EP1930544A1 - Aube de turbine - Google Patents

Aube de turbine Download PDF

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Publication number
EP1930544A1
EP1930544A1 EP06022622A EP06022622A EP1930544A1 EP 1930544 A1 EP1930544 A1 EP 1930544A1 EP 06022622 A EP06022622 A EP 06022622A EP 06022622 A EP06022622 A EP 06022622A EP 1930544 A1 EP1930544 A1 EP 1930544A1
Authority
EP
European Patent Office
Prior art keywords
blade
heat protection
protection element
turbine
cooling
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Withdrawn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
EP06022622A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
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 EP06022622A priority Critical patent/EP1930544A1/fr
Priority to PCT/EP2007/059989 priority patent/WO2008052846A1/fr
Priority to CNA2007800407001A priority patent/CN101535601A/zh
Priority to JP2009533769A priority patent/JP2010508461A/ja
Priority to AT07820422T priority patent/ATE459784T1/de
Priority to EP07820422A priority patent/EP2084368B1/fr
Priority to RU2009120554/06A priority patent/RU2405940C1/ru
Priority to US12/447,972 priority patent/US20100068069A1/en
Priority to DE502007003043T priority patent/DE502007003043D1/de
Priority to PL07820422T priority patent/PL2084368T3/pl
Priority to ES07820422T priority patent/ES2341384T3/es
Publication of EP1930544A1 publication Critical patent/EP1930544A1/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Withdrawn legal-status Critical Current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
    • F01D5/186—Film cooling

