US2843354A - Turbine and like blades - Google Patents

Turbine and like blades Download PDF

Info

Publication number
US2843354A
US2843354A US171263A US17126350A US2843354A US 2843354 A US2843354 A US 2843354A US 171263 A US171263 A US 171263A US 17126350 A US17126350 A US 17126350A US 2843354 A US2843354 A US 2843354A
Authority
US
United States
Prior art keywords
elements
blade
passage
passageway
flow
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.)
Expired - Lifetime
Application number
US171263A
Other languages
English (en)
Inventor
Austin G Smith
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.)
Power Jets Research and Development Ltd
Original Assignee
Power Jets Research and Development Ltd
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 Power Jets Research and Development Ltd filed Critical Power Jets Research and Development Ltd
Application granted granted Critical
Publication of US2843354A publication Critical patent/US2843354A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • 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
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2212—Improvement of heat transfer by creating turbulence

Definitions

  • This invention relates to cooling arrangements for tun bine and like blades required to operate at high tern peratures (e. g. above 700 C.), and is concerned with that class of cooling arrangement in which a blade is provided with an internal passage through which, in the operation of the blade, flows a coolant fluid.
  • the present invention resides in the realisation that in such a cooling arrangement it is advantageous to vary the internal conductance from place to place along the length of the passage, co iductance, being defined as ocS, where a is the coefilcient of heattransfer and S the surface area per unit length'of the passage. It is envisaged that the manner of such variation should be determined on the basis of either or both of two con siderations: firstly, the need to, vary the cooling provision along the length of the passage in the general sense required by variation of the gas temperature and/ or mechanical stress; secondly, the desirability of effecting the requisite cooling with maximum economy in the use of coolant, which implies a close correspondence between the cooling required at any station in the passage and that actually provided.
  • the coolant when discharged into the turbine working fluid it should, from the point of view of plant efficiency, be at the maximum practicable temperature when it emerges into the working fluid flow, which implies limitation of the rate of flow of coolant and a high coolant temperature at the outlet end of the coolant passage resulting in correspondingly reduced cooling capacity in that region.
  • the former consideration would indicate on the one hand that the internal heat transfer surface should be designed on the basis of securing maximumconductance in the region at which the most onerous stress-temperature combination occurs (which might be, for example, in the case of an axial flow turbine,- at about on third of the blade height from the root), and on'the other that the conductance should increase progressively in the direction of coolant flow to compensate for the rising temperature of the coolant.
  • the second factor implies the formulation of some method of determining cooling requirements. Such a method will be described more fully hereafter with referenceto. theaccompanying drawings.
  • the internal conductance OLS can be varied, namely by varying or by changing the form of the heat conducting surface while keeping S constant, or varying the value of S while keeping a constant, or by varying both these quantities, as by varying the cross sectional area of the passage and hence the coolant velocity.
  • the present invention also makes use within the blade cooling passage of an extended heat conducting formation to increase the aggregate internal conductance of the passage, but additionally proposes to vary the conductance by varying from place to place the form or distribution of the heat exchange surface provided by said formation.
  • an inserted heat conducting structure from a plurality of discrete, elements mounted in succession along the length of and Within the. coolant passage as an assembly unitary with the blade and in heatconducting relation therewith, each element being gapped to permit the flow of coolant through or around it and along the passage.
  • the discrete elements making up the heat conducting insert may take a variety of forms such as to affordv a succession of partitions or similar obstructions across or in the coolant passage, the flow of coolant being permitted by passages slaughterded either by the shape of or by holes in the elements, and successive elements may be designed to provide for a tortuous coolant path either by the use of diiferent combinations of holes therein or by diiferentlyarranging successive similar elements with respect to the axis of the passage.
  • the elements may also be either separately introduced into the blade passage and subsequently secured therein, in which case their location can readily be controlled by the use of a rarnrod whose penetration into the passage is accurately controlled, or may be previously as Sild'on acarrier and inserted into and secured in the blade passage as a group.
  • the final ses curing of the elements in the passage may be achieved by brazing,vthe efiectiveness of which may be aided by the use of elements having an extended rim, and in the second case the carrier (e. g. a central rod) may remain as a partof the finished unit.
  • the passages into which the elements are positione may be formed in various ways.
  • the blade may be solid and the passages drilled out from the solid blade; or the blade may be cast with preformed passages which may be subsequently suitably worked to the required dimensions; or (more particularly in the case when it is desired that the elements should be noncircular) the blade may be formed from a number of laminae each having an aperture whose section corresponds to that ofthe required passage, these laminae being subsequently brazed together to form a block from which the blade may be machined.
