US3032314A - Method of and device for cooling the component elements of machines - Google Patents

Method of and device for cooling the component elements of machines Download PDF

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Publication number
US3032314A
US3032314A US735243A US73524358A US3032314A US 3032314 A US3032314 A US 3032314A US 735243 A US735243 A US 735243A US 73524358 A US73524358 A US 73524358A US 3032314 A US3032314 A US 3032314A
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Prior art keywords
blade
cooling
jacket
machines
component elements
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US735243A
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David Otto Arthur
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Safran Aircraft Engines SAS
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SNECMA SAS
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • 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
    • F01D5/182—Transpiration 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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 action of the cooling fluid is effected by simple circulation thereof along one face of the element to be cooled, that is to say if the cooling fluid only washes along the said face as in the case of conventional cooling by circulation, there is a risk that the distribution of temperature at the cooled element may not be uniform. This is particularly the case if the various parts of the element to be cooled are subjected to different temperatures and if on the other hand the fluid travels along a duct or conduit wherein the temperatures vary and along which the cooling fluid itself becomes heated.
  • This non-uniform distribution causes stresses in the molecular structure of this element, and diminishes the mechanical strength thereof.
  • the present invention is concerned with a more rational cooling method which obviates the defects mentioned he-reinbefore.
  • the cooling fluid is made to arrive at a uniform temperature through a jacket which itself has a substantially uniform temperature distribution, and at a pressure greater than the pressure prevailing externally of the jacket, the said fluid escaping from the jacket through appropriately formed and distributed perforations in the foirnof fine jets perpendicular to the wall which is to be cooled, so that the cooling fluid impinges violently against the said wall, instead of merely washing along it as in the case of the conventional methods.
  • the present invention also extends to cooling devices carrying the said methods into effect, and to component elements of machines arranged so as to permit the said method to be carried into effect.
  • FIG. 1 is a diagrammatic sectional view taken on the line I-I of FIG. 2, of a guide blade of the turbine of a turbo-jet engine.
  • FIG. 2 is an elevational view of this blade with a portion thereofbroken away.
  • FIG. 3 is a detail view showing the form of an inner jacket which is lodged in such a blade.
  • FIGS. 1 and 2 The form of embodiment illustrated by way of example in FIGS. 1 and 2 is concerned with a hollow blade 1 made of sheet metal which is to be subjected to cooling atent by internal circulation of air and is welded along the trailing edge on a rib 2 in which are formed apertures 3 whereby the air which has. been used for cooling can escape from the blade and mingle with the passing flow of hot gases.
  • a jacket 4 is fitted in the interior of this blade. It is so dimensioned and profiled that between the wall of the blade and the said jacket there is formed a space 9 which is sufiicient to permit the circulation of the cooling air.
  • This jacket closed at its upper end by a wall 5, is held in place by bosses 6 (FIG. 3) which bear against the wall 1 of the blade and are staggered in a quincunx formation.
  • the jacket 4 comprises holes or perforations 8 situated opposite those parts of the blade wall which are to be more particularly cooled, that is to say the parts where there is a risk of excessive heating.
  • the perforations 8 have been provided in the vicinity of the leading edge of the blade and also in the vicinity of the trailing edge thereof.
  • the cooling air can be taken from the output side of the compressor and penetrates into the jacket 4 through the root of the blade in the direction of the arrows 7. It issues from the jacket through the perforations 8 in the form of fine jets which strike perpendicularly against the Wall of the blade exactly at the points to be cooled.
  • Thethin jets issuing from the perforations 8 and impinging perpendicularly against the wall of the blade 1 produce a cooling effect by impact which gives better results than simple circulation of air through a blade.
  • the method of cooling according to the invention therefore makes it possible, by sending through appropriately arranged perforations jets of air directed perpendicularly to the parts to be cooled, to obtain a better etficiency since on the one hand the air is still relatively cool when issuing from the perforations 8 and its temperature is the same at perforations situated near the blade root and perforations situated near the blade tip, and on the other hand, this air not being heated as it passes through the jacket 4, the heat exchange over the whole height of the blade will be more uniform and therefore stresses caused by great differences in temperature can be avoided.
  • the example chosen has been an entry guide blade for a gas turbine.
  • the invention can be applied to any other component element of a machine which has to be effectively cooled, and more particularly to the rotor blades of the turbine, to support arms situated in the flow of hot gas or to the walls of the combustion chamber or the post-combustion duct.
  • a gas-turbine hollow blade having a leading edge and a trailing edge, with an orifice formed through the said trailing edge, said blade comprising internal means for discharging a coolant gas under pressure onto the inner surface of the leading edge of said hollow blade, means for directing said coolant gas as a thin flow in bathing contact with the inner surface of said hollow blade from the leading edge to the trailing edge thereof, to be discharged through said orifice, and further internal means for discharging additional fresh coolant gas onto the inner surface of said hollow blade adjacent and in ad- Vance of the trailing edge thereof, in a substantially transverse direction with respect to said thin flow.
  • a gas-turbine hollow blade having a leading edge and a trailing edge, with an orifice formed through the said trailing edge, said blade comprising an inner chamher having a wall extending substantially parallel and close to the inner surface of said hollow blade to define therewith a narrow passage extending from the leading edge of said blade to said orifice, a supply of coolant gas under pressure to said chamber, a first series of nozzles formed through said wall opposite the inner surface of said leading edge to discharge thereonto a fraction of the coolant gas supply which is thereby caused to flow towards said orifice through said narrow passage in bathing contact with the inner surface of said hollow blade, and a second series of nozzles formed through said wall opposite the inner surface of said hollow blade adjacent the trailing edge thereof, said latter-mentioned nozzles being directed transversely with respect to said narrow passage.