Definitions

  • the invention relates to a turbine blade, in particular for use in a gas turbine, with a blade body. It further relates to a turbomachine provided with a number of such turbine blades.
  • Turbomachines in particular gas turbines, are used in many areas to drive generators or work machines.
  • a gas turbine usually has a rotatably mounted rotor surrounded by a stationary housing.
  • the fixed assemblies of a gas turbine are collectively referred to as a stator.
  • the energy content of a fuel is used to generate a rotational movement of the rotor component.
  • the fuel is burned in a combustion chamber, compressed air being supplied by an air compressor.
  • the working medium produced in the combustion chamber by the combustion of the fuel, under high pressure and at high temperature, is guided via a turbine unit arranged downstream of the combustion chamber, where it relaxes to perform work.
  • a number of rotor blades which are usually combined into blade groups or rows of blades, are arranged thereon and drive the rotor components by impulse transmission of the flow medium.
  • To guide the flow medium in the turbine unit also commonly associated guide blade rows are arranged between adjacent blade rows with the turbine housing.
  • the turbine blades, in particular the guide vanes usually have an airfoil extending along a blade axis for suitable guidance of the working medium, to the end for attachment of the turbine blade to the respective one Carrier body can be integrally formed transversely to the blade axis extending platform.
  • the exposed to these high temperatures of the working fluid components and components of the gas turbine are thus subject to high thermal stress.
  • the affected components in particular the rotor blades and / or guide vanes of the turbine unit, are cooled.
  • the turbine blades are usually provided with cooling channels, in particular, an effective and reliable cooling of the particular thermally loaded front edge of the respective turbine blade should be ensured.
  • Coolant is usually used as coolant. This can be supplied to the respective turbine blade via a number of integrated into the blade or the blade profile coolant channels. Starting from these, the cooling air flows through the respectively provided regions of the turbine blades in exit channels branching off from them, as a result of which convective cooling of the blade interior and of the blade wall is achieved. On the exit side, these channels are left open, so that the cooling air, after flowing through the turbine blade, flows out of the outlet openings, also referred to as film cooling holes, and forms a cooling film on the surface of the airfoil. Through this cooling air film of the blade body is at the Surface largely protected from direct and intensive contact with the hot working medium flowing past at high speed.
  • the outlet openings in this region are usually arranged uniformly along at least two rows aligned parallel to the front edge.
  • the exit channels are also generally oriented obliquely to the longitudinal direction of the turbine blade, which supports the formation of the protective, flowing on the surface cooling air film.
  • the front edge of the blade can also be provided with a heat protection layer.
  • This heat protection layer is expediently made of a more temperature-resistant material than that of the blade main body.
  • the heat protection layer is characterized by a low coefficient of thermal conductivity, whereby the temperature load of the base material of the blade body is reduced.
  • the life of the turbine blade is increased by such a heat protection layer in conjunction with cooling of the leading edge region of the blade.
  • this heat protection has the disadvantage that cracks in the heat protection layer occur after a certain time. These cracks reduce the protection of the blade body from the hot exhaust gas of the gas turbine, so that it may also cause cracking due to the increased thermal load in the main body of the turbine blade. Such cracks in the blade body endanger the reliability and can lead to failure of the gas turbine.
  • the invention is therefore an object of the invention to provide a turbine blade of the type mentioned above, which with simple Means guaranteed even when used in high flow temperatures, a particularly high reliability of the gas turbine.
  • This object is achieved according to the invention by providing the outer surface of the blade main body in the leading edge region with a heat protection element spaced therefrom.
  • the invention is based on the consideration that, especially with regard to the reliability and economy of a gas turbine, the turbine blades should have as long as possible by a suitably selected heat protection.
  • the fact should be taken into account that just the leading edge of the turbine blade is thermally heavily loaded. This should therefore be specially protected
  • the heat protection element is mounted at a distance to the blade body in the leading edge region, whereby a direct contact of the heat protection element is avoided with the blade body.
  • the blade base body For cooling the blade base body while the outer surface is provided in the leading edge region with a number of secondary cooling channels, which extend from the main cooling channel to the blade main body outer surface.
  • These auxiliary cooling channels are preferably distributed homogeneously behind the heat protection element for effective cooling in the leading edge region of the blade main body.
  • this has a number of its outer surface extending in the direction of the blade main body outlet channels.
  • This trained to guide a cooling flow channel is preferably also used as a connecting element between the heat protection element and the blade body.
  • the outlet channel protrudes at one end into a blade inside trained Main cooling channel into it, wherein the medium flowing in the main cooling channel medium can flow to cool the heat protection element on the outer surface.
  • the particularly critical region namely the leading edge region of the turbine blade
  • the particularly critical region is particularly effectively protected against the high temperatures of the working medium of the turbine.
  • the cooling is preferably carried out in such a way that a cooling flow from the main cooling channel is partly passed through the outlet channels of the heat protection element on the outer surface and flows partly from the main cooling channel, via the auxiliary cooling channels of the blade body, through the space formed by the heat protection element and the blade body.