  • Figure 1 is a transverse section through a turbine blade having a number of internal passages.
  • Figure 2 is a fragmentary longitudinal section on the line IIII in Figure 1 and to an enlarged scale showing in perspective the arrangement of inserted elements.
  • Figure 3 is a side view of a blade partly in longitudinal section showing the general arrangement of a coolant p-assage therein and a possible distribution of inserted elements in said passage.
  • Figure 4 is a view similar to that of Figure 2 showing an alternative form of element.
  • Figures 5 and 6 are respectively a face view and axial section of an alternative form of insert element.
  • Figure 7 is a side view of a spacer for use therewith.
  • Figure 8 is a fragmentary section to an enlarged scale of the insert elements and spacers of Figures 5 to 7 assembled in a blade passage.
  • Figure 9 shows curves illustrating a method of choosing the manner of variation of the internal conductance of a passage.
  • Figure 10 shows a modification of the arrangement shown in Figure 8.
  • a blade 1 is provided with longitudinal passages 2a, 2b, 2c, 2d and 2e.
  • the passages 2a, 2b, and 2c are circular or the passage 2d is approximately rectangular and the passage 2e is approximately triangular.
  • Within these passages are placed a number of thin sheet elements extending transversely across the passage and acting as an assemblage of cooling fins.
  • the elements in the passage 2a take the form of perforated discs 3 having a number of apertures 3a which are arranged in a regular pattern over the entire surface of the disc in simulation of wire mesh, those in passage 2b are gauze discs 4 having any number of wires, and those in passages 2d and 2c are perforated plates 5, 6.
  • the elements comprise a number of perforated discs 7 alternating with imperforate discs 8.
  • Each disc 7 has a central aperture 7a, while each disc 8 is larger than the aperture 7a andis secured to the inner wall of the passage 2 by means of radial arms 9.
  • Figure 3 shows a blade 1 having a longitudinal passage 2, in which a number of elements are arranged at varied spacing along the passage, being closer at the region 19 intermediate the root and tip ends of the passage 2 than at other regions along the passage 2,, this distribution typifying one arrived at solely on the assumed basis that 10 is the region of most onerous stresstemperature relationship.
  • the elements may be entered into and located in the passage by the use of a ramrod and secured by copper or other high melting point brazing.
  • Figure 4 shows a. perforated disc element 11, which is dished in order to enable centrifugal force to assist the attachment of the element to the inner wall of the passage 2.
  • the inserted elements are in good heat conducting contact at their edges with the inner wall of the passage in which they are inserted. Further they have surfaces exposed to the flow of coolant fluid through the passages, which surfaces constitute heat exchange surfaces for transferring heat from the elements and hence from the blade to the coolant fluid.
  • the blanks for the elements 12, aftr drilling may be mounted on the rod 13 without spacers and milled to form their radial arms, the group being thereafter dismantled and reassembled with appropriate spacers and with their radial arms out of phase, and the whole group inserted into a. blade passage and copper brazed in position.
  • the elements 12 may then be thought of as cooling fins extending from the rod 13. It may be contrived that at this stage any securing nuts, if used, and possibly also the carrier rod, are dispensed with, leaving the elements and spacers secured to each other and the blade solely by brazing.
  • the heat exchange surfaces are constituted by the side and end surfaces of the radial arms of the elements 12.
  • the spacing of the elements is varied along the length of the passage 2, the spacing decreasing from one end to the other, and so the area of the heat exchange surface per unit length of the passage and the internal conductance of the passage varies in a corresponding manner.
  • the area of the heat exchange surface per unit length of passage is further varied by making the elements 12 of different axial lengths.
  • a turbine blade having root and tip ends and a portion defining an internal passageway extending lengthwise of the blade from the root end to the tip end to permit the flow of coolant fluid, and a plurality of elements within said passageway and extending transversely with respect thereto and axially spaced from one another along the longitudinal axis thereof, said elements being apertured to permit the flow of said fluid along the passageway, being in good heat conducting contact at their edges with the passagewaydefining portion of the blade and having heat exchange surfaces exposed to the flow of said fluid, said elements being variably spaced along at least part of the passageway so that the area of said heat exchange surfaces per unit length of the passageway varies along at least this part of the passageway.
  • a turbine blade having root and tip ends and a portion defining an internal passageway extending lengthwise of the blade from the root end to the tip end to permit the flow of coolant fluid, and a plurality of elements within said passageway and extending transversely with respect thereto and axially spaced from one another along the longitudinal axis thereof, said elements being apertured to permit the flow of said fluid along the passageway, being in good heat-conducting contact at their edges with the passageway-defining portion of the blade and having heat exchange surfaces exposed to the flow of said fluid, said elements being variably spaced along the passageway and being closer together at the region intermediate the root and tip ends of the blade than at the root and tip ends.
