- 12) .
  • the pressure and suction sidewall surfaces 39, 40 of the pressure and suction sidewalls 42, 44 respectively widthwise bound the hole 30.
  • FIG. 4 has a fixed or constant width W through the cooling hole 30 and the pressure and suction sidewall surfaces 39, 40 are parallel through the entire first and second lengths LI, L2 of the cooling hole 30.
  • the pressure sidewall surface 39 is flat or planar through the entire metering and diverging sections 100, 102 and their corresponding first and second lengths LI, L2 of the cooling hole 30.
  • the suction sidewall surface 40 is flat or planar through the entire metering and diverging sections 100, 102 and their corresponding first and second lengths LI, L2 of the cooling hole 30.
  • the deck 130 is coplanar with suction sidewall surface 40 in the hole 30.
  • the inlet 70, the metering section 100, and the diverging section 102 have the same hole width W or are of constant width W in the embodiment of the trailing edge cooling holes 30 illustrated in FIG. 3 and
  • the diverging section 102 diverges in a spanwise direction.
  • the cooling holes 30 and trailing edge cooling slot 66 are cast in cooling features. Casting these features provides good strength, low manufacturing costs, and durability for the airfoil and blades and vanes.
  • the bottom corner radii RT contribute to good cooling, castability, and strength of these cooling features.
  • FIGS. 9-12 Four exemplary shapes suitable for the flow cross section 74 are illustrated in FIGS. 9-12.
  • the race track shaped flow cross section 74 illustrated in FIG. 9 is spanwise elongated, has four equal corner radii R, and has a width to height ratio W/H in a range of 0.25-0.50.
  • the race track shaped flow cross section 74 illustrated in FIG. 10 is spanwise elongated, has four equal corner radii R, and has a width to height ratio W/H in a range of 0.15-0.50.
  • the race track shaped flow cross section 74 illustrated in FIG. 11 is similar to the one
  • the race track shaped flow cross section 74 illustrated in FIG. 12 is spanwise elongated and fully curved and includes curved quarter sides 78 that may be elliptical, parabolic, or polynomial blends.
  • the spanwise elongated metering section 100 with the constant width W is sized to control the
  • diverging section 102 expand the flow coverage at the breakout 58, evenly distributes coolant flow 52 in the trailing edge cooling slot 66.
  • the constant width W metering section 100 upstream of the diverging section 102 of the hole 30 helps keep the coolant flow 52 fully attached in the diverging section 102.
  • the constant width and separately the planar pressure sidewall surface 39 of the cooling hole 30 helps keep a coolant velocity of the coolant flow 52 and a gas velocity of the hot combustion gases along the external surface 43 of the pressure sidewall 42 at the breakout about equal to minimize aero losses which could result in a negative effect on turbine efficiency. These two features also help keep the coolant flow 52 flow attached in the slot 66.

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

Abstract

L'invention porte sur une ailette de turbine qui présente des flancs de pression et d'aspiration qui s'étendent selon une longueur, d'une base à une pointe. Des trous de refroidissement de bord de fuite espacés dans le sens de la longueur, dans le flanc de pression, se terminent au niveau d'une fente de refroidissement de bord de fuite s'étendant dans la direction de la corde, sensiblement jusqu'à un bord de fuite d'ailette. Chaque trou de refroidissement comprend une entrée incurvée, une section doseuse ayant une partie d'écoulement à une aire constante et à largeur constante et une section divergente sur la longueur qui débouche dans une fente. Des cloisons axiales s'étendent dans la direction de la corde entre les trous de refroidissement espacés radialement sur la longueur. Les extrémités arrière des cloisons présentent des parties rétrécies infléchies. Des flancs de ponts supérieur et inférieur atteignent, dans le sens de la longueur, un pont d'une fente et s'étendent vers l'extérieur jusqu'à une surface extérieure du flanc de pression. Des congés entre les flancs et le pont ont des rayons de congé sensiblement de même dimension que les rayons d'angle inférieur de la section de passage de flux de sections divergentes adjacents aux rayons d'angle inférieur.
PCT/US2013/037510 2012-05-08 2013-04-22 Fente de refroidissement de bord de fuite d'ailette de turbine WO2013169471A1 (fr)

Priority Applications (5)

Application Number Priority Date Filing Date Title
CA2872265A CA2872265A1 (fr) 2012-05-08 2013-04-22 Fente de refroidissement de bord de fuite d'ailette de turbine
JP2015511494A JP2015516539A (ja) 2012-05-08 2013-04-22 タービン翼形部の後縁冷却スロット
CN201380024101.6A CN104285038A (zh) 2012-05-08 2013-04-22 涡轮翼型件后缘冷却槽口
EP13720668.6A EP2855850A1 (fr) 2012-05-08 2013-04-22 Fente de refroidissement de bord de fuite d'ailette de turbine
BR112014026785A BR112014026785A2 (pt) 2012-05-08 2013-04-22 aerofólio de turbina.