  • a protective film is formed on the outer surface of the heat protection element.
  • This cooling film prevents direct contact of the hot working medium of the turbine with the heat protection element, whereby the temperature load of the impinged outer surface is reduced.
  • the still occurring increase in the temperature of the heat protection element does not affect directly on the temperature of the blade body in the leading edge region, since the heat protection element is arranged at a distance from the blade body.
  • the heat transfer between the heat protection element and the blade body is also substantially reduced by the flowing between the inner surface of the heat protection member and the blade body outer surface cooling medium by the heat in the leading edge region is removed by the inner cooling flow.
  • the heat protection element has a shape adapted to the profile of the blade body in the leading edge region. This will ensure that the turbine blade even after the attachment of the heat protection element has a flow-optimized shape in the leading edge region.
  • the form of the heat protection element corresponding to the blade body leads to a uniform expansion of the gap in the leading edge region.
  • the cooling medium flows at a predominantly constant speed along the blade body outer surface and the inner surface of the heat protection element, whereby the cooling in the leading edge region of the turbine blade is particularly uniform.
  • the heat protection element is made of a material which is more temperature-resistant than the blade body. Since the heat protection element is directly flowed during operation of the turbine from the hot working fluid, just this component is exposed to a high temperature load. Therefore, the heat protection element should be made of a particularly temperature-resistant material to ensure above all the reliability and minimize the downtime of the turbine.
  • the heat protection element is designed for an impingement cooling for a particularly effective cooling, this is achieved by keeping the distance of the heat protection element to the blade main body sufficiently low.
  • the heat protection element is preferably arranged at a distance of 1 mm to 3 mm from the main blade body. Just a heat protection element attached to this distance in the leading edge region of the turbine blade ensures a sufficiently high impact velocity of the cooling medium on the inner surface of the heat protection element, whereby a particularly effective cooling is achieved by impingement cooling.
  • the impact velocity the cooling flow in addition to, for example, the diameter of the secondary cooling channels determined primarily by the distance of the heat protection element to the blade body.
  • a sufficiently high velocity of the cooling medium immediately before impinging on the inner surface of the heat protection element is necessary because such an intimate contact between the cooling medium and the inner surface of the heat protection element comes about.
  • the auxiliary cooling channels are arranged aligned substantially perpendicular to the inner surface of the heat protection element.
  • the cooling flow from the main cooling channel is perpendicular to the inner surface of the heat protection element, whereby a large part of the kinetic energy of the cooling medium is used for a particularly intimate contact between the particles in the cooling flow and the inner surface of the heat protection element.
  • the heat of the heat protection element is transferred to the inner cooling stream in a particularly effective manner and transported away.
  • the heat protection element in the edge regions of the turbine blade is connected to the blade main body.
  • the heat protection element in the edge regions of the turbine blade is connected to the blade main body.
  • the described heat protection element can advantageously be used at the points of the turbomachine where components and assemblies of the thermal turbomachine are charged with the hot working medium.
  • the use of the heat protection element for protecting the leading edge region of the turbine blade is particularly preferred, since the temperature load of the blade main body in this area is particularly high.
  • the downtime of the gas turbine is minimized by such a heat protection system, since the life is increased by the heat protection element.
  • the advantages achieved by the invention are, in particular, that an effective protection of the leading edge region of the turbine blade from the high temperatures of the working medium of the turbines is provided by the heat shield element upstream of the blade body of the turbine blade.
  • a heat protection system allows the use of an impingement cooling, whereby the heat protection element can be cooled particularly effectively.
  • it is prevented by the heat protection element that any cracks, which originate from the outer surface of the heat protection element, propagate into the blade main body.
  • the heat protection elements according to the invention can be retrofitted to the turbine blades in a simple manner and with relatively little effort.
  • the gas turbine 1 has a compressor 2 for combustion air, a combustion chamber 4 and a turbine unit 6 for driving the compressor 2 and a generator, not shown, or a working machine.
  • the turbine unit 6 and the compressor 2 are arranged on a common, also called turbine rotor turbine shaft 8, with which the generator or the working machine is connected, and which is rotatably mounted about its central axis 9.
  • the running in the manner of an annular combustion chamber 4 is equipped with a number of burners 10 for the combustion of a liquid or gaseous fuel.
  • the turbine unit 6 has a number of rotatable blades 12 connected to the turbine shaft 8.
  • the blades 12 are arranged in a ring on the turbine shaft 8 and thus form a number of blade rows.
  • the turbine unit 6 comprises a number of fixed vanes 14, which are also annular in shape to form vanes on an inner casing 16 the turbine unit 6 are attached.
  • the blades 12 serve to drive the turbine shaft 8 by momentum transfer from the turbine unit 6 flowing through the working medium M.
  • the vanes 14, however, serve to guide the flow of the working medium M between two seen in the flow direction of the working medium M consecutive blade rows or blade rings.
  • a successive pair of a ring of vanes 14 or a row of vanes and a ring of blades 12 or a blade row is also referred to as a turbine stage.