  • a turbine blade having root and tip ends and a portion defining an internal passageway extending lengthwise of the blade from the root end to the tip end to permit the flow of coolant fluid, and a plurality of elements within said passageway and extending transversely with respect thereto and axially spaced from one another along the longitudinal axis thereof, said elements being apertured to permit the flow of said fluid along the passageway, being in good heat-conducting contact at their edges with the passage-defining portion of the blade and having heat exchange surfaces exposed to the flow of said fluid, said elements being variably spaced along the passageway, the spacing decreasing from one end of the passageway to the other.
  • a turbine blade having root and tip ends and a portion defining an internal passageway extending lengthwise of the blade from the root end to the tip end to permit the flow of coolant fluid, a central support extending lengthwise of the passageway and a plurality of elements mounted on said support within saidpassageway and, extending transversely, with respect theretoand axially spaced from one another along the longitudinal axis thereof, said elements being apertured to permit the flow of saidfiuid along the passageway, being in good heatconducting contact at their edges with the passageway-defining portion of the blade and having heat exchange; surfaces exposed to the flow of said fluid, said, elements being variably spacedqalong at least partof the passageway so that the area of said heat exchange surfaces per unitlength of thepassageway varies along at least this parttofthe assageway.
  • aturbine blade having-root and tip ends and a portion, defining; an internal passageway extending lengthwiseof the blade from the rootend to the tip end to permit the flow'of coolant fluid,.and;a .plurality of elements within said passageway and extending transversely with respect thereto and axially spaced from one another along thelongitudinal axisythereof, said elements being apertured to permit the flow of said fluid, along the passageway, being in good heat-conducting contact at their edges with the passageway-defining portion of the blade and having heat exchange surfaces exposed to the flow of said fluid, said elements being of different dimensions along at least part of the passageway so that the area of said heat exchange surfaces per unit length of the passageway varies along at least this part of the passageway.
  • a turbine blade having root and tip ends and a portion defining an internal passageway extending lengthwise of the blade from the root end to the tip end to permit the flow of coolant fluid, and a plnraiity of elements within said passageway and extending transversely with respect thereto and axially spaced from one another along the longitudinal axis thereof, said elements being apertured to permit the flow of said fluid along the passageway, being in good heat-conducting contact at their edges with the passageway-defining portion of the blade and having heat exchange surfaces exposed to the flow or" said fluid, said elements being of different lengths along at least part of the passageway, so that the area of said heat exchange surfaces per unit length of the passageway varies along at least this part of the passageway.
  • a turbine blade having root and tip ends and a portion defining an internal passageway extending lengthwise of the blade from the root end to the tip end to permit the flow of coolant fluid, and a plurality of elements within said passageway and extending transversely with respect thereto and axially spaced from one another along the longitudinal axis thereof, said elements being apertured to permit the flow of said fluid along the passageway, being in good heat-conducting contact at their edges with the passageway-defining portion of the blade and having heat exchange surfaces exposed to the flow of said fluid, said elements being of different lengths with the greater length elements being at the region intermediate the root and tip ends of the blade.
  • a turbine blade having root and tip ends and a portion defining an internal passageway extending lengthwise of the blade from the root end to the tip end to permit the flow of coolant fluid, and a plurality of elements within said passageway and extending transversely with respect thereto and axially spaced from one another along the longitudinal axis thereof, said ele ments being apertured to permit the flow of said fluid along the passageway, being in good heat-conducting contact at their edges with the passageway-defining portion of the blade and having heat exchange surfaces exposed to the flow of said fluid, said elements being of different lengths and positioned along the longitudinal axis of the passageway with the lengths of said elements decreasing from one end of the passageway to the other.
  • a turbine blade having root and tip ends and a portion defining an internal passageway extending lengthwise of the blade from the root end to the tip end to permit the flow of coolant fluid, a central support extending lengthwise of the passageway and a plurality of elements mounted on said support within said passageway and extending transversely with respect thereto and axially spaced from one another along the longitudinal axis thereof, said elements being apertured to permit the flow of said fluid along the passageway, being in good heat-conducting contact at their edges with the passageway-defining portion of the blade and having heat exchange surfaces exposed to the flow of said fluid, said elements being of different lengths along at least part of the passageway so that the area of said heat exchange surfaces per unit length of the passageway varies along at least this part of the passageway.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US171263A 1949-07-06 1950-06-30 Turbine and like blades Expired - Lifetime US2843354A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB286626X 1949-07-06