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

Description

y 1, 1962 o. A. DAVID METHOD OF AND DEVICE FOR COOLING THE COMPONENT ELEMENTS OF MACHINES Flled May 14 1958 United States When the component elements of a machine, and more particularly of a heat engine, have to work at elevated temperatures, technological problems are encountered which are quite diflicult to solve since the maximum temperature permitted by present known materials is relatively low.
Recourse has therefore been had to cooling these elements by subjecting them to the action of a fluid having a low temperature.
If the action of the cooling fluid is effected by simple circulation thereof along one face of the element to be cooled, that is to say if the cooling fluid only washes along the said face as in the case of conventional cooling by circulation, there is a risk that the distribution of temperature at the cooled element may not be uniform. This is particularly the case if the various parts of the element to be cooled are subjected to different temperatures and if on the other hand the fluid travels along a duct or conduit wherein the temperatures vary and along which the cooling fluid itself becomes heated.
This non-uniform distribution causes stresses in the molecular structure of this element, and diminishes the mechanical strength thereof.
The present invention is concerned with a more rational cooling method which obviates the defects mentioned he-reinbefore. According to this method, the cooling fluid is made to arrive at a uniform temperature through a jacket which itself has a substantially uniform temperature distribution, and at a pressure greater than the pressure prevailing externally of the jacket, the said fluid escaping from the jacket through appropriately formed and distributed perforations in the foirnof fine jets perpendicular to the wall which is to be cooled, so that the cooling fluid impinges violently against the said wall, instead of merely washing along it as in the case of the conventional methods.
Tests carried out by the applicant have shown conclusively that the action of jets of cooling fluid directed perpendicularly to the hottest parts is distinctly superior to what can be obtained by simply washing along a part with the said fluid, and that the heat transfer obtained, for example at a blade profile, is markedly increased.
The present invention also extends to cooling devices carrying the said methods into effect, and to component elements of machines arranged so as to permit the said method to be carried into effect.
The description which will now begiven with reference to the accompanying drawings will show clearly the various features of the invention and the manner of carrying them into effect, any feature brought out from the text and from the figures being of course understood to come within the scope of the present invention.
FIG. 1 is a diagrammatic sectional view taken on the line I-I of FIG. 2, of a guide blade of the turbine of a turbo-jet engine.
FIG. 2 is an elevational view of this blade with a portion thereofbroken away.
FIG. 3 is a detail view showing the form of an inner jacket which is lodged in such a blade.
The form of embodiment illustrated by way of example in FIGS. 1 and 2 is concerned with a hollow blade 1 made of sheet metal which is to be subjected to cooling atent by internal circulation of air and is welded along the trailing edge on a rib 2 in which are formed apertures 3 whereby the air which has. been used for cooling can escape from the blade and mingle with the passing flow of hot gases.
According to the present invention, a jacket 4 is fitted in the interior of this blade. It is so dimensioned and profiled that between the wall of the blade and the said jacket there is formed a space 9 which is sufiicient to permit the circulation of the cooling air. This jacket, closed at its upper end by a wall 5, is held in place by bosses 6 (FIG. 3) which bear against the wall 1 of the blade and are staggered in a quincunx formation.