  • Each vane 14 has a platform 18, which is arranged to fix the respective vane on the inner housing 16 of the turbine unit 6 as a wall element.
  • the platform 18 is - as well as the turbine blade 12, 14 - while a comparatively thermally heavily loaded component.
  • Each blade 12 is attached to the turbine shaft 8 in an analogous manner via a platform 19, also referred to as a blade root.
  • each guide ring 21 on the inner housing 16 of the turbine unit 6 is arranged between the spaced-apart platforms 18 of the guide vanes 14 of two adjacent rows of guide vanes.
  • the outer surface of each guide ring 21 is also exposed to the hot, the turbine unit 6 flowing through the working medium M and spaced in the radial direction from the outer end of the opposite blades 12 by a gap.
  • the guide rings 21 arranged between adjacent rows of guide blades serve, in particular, as cover elements which protect the inner housing 16 or other housing mounting parts against overstressing by the working medium M flowing through the turbine 6.
  • the combustion chamber 4 is configured in the exemplary embodiment as a so-called annular combustion chamber, in which a plurality of arranged circumferentially around the turbine shaft 8 around Burners 10 opens into a common combustion chamber.
  • the combustion chamber 4 is configured in its entirety as an annular structure which is positioned around the turbine shaft 8 around.
  • the combustion chamber 4 is designed for a comparatively high temperature of the working medium M of about 1000 ° C to 1600 ° C.
  • the rotor blades 12 -as shown in FIG. 2 - have a heat protection element 22 mounted in the leading edge region.
  • Each of the heat shield elements 22 attached to the rotor blades 12 is equipped on the working medium side with a particularly heat-resistant protective layer such as, for example, ceramic or made of a high-temperature-resistant material.
  • the turbine blade 12, 14 is provided with a number of secondary cooling channels 24 in the leading edge region.
  • the likewise in the leading edge region of the turbine blade 12, 14 mounted and projecting into a main cooling channel 26 outlet channels 28 are used in addition to the leadership of the cooling medium K as fastening elements for the heat protection element 22.
  • cooling medium K cooling air K flows, due to the main cooling channel 26 of the blade body 30 prevailing higher pressure relative to the ambient pressure in the turbine unit 6, via the subcooling channels 24 in the intermediate space formed between the outer surface 32 of the blade body 30 and the inner surface 34 of the heat protection member 22 and through the outlet channels 28 of the heat protection member 22, wherein the effluent from the outlet channels 28 Cooling air K forms a protective film between the working medium M and the outer surface 36 of the heat protection member 22.
  • the cooling air K escaping from the secondary cooling channels 24 of the blade main body 30 flows against the inner surface 34 of the heat protection element 22 and cools them by the resulting impact chilling effect.
  • FIG. 3 and 4 show the heat protection element 22 in two different sectional views, it being apparent from the longitudinal section of the heat protection element 22 shown in FIG. 3 that the outlet channels 28 are arranged one behind the other viewed in the longitudinal direction of the heat protection element 22 and wherein each outlet channel 28 extends from the outer surface 36 of the heat protection member 22 extends to the inner surface 34 toward. As shown in FIG. 4, the exit channels 28 can be arranged centrally in the longitudinal direction of the heat protection element 22.
  • the heat protection element 22 has a shape adapted to the profile of the blade body 30 in the leading edge region.
  • the turbine blade 12, 14 also has a flow-optimized shape after attachment of the heat protection element 22 to the blade main body 30.
  • such a bent heat protection element 22 causes a constant distance between the inner surface 34 of the heat protection element 22 and the outer surface 32 of the blade body 30, whereby a particularly effective cooling in this area is made possible.
  • the cooling air K required for cooling flows from the main cooling channel 26 of the turbine blade 12, 14 through the auxiliary cooling channels 24 and the outlet channels 28, resulting in a cooling film on the Outer surface 36 of the heat protection element 22 forms.
  • the auxiliary cooling channels 24 are preferably arranged such that the cooling air K flowing out of the auxiliary cooling channels 24 impinges perpendicular to the inner surface 34 of the heat protection element 22.
  • the distance of the heat protection element 22 to the blade body 30 is preferably to be chosen so that an intimate contact between the cooling air K and the baffle is effected by a sufficiently high flow rate of the cooling medium K when hitting the inner surface 34 of the heat protection element 22 and thus the impact cooling effect established.
  • FIG. 1 A particularly expedient embodiment of the turbine blade 12, 14 with the heat protection element 22 according to the invention is shown in FIG.
  • the heat protection element 22 has been integrated into the leading edge region of the blade body 30, whereby the original outer shape of the turbine blade 12, 14 is maintained in an advantageous manner.
  • the aerodynamic design of the turbomachine is thus not changed, whereby a reduction in the efficiency of the gas turbine, for example, due to vortex formations on the outer edges in a externally attached to the blade body 30 heat protection element 22 is prevented.
  • the gap between the heat protection element 22 and the blade main body 30 required for producing impingement cooling is achieved in this special embodiment of the turbine blade 12, 14 by placing the heat protection element 22 on a recess 38 present in the blade base body 30. In this way, the outer surface of the turbine blade 12, 14 extending into the flow channel of the gas turbine is partially formed by the outer surface of the heat protection element 22.
  • the free ends of the heat protection element 22 according to FIG 5 are in the embodiment of FIG 6 to that of the base body 30th formed blade walls flush formed to achieve a non-offset surface of the turbine blade 12, 14.
  • the part of the main body 30, which is opposite to the heat protection element 22 set back to the blade interior, so that the edge regions of the heat protection element 22 is connected to the blade body.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP06022622A 2006-10-30 2006-10-30 Aube de turbine Withdrawn EP1930544A1 (fr)