Publications (1)

Publication Number Publication Date
US2843354A true US2843354A (en) 1958-07-15

Family

ID=10279147

Family Applications (1)

Application Number Title Priority Date Filing Date
US171263A Expired - Lifetime US2843354A (en) 1949-07-06 1950-06-30 Turbine and like blades

Country Status (6)

Country Link
US (1) US2843354A (enrdf_load_stackoverflow)
BE (1) BE496812A (enrdf_load_stackoverflow)
CH (1) CH286626A (enrdf_load_stackoverflow)
FR (2) FR1022398A (enrdf_load_stackoverflow)
GB (2) GB680118A (enrdf_load_stackoverflow)
NL (1) NL73916C (enrdf_load_stackoverflow)

Cited By (27)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2999669A (en) * 1958-11-21 1961-09-12 Westinghouse Electric Corp Damping apparatus
US3094310A (en) * 1959-12-09 1963-06-18 Rolls Royce Blades for fluid flow machines
US3738771A (en) * 1970-07-20 1973-06-12 Onera (Off Nat Aerospatiale) Rotor blades of rotary machines, provided with an internal cooling system
US4119390A (en) * 1976-11-19 1978-10-10 General Electric Company Liquid-cooled, turbine bucket with enhanced heat transfer performance
US4383854A (en) * 1980-12-29 1983-05-17 General Electric Company Method of creating a controlled interior surface configuration of passages within a substrate
DE3416087A1 (de) * 1984-04-30 1985-10-31 Klöckner-Humboldt-Deutz AG, 5000 Köln Gekuehlte turbinenschaufel
US5002460A (en) * 1989-10-02 1991-03-26 General Electric Company Internally cooled airfoil blade
US5704763A (en) * 1990-08-01 1998-01-06 General Electric Company Shear jet cooling passages for internally cooled machine elements
EP0887515A1 (fr) * 1997-06-26 1998-12-30 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Aubage refroidi par rampe hélicoidale, par impact en cascade et par système à pontets dans une double peau
US20060171808A1 (en) * 2005-02-02 2006-08-03 Siemens Westinghouse Power Corp. Vortex dissipation device for a cooling system within a turbine blade of a turbine engine
US20070014664A1 (en) * 2004-07-26 2007-01-18 Jurgen Dellmann Cooled component of a fluid-flow machine, method of casting a cooled component, and a gas turbine
WO2011117395A1 (de) * 2010-03-26 2011-09-29 Siemens Aktiengesellschaft Komponente mit einer einem heissgas einer gasturbine aussetzbaren aussenwand und verfahren zum herstellen einer derartigen komponente
US8297927B1 (en) * 2008-03-04 2012-10-30 Florida Turbine Technologies, Inc. Near wall multiple impingement serpentine flow cooled airfoil
DE102012017491A1 (de) * 2012-09-04 2014-03-06 Rolls-Royce Deutschland Ltd & Co Kg Turbinenschaufel einer Gasturbine mit Drallerzeugungselement
US8864438B1 (en) * 2013-12-05 2014-10-21 Siemens Energy, Inc. Flow control insert in cooling passage for turbine vane
GB2518379A (en) * 2013-09-19 2015-03-25 Rolls Royce Deutschland Aerofoil cooling system and method