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US13/466,487 US20130302176A1 (en) 2012-05-08 2012-05-08 Turbine airfoil trailing edge cooling slot
US13/466,487 2012-05-08

Publications (1)

Publication Number Publication Date
WO2013169471A1 true WO2013169471A1 (fr) 2013-11-14

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ID=48289651

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PCT/US2013/037510 WO2013169471A1 (fr) 2012-05-08 2013-04-22 Fente de refroidissement de bord de fuite d'ailette de turbine

Country Status (7)

Country Link
US (1) US20130302176A1 (fr)
EP (1) EP2855850A1 (fr)
JP (1) JP2015516539A (fr)
CN (1) CN104285038A (fr)
BR (1) BR112014026785A2 (fr)
CA (1) CA2872265A1 (fr)
WO (1) WO2013169471A1 (fr)

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WO2021186122A1 (fr) * 2020-03-18 2021-09-23 Safran Aircraft Engines Aube de turbine comportant des nervures entre des sorties de refroidissement avec des orifices de refroidissement

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US9732617B2 (en) 2013-11-26 2017-08-15 General Electric Company Cooled airfoil trailing edge and method of cooling the airfoil trailing edge
JP6671149B2 (ja) 2015-11-05 2020-03-25 三菱日立パワーシステムズ株式会社 タービン翼及びガスタービン、タービン翼の中間加工品、タービン翼の製造方法
US10301954B2 (en) 2016-01-08 2019-05-28 General Electric Company Turbine airfoil trailing edge cooling passage
US10260354B2 (en) 2016-02-12 2019-04-16 General Electric Company Airfoil trailing edge cooling
US10563518B2 (en) 2016-02-15 2020-02-18 General Electric Company Gas turbine engine trailing edge ejection holes
KR101937589B1 (ko) * 2017-09-18 2019-04-09 두산중공업 주식회사 터빈의 터빈 블레이드와 터빈 베인 및 이를 포함하는 터빈 및 가스터빈
US10787932B2 (en) * 2018-07-13 2020-09-29 Honeywell International Inc. Turbine blade with dust tolerant cooling system
FR3094033B1 (fr) * 2019-03-22 2021-06-11 Safran Aircraft Engines Aube de turbomachine equipee d’un circuit de refroidissement optimise
US11608746B2 (en) * 2021-01-13 2023-03-21 General Electric Company Airfoils for gas turbine engines
EP4105441A1 (fr) * 2021-06-18 2022-12-21 ITP Next Generation Turbines S.L. Profil aérodynamique pour un moteur à turbine

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GB2036884A (en) * 1978-12-15 1980-07-02 Gen Electric Film cooled airfoil
EP1213442A1 (fr) * 2000-12-05 2002-06-12 United Technologies Corporation Structure d'aube refroidissable
EP1326006A2 (fr) * 2002-01-04 2003-07-09 General Electric Company Procédé et dispositif pour le refroidissement des aubes de guidage d'une turbine à gaz
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Publication number Priority date Publication date Assignee Title
WO2021186122A1 (fr) * 2020-03-18 2021-09-23 Safran Aircraft Engines Aube de turbine comportant des nervures entre des sorties de refroidissement avec des orifices de refroidissement
FR3108364A1 (fr) * 2020-03-18 2021-09-24 Safran Aircraft Engines Aube de turbine comportant des nervures entre des sorties de refroidissement avec des orifices de refroidissement
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Also Published As

Publication number Publication date
CA2872265A1 (fr) 2013-11-14
EP2855850A1 (fr) 2015-04-08
JP2015516539A (ja) 2015-06-11
US20130302176A1 (en) 2013-11-14
CN104285038A (zh) 2015-01-14
BR112014026785A2 (pt) 2017-06-27

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