Priority Applications (11)

Application Number Priority Date Filing Date Title
EP06022622A EP1930544A1 (fr) 2006-10-30 2006-10-30 Aube de turbine
PCT/EP2007/059989 WO2008052846A1 (fr) 2006-10-30 2007-09-20 Aube de turbine
CNA2007800407001A CN101535601A (zh) 2006-10-30 2007-09-20 涡轮叶片
JP2009533769A JP2010508461A (ja) 2006-10-30 2007-09-20 タービン翼
AT07820422T ATE459784T1 (de) 2006-10-30 2007-09-20 Turbinenschaufel
EP07820422A EP2084368B1 (fr) 2006-10-30 2007-09-20 Aube de turbine
RU2009120554/06A RU2405940C1 (ru) 2006-10-30 2007-09-20 Турбинная лопатка
US12/447,972 US20100068069A1 (en) 2006-10-30 2007-09-20 Turbine Blade
DE502007003043T DE502007003043D1 (de) 2006-10-30 2007-09-20 Turbinenschaufel
PL07820422T PL2084368T3 (pl) 2006-10-30 2007-09-20 Łopatka turbiny
ES07820422T ES2341384T3 (es) 2006-10-30 2007-09-20 Alabe para turbina.

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP06022622A EP1930544A1 (fr) 2006-10-30 2006-10-30 Aube de turbine

Publications (1)

Publication Number Publication Date
EP1930544A1 true EP1930544A1 (fr) 2008-06-11

Family

ID=37945313

Family Applications (2)

Application Number Title Priority Date Filing Date
EP06022622A Withdrawn EP1930544A1 (fr) 2006-10-30 2006-10-30 Aube de turbine
EP07820422A Not-in-force EP2084368B1 (fr) 2006-10-30 2007-09-20 Aube de turbine

Family Applications After (1)

Application Number Title Priority Date Filing Date
EP07820422A Not-in-force EP2084368B1 (fr) 2006-10-30 2007-09-20 Aube de turbine

Country Status (10)

Country Link
US (1) US20100068069A1 (fr)
EP (2) EP1930544A1 (fr)
JP (1) JP2010508461A (fr)
CN (1) CN101535601A (fr)
AT (1) ATE459784T1 (fr)
DE (1) DE502007003043D1 (fr)
ES (1) ES2341384T3 (fr)
PL (1) PL2084368T3 (fr)
RU (1) RU2405940C1 (fr)
WO (1) WO2008052846A1 (fr)