US20150337667A1 (en) * 2014-05-23 2015-11-26 United Technologies Corporation Airfoil cooling device and method of manufacture
WO2016014056A1 (en) * 2014-07-24 2016-01-28 Siemens Aktiengesellschaft Turbine airfoil cooling system with spanwise extending flow blockers
US20170321569A1 (en) * 2016-05-06 2017-11-09 General Electric Company Turbomachine including clearance control system
US20190055849A1 (en) * 2015-11-10 2019-02-21 Siemens Aktiengesellschaft Laminated airfoil for a gas turbine
US10309246B2 (en) 2016-06-07 2019-06-04 General Electric Company Passive clearance control system for gas turbomachine
US10392944B2 (en) 2016-07-12 2019-08-27 General Electric Company Turbomachine component having impingement heat transfer feature, related turbomachine and storage medium
US10605093B2 (en) 2016-07-12 2020-03-31 General Electric Company Heat transfer device and related turbine airfoil
US10738622B2 (en) 2016-08-09 2020-08-11 General Electric Company Components having outer wall recesses for impingement cooling
US20200256194A1 (en) * 2019-02-07 2020-08-13 United Technologies Corporation Blade neck transition
US10871074B2 (en) 2019-02-28 2020-12-22 Raytheon Technologies Corporation Blade/vane cooling passages
RU231796U1 (ru) * 2024-09-30 2025-02-11 Федеральное государственное бюджетное образовательное учреждение высшего образования "Рыбинский государственный авиационный технический университет имени П.А. Соловьева" Охлаждаемая рабочая лопатка турбины газотурбинного двигателя со спиралевидными турбулизаторами

Families Citing this family (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2921769A (en) * 1953-06-08 1960-01-19 Peerless Turbine Corp Turbine rotor
GB778672A (en) * 1954-10-18 1957-07-10 Parsons & Marine Eng Turbine Improvements in and relating to the cooling of bodies subject to a hot gas stream, for example turbine blades
US3220697A (en) * 1963-08-30 1965-11-30 Gen Electric Hollow turbine or compressor vane
US3370829A (en) * 1965-12-20 1968-02-27 Avco Corp Gas turbine blade construction
US4080095A (en) * 1976-09-02 1978-03-21 Westinghouse Electric Corporation Cooled turbine vane
NL7712519A (nl) * 1976-11-19 1978-05-23 Gen Electric Vloeistof gekoelde turbineschoep met verbeterd warmte-overdrachtsgedrag.
US4142831A (en) * 1977-06-15 1979-03-06 General Electric Company Liquid-cooled turbine bucket with enhanced heat transfer performance
GB2163219B (en) * 1981-10-31 1986-08-13 Rolls Royce Cooled turbine blade
US5253976A (en) * 1991-11-19 1993-10-19 General Electric Company Integrated steam and air cooling for combined cycle gas turbines
US5320483A (en) * 1992-12-30 1994-06-14 General Electric Company Steam and air cooling for stator stage of a turbine
US5486093A (en) * 1993-09-08 1996-01-23 United Technologies Corporation Leading edge cooling of turbine airfoils
CN112483191B (zh) * 2020-11-30 2022-07-19 日照黎阳工业装备有限公司 一种适用于燃气轮机具备对流换热功能的涡轮叶片

Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US970349A (en) * 1910-02-25 1910-09-13 Theodore B Stubbs Crude-oil burner.
US1864742A (en) * 1923-06-15 1932-06-28 Charles N Koch Combustion turbine and method of burning fuel
US1928504A (en) * 1932-01-09 1933-09-26 Holzwarth Gas Turbine Co Cooled nozzle segment for combustion gas turbines
GB426885A (en) * 1933-11-03 1935-04-11 Dewandre Co Ltd C Improvements in or relating to heat transmitting tubes
US2043644A (en) * 1933-09-21 1936-06-09 Cyril C Young Heat transfer apparatus
CH213483A (de) * 1938-11-23 1941-02-15 Porsche Kg Laufrad für Brennkraft- oder Dampfturbinen und Verfahren zur Herstellung desselben.
US2297446A (en) * 1938-12-03 1942-09-29 Zellbeck Gustav Hollow blade for exhaust gas turbine rotors
US2407531A (en) * 1942-05-02 1946-09-10 Fed Reserve Bank Elastic fluid mechanism
GB625693A (en) * 1946-10-25 1949-07-01 Brush Electrical Eng Improvements in and relating to turbine blades
US2520373A (en) * 1945-01-24 1950-08-29 Lockheed Aircraft Corp Turbine blade and method of making the same
US2568726A (en) * 1949-08-03 1951-09-25 Franz Anselm Air-cooled turbine blade
US2581252A (en) * 1947-12-31 1952-01-01 Sintercast Corp America Powder metallurgy articles
US2665881A (en) * 1948-06-15 1954-01-12 Chrysler Corp Cooled turbine blade

Patent Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US970349A (en) * 1910-02-25 1910-09-13 Theodore B Stubbs Crude-oil burner.
US1864742A (en) * 1923-06-15 1932-06-28 Charles N Koch Combustion turbine and method of burning fuel
US1928504A (en) * 1932-01-09 1933-09-26 Holzwarth Gas Turbine Co Cooled nozzle segment for combustion gas turbines
US2043644A (en) * 1933-09-21 1936-06-09 Cyril C Young Heat transfer apparatus
GB426885A (en) * 1933-11-03 1935-04-11 Dewandre Co Ltd C Improvements in or relating to heat transmitting tubes
CH213483A (de) * 1938-11-23 1941-02-15 Porsche Kg Laufrad für Brennkraft- oder Dampfturbinen und Verfahren zur Herstellung desselben.
US2297446A (en) * 1938-12-03 1942-09-29 Zellbeck Gustav Hollow blade for exhaust gas turbine rotors
US2407531A (en) * 1942-05-02 1946-09-10 Fed Reserve Bank Elastic fluid mechanism
US2520373A (en) * 1945-01-24 1950-08-29 Lockheed Aircraft Corp Turbine blade and method of making the same
GB625693A (en) * 1946-10-25 1949-07-01 Brush Electrical Eng Improvements in and relating to turbine blades
US2581252A (en) * 1947-12-31 1952-01-01 Sintercast Corp America Powder metallurgy articles
US2665881A (en) * 1948-06-15 1954-01-12 Chrysler Corp Cooled turbine blade
US2568726A (en) * 1949-08-03 1951-09-25 Franz Anselm Air-cooled turbine blade