The jacket 4 comprises holes or perforations 8 situated opposite those parts of the blade wall which are to be more particularly cooled, that is to say the parts where there is a risk of excessive heating. In the example illustrated in the drawings, the perforations 8 have been provided in the vicinity of the leading edge of the blade and also in the vicinity of the trailing edge thereof.
The cooling air can be taken from the output side of the compressor and penetrates into the jacket 4 through the root of the blade in the direction of the arrows 7. It issues from the jacket through the perforations 8 in the form of fine jets which strike perpendicularly against the Wall of the blade exactly at the points to be cooled.
This air circulates then through the space 9 in the direction of the arrows 10 and finally escapes therefrom through the apertures 3. It should be noted that the staggered bosses 6 increase the turbulence of the air circulating in the space 9 and promote heat exchange.
Thethin jets issuing from the perforations 8 and impinging perpendicularly against the wall of the blade 1 produce a cooling effect by impact which gives better results than simple circulation of air through a blade. The method of cooling according to the invention therefore makes it possible, by sending through appropriately arranged perforations jets of air directed perpendicularly to the parts to be cooled, to obtain a better etficiency since on the one hand the air is still relatively cool when issuing from the perforations 8 and its temperature is the same at perforations situated near the blade root and perforations situated near the blade tip, and on the other hand, this air not being heated as it passes through the jacket 4, the heat exchange over the whole height of the blade will be more uniform and therefore stresses caused by great differences in temperature can be avoided.
In the preceding description the example chosen has been an entry guide blade for a gas turbine. However it will be apparent that the invention can be applied to any other component element of a machine which has to be effectively cooled, and more particularly to the rotor blades of the turbine, to support arms situated in the flow of hot gas or to the walls of the combustion chamber or the post-combustion duct.
What I claim is:
l. A gas-turbine hollow blade having a leading edge and a trailing edge, with an orifice formed through the said trailing edge, said blade comprising internal means for discharging a coolant gas under pressure onto the inner surface of the leading edge of said hollow blade, means for directing said coolant gas as a thin flow in bathing contact with the inner surface of said hollow blade from the leading edge to the trailing edge thereof, to be discharged through said orifice, and further internal means for discharging additional fresh coolant gas onto the inner surface of said hollow blade adjacent and in ad- Vance of the trailing edge thereof, in a substantially transverse direction with respect to said thin flow.
2. A gas-turbine hollow blade having a leading edge and a trailing edge, with an orifice formed through the said trailing edge, said blade comprising an inner chamher having a wall extending substantially parallel and close to the inner surface of said hollow blade to define therewith a narrow passage extending from the leading edge of said blade to said orifice, a supply of coolant gas under pressure to said chamber, a first series of nozzles formed through said wall opposite the inner surface of said leading edge to discharge thereonto a fraction of the coolant gas supply which is thereby caused to flow towards said orifice through said narrow passage in bathing contact with the inner surface of said hollow blade, and a second series of nozzles formed through said wall opposite the inner surface of said hollow blade adjacent the trailing edge thereof, said latter-mentioned nozzles being directed transversely with respect to said narrow passage.
UNITED STATES PATENTS Meyer Nov. 5, 1940 Thomas July 4, 1950 Newcornb May 22, 1956 Stalker Apr. 2, 1957 Petrie Aug. 12, 1958 Scanlan et al. Sept. 9, 1958 Zimmerman Nov. 4, 1958 Wiese et al. Feb. '17, 1959 FOREIGN PATENTS Switzerland Sept. 1, 1945 France Oct. 18, 1960
US735243A 1957-05-28 1958-05-14 Method of and device for cooling the component elements of machines Expired - Lifetime US3032314A (en)