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CH700687A1 (de) * 2009-03-30 2010-09-30 Alstom Technology Ltd Gekühltes bauteil für eine gasturbine.
US9399271B2 (en) * 2010-12-28 2016-07-26 Rolls-Royce Corporation Heat treating and brazing of an object
US10370981B2 (en) 2014-02-13 2019-08-06 United Technologies Corporation Gas turbine engine component cooling circuit with respirating pedestal
US10301945B2 (en) * 2015-12-18 2019-05-28 General Electric Company Interior cooling configurations in turbine rotor blades
CN105422188A (zh) * 2016-01-13 2016-03-23 北京航空航天大学 一种带隔热罩式复合冷却结构的涡轮叶片
US10577942B2 (en) * 2016-11-17 2020-03-03 General Electric Company Double impingement slot cap assembly
US20190024513A1 (en) * 2017-07-19 2019-01-24 General Electric Company Shield for a turbine engine airfoil
KR102028803B1 (ko) * 2017-09-29 2019-10-04 두산중공업 주식회사 가스 터빈
US11286783B2 (en) * 2020-04-27 2022-03-29 Raytheon Technologies Corporation Airfoil with CMC liner and multi-piece monolithic ceramic shell

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB841117A (en) * 1957-08-02 1960-07-13 Rolls Royce Improvements in or relating to stator blades of fluid machines
GB995182A (en) * 1963-06-20 1965-06-16 Rolls Royce Improvements in or relating to gas turbine engine combustion equipment
US3269700A (en) * 1964-12-07 1966-08-30 United Aircraft Corp Heat shield for turbine strut
GB1218371A (en) * 1967-09-29 1971-01-06 Trw Inc Improvements in or relating to aerofoil vanes or blades for high temperature use

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS527482B2 (fr) * 1972-05-08 1977-03-02
US4026659A (en) * 1975-10-16 1977-05-31 Avco Corporation Cooled composite vanes for turbine nozzles
JPS5390509A (en) * 1977-01-20 1978-08-09 Koukuu Uchiyuu Gijiyutsu Kenki Structure of air cooled turbine blade
EP0954680B1 (fr) * 1996-12-02 2002-02-06 Siemens Aktiengesellschaft Aube de turbine et son utilisation dans un systeme de turbine a gaz
DE19848104A1 (de) * 1998-10-19 2000-04-20 Asea Brown Boveri Turbinenschaufel
US6709230B2 (en) * 2002-05-31 2004-03-23 Siemens Westinghouse Power Corporation Ceramic matrix composite gas turbine vane
US7011502B2 (en) * 2004-04-15 2006-03-14 General Electric Company Thermal shield turbine airfoil
US7871246B2 (en) * 2007-02-15 2011-01-18 Siemens Energy, Inc. Airfoil for a gas turbine

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB841117A (en) * 1957-08-02 1960-07-13 Rolls Royce Improvements in or relating to stator blades of fluid machines
GB995182A (en) * 1963-06-20 1965-06-16 Rolls Royce Improvements in or relating to gas turbine engine combustion equipment
US3269700A (en) * 1964-12-07 1966-08-30 United Aircraft Corp Heat shield for turbine strut
GB1218371A (en) * 1967-09-29 1971-01-06 Trw Inc Improvements in or relating to aerofoil vanes or blades for high temperature use

Also Published As

Publication number Publication date
RU2405940C1 (ru) 2010-12-10
EP2084368A1 (fr) 2009-08-05
ATE459784T1 (de) 2010-03-15
EP2084368B1 (fr) 2010-03-03
PL2084368T3 (pl) 2010-07-30
US20100068069A1 (en) 2010-03-18
ES2341384T3 (es) 2010-06-18
JP2010508461A (ja) 2010-03-18
DE502007003043D1 (de) 2010-04-15
WO2008052846A1 (fr) 2008-05-08
CN101535601A (zh) 2009-09-16

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