Cited By (41)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2999669A (en) * 1958-11-21 1961-09-12 Westinghouse Electric Corp Damping apparatus
US3094310A (en) * 1959-12-09 1963-06-18 Rolls Royce Blades for fluid flow machines
US3738771A (en) * 1970-07-20 1973-06-12 Onera (Off Nat Aerospatiale) Rotor blades of rotary machines, provided with an internal cooling system
US4119390A (en) * 1976-11-19 1978-10-10 General Electric Company Liquid-cooled, turbine bucket with enhanced heat transfer performance
US4383854A (en) * 1980-12-29 1983-05-17 General Electric Company Method of creating a controlled interior surface configuration of passages within a substrate
DE3416087A1 (de) * 1984-04-30 1985-10-31 Klöckner-Humboldt-Deutz AG, 5000 Köln Gekuehlte turbinenschaufel
US5002460A (en) * 1989-10-02 1991-03-26 General Electric Company Internally cooled airfoil blade
US5704763A (en) * 1990-08-01 1998-01-06 General Electric Company Shear jet cooling passages for internally cooled machine elements
US5993156A (en) * 1997-06-26 1999-11-30 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Snecma Turbine vane cooling system
FR2765265A1 (fr) * 1997-06-26 1998-12-31 Snecma Aubage refroidi par rampe helicoidale, par impact en cascade et par systeme a pontets dans une double peau
EP0887515A1 (fr) * 1997-06-26 1998-12-30 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Aubage refroidi par rampe hélicoidale, par impact en cascade et par système à pontets dans une double peau
US20070014664A1 (en) * 2004-07-26 2007-01-18 Jurgen Dellmann Cooled component of a fluid-flow machine, method of casting a cooled component, and a gas turbine
US7824156B2 (en) * 2004-07-26 2010-11-02 Siemens Aktiengesellschaft Cooled component of a fluid-flow machine, method of casting a cooled component, and a gas turbine
US20060171808A1 (en) * 2005-02-02 2006-08-03 Siemens Westinghouse Power Corp. Vortex dissipation device for a cooling system within a turbine blade of a turbine engine
US7163373B2 (en) 2005-02-02 2007-01-16 Siemens Power Generation, Inc. Vortex dissipation device for a cooling system within a turbine blade of a turbine engine
US8297927B1 (en) * 2008-03-04 2012-10-30 Florida Turbine Technologies, Inc. Near wall multiple impingement serpentine flow cooled airfoil
WO2011117395A1 (de) * 2010-03-26 2011-09-29 Siemens Aktiengesellschaft Komponente mit einer einem heissgas einer gasturbine aussetzbaren aussenwand und verfahren zum herstellen einer derartigen komponente
EP2372087A1 (de) * 2010-03-26 2011-10-05 Siemens Aktiengesellschaft Komponente mit einer einem Heissgas einer Gasturbine aussetzbaren Aussenwand und Verfahren zum Herstellen einer derartigen Komponente
DE102012017491A1 (de) * 2012-09-04 2014-03-06 Rolls-Royce Deutschland Ltd & Co Kg Turbinenschaufel einer Gasturbine mit Drallerzeugungselement
US9506352B2 (en) 2012-09-04 2016-11-29 Rolls-Royce Deutschland Ltd & Co Kg Turbine blade of a gas turbine with swirl-generating element and method for its manufacture
GB2518379A (en) * 2013-09-19 2015-03-25 Rolls Royce Deutschland Aerofoil cooling system and method
US8864438B1 (en) * 2013-12-05 2014-10-21 Siemens Energy, Inc. Flow control insert in cooling passage for turbine vane
US20150337667A1 (en) * 2014-05-23 2015-11-26 United Technologies Corporation Airfoil cooling device and method of manufacture
US9932835B2 (en) * 2014-05-23 2018-04-03 United Technologies Corporation Airfoil cooling device and method of manufacture
CN106536858B (zh) * 2014-07-24 2019-01-01 西门子公司 具有顺翼展延伸流阻断器的涡轮翼型件冷却系统
WO2016014056A1 (en) * 2014-07-24 2016-01-28 Siemens Aktiengesellschaft Turbine airfoil cooling system with spanwise extending flow blockers
CN106536858A (zh) * 2014-07-24 2017-03-22 西门子公司 具有顺翼展延伸流阻断器的涡轮翼型件冷却系统
US9822646B2 (en) 2014-07-24 2017-11-21 Siemens Aktiengesellschaft Turbine airfoil cooling system with spanwise extending fins
US20190055849A1 (en) * 2015-11-10 2019-02-21 Siemens Aktiengesellschaft Laminated airfoil for a gas turbine
US10221717B2 (en) * 2016-05-06 2019-03-05 General Electric Company Turbomachine including clearance control system
CN107345488A (zh) * 2016-05-06 2017-11-14 通用电气公司 包括间隙控制系统的涡轮机
US20170321569A1 (en) * 2016-05-06 2017-11-09 General Electric Company Turbomachine including clearance control system
US10309246B2 (en) 2016-06-07 2019-06-04 General Electric Company Passive clearance control system for gas turbomachine
US10392944B2 (en) 2016-07-12 2019-08-27 General Electric Company Turbomachine component having impingement heat transfer feature, related turbomachine and storage medium
US10605093B2 (en) 2016-07-12 2020-03-31 General Electric Company Heat transfer device and related turbine airfoil
US10738622B2 (en) 2016-08-09 2020-08-11 General Electric Company Components having outer wall recesses for impingement cooling
US20200256194A1 (en) * 2019-02-07 2020-08-13 United Technologies Corporation Blade neck transition
US11149550B2 (en) * 2019-02-07 2021-10-19 Raytheon Technologies Corporation Blade neck transition
US10871074B2 (en) 2019-02-28 2020-12-22 Raytheon Technologies Corporation Blade/vane cooling passages
RU231796U1 (ru) * 2024-09-30 2025-02-11 Федеральное государственное бюджетное образовательное учреждение высшего образования "Рыбинский государственный авиационный технический университет имени П.А. Соловьева" Охлаждаемая рабочая лопатка турбины газотурбинного двигателя со спиралевидными турбулизаторами
RU2845772C1 (ru) * 2024-12-26 2025-08-25 Федеральное государственное бюджетное образовательное учреждение высшего образования "Рыбинский государственный авиационный технический университет имени П.А. Соловьева" Охлаждаемая сопловая лопатка газовой турбины

Also Published As

Publication number Publication date
GB728834A (en) 1955-04-27
CH286626A (de) 1952-10-31
NL73916C (enrdf_load_stackoverflow) 1900-01-01
BE496812A (enrdf_load_stackoverflow) 1900-01-01
FR1022398A (fr) 1953-03-04
GB680118A (en) 1952-10-01
FR66874E (fr) 1957-10-31

Similar Documents

Publication Publication Date Title
US2843354A (en) Turbine and like blades
US3635585A (en) Gas-cooled turbine blade
US10808551B2 (en) Airfoil cooling circuits
JP2668207B2 (ja) ガスタービンエンジンのタービンのエーロフオイルセクシヨン
US8807943B1 (en) Turbine blade with trailing edge cooling circuit
US10961856B2 (en) Ceramic core for a multi-cavity turbine blade
US3849025A (en) Serpentine cooling channel construction for open-circuit liquid cooled turbine buckets
US7137776B2 (en) Film cooling for microcircuits
JP3053174B2 (ja) ターボ機械に使用するための翼部及びその製造方法
US8162609B1 (en) Turbine airfoil formed as a single piece but with multiple materials
US3017159A (en) Hollow blade construction
US8790083B1 (en) Turbine airfoil with trailing edge cooling
US8303253B1 (en) Turbine airfoil with near-wall mini serpentine cooling channels
US3370829A (en) Gas turbine blade construction
GB895077A (en) Blades for fluid flow machines such as axial flow turbines
US1673554A (en) Blading for reaction turbines
US20160245097A1 (en) Airfoil and method for manufacturing an airfoil
GB1257041A (enrdf_load_stackoverflow)
CN104736797A (zh) 一种翼型及用于构造该翼型的方法
US3314650A (en) Cooled blade
GB791751A (en) Improvements in or relating to blades for axial flow gas turbine engines, and to methods of making such blades
EP1538305B1 (en) Airfoil with variable density array of pedestals at the trailing edge
US2823894A (en) Air-cooled turbine buckets
WO1998045577A1 (de) Verfahren zur kühlung einer turbinenschaufel
CN111485957B (zh) 涡轮导向冷却叶片