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Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3191908A (en) * 1961-05-02 1965-06-29 Rolls Royce Blades for fluid flow machines
US3240468A (en) * 1964-12-28 1966-03-15 Curtiss Wright Corp Transpiration cooled blades for turbines, compressors, and the like
US3301527A (en) * 1965-05-03 1967-01-31 Gen Electric Turbine diaphragm structure
US3319593A (en) * 1962-12-24 1967-05-16 Papst Hermann Boundary layer control
US3346235A (en) * 1963-12-23 1967-10-10 Papst Hermann Boundary layer control
US3373970A (en) * 1965-12-11 1968-03-19 Daimler Benz Ag Gas turbine blade
US3384346A (en) * 1966-02-01 1968-05-21 Rolls Royce Aerofoil shaped blade for a fluid flow machine such as a gas turbine engine
US3540810A (en) * 1966-03-17 1970-11-17 Gen Electric Slanted partition for hollow airfoil vane insert
US3574481A (en) * 1968-05-09 1971-04-13 James A Pyne Jr Variable area cooled airfoil construction for gas turbines
US3700348A (en) * 1968-08-13 1972-10-24 Gen Electric Turbomachinery blade structure
US3715170A (en) * 1970-12-11 1973-02-06 Gen Electric Cooled turbine blade
US3767322A (en) * 1971-07-30 1973-10-23 Westinghouse Electric Corp Internal cooling for turbine vanes
US3806276A (en) * 1972-08-30 1974-04-23 Gen Motors Corp Cooled turbine blade
US3806275A (en) * 1972-08-30 1974-04-23 Gen Motors Corp Cooled airfoil
US4021139A (en) * 1974-11-08 1977-05-03 Brown Boveri Sulzer Turbomachinery, Ltd. Gas turbine guide vane
US4105364A (en) * 1975-12-20 1978-08-08 Rolls-Royce Limited Vane for a gas turbine engine having means for impingement cooling thereof
US4359310A (en) * 1979-12-12 1982-11-16 Bbc Brown, Boveri & Company Limited Cooled wall
US6237344B1 (en) 1998-07-20 2001-05-29 General Electric Company Dimpled impingement baffle
EP1039096A3 (en) * 1999-03-22 2003-03-05 General Electric Company Turbine nozzle
US20140321965A1 (en) * 2013-04-24 2014-10-30 Honeywell International Inc. Turbine nozzles and methods of manufacturing the same
US20150322860A1 (en) * 2014-05-07 2015-11-12 United Technologies Corporation Variable vane segment

Families Citing this family (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1204021B (en) * 1959-04-27 1965-10-28 Rolls Royce Blade for axial flow machines, especially gas turbines
US3171631A (en) * 1962-12-05 1965-03-02 Gen Motors Corp Turbine blade
US3316714A (en) * 1963-06-20 1967-05-02 Rolls Royce Gas turbine engine combustion equipment
GB1034260A (en) * 1964-12-02 1966-06-29 Rolls Royce Aerofoil-shaped blade for use in a fluid flow machine
IN163070B (en) * 1984-11-15 1988-08-06 Westinghouse Electric Corp
GB2439330B (en) * 2006-06-22 2008-09-17 Rolls Royce Plc Aerofoil

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* Cited by examiner, † Cited by third party
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US2220420A (en) * 1938-02-08 1940-11-05 Bbc Brown Boveri & Cie Means for cooling machine parts
CH237475A (en) * 1942-06-09 1945-04-30 Vorkauf Heinrich Method and device for cooling gas turbine blades.
US2514105A (en) * 1945-12-07 1950-07-04 Thomas Wilfred Airfoil conditioning means
US2746671A (en) * 1950-04-14 1956-05-22 United Aircraft Corp Compressor deicing and thrust balancing arrangement
US2787049A (en) * 1952-05-23 1957-04-02 Stalkcr Dev Company Process of fabricating blades for turbines, compressors and the like
US2847185A (en) * 1953-04-13 1958-08-12 Rolls Royce Hollow blading with means to supply fluid thereinto for turbines or compressors
US2851216A (en) * 1954-01-13 1958-09-09 Schwarzkopf Dev Co Device adapted for respiration cooling and process of making same
US2859011A (en) * 1953-07-27 1958-11-04 Gen Motors Corp Turbine bucket and liner
US2873944A (en) * 1952-09-10 1959-02-17 Gen Motors Corp Turbine blade cooling

Family Cites Families (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE906636C (en) * 1944-01-23 1954-03-15 Herwig Kress Dr Ing Gas or exhaust gas turbines with hollow blade cooling
BE498667A (en) * 1949-08-27
DE949954C (en) * 1952-02-09 1956-09-27 Helmut Reining Device on cooling frames for Siemens-Martin ovens or other industrial ovens
GB740597A (en) * 1953-11-07 1955-11-16 Gen Motors Corp Improvements relating to gas turbine or compressor blades
US2780435A (en) * 1953-01-12 1957-02-05 Jackson Thomas Woodrow Turbine blade cooling structure

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2220420A (en) * 1938-02-08 1940-11-05 Bbc Brown Boveri & Cie Means for cooling machine parts
CH237475A (en) * 1942-06-09 1945-04-30 Vorkauf Heinrich Method and device for cooling gas turbine blades.
US2514105A (en) * 1945-12-07 1950-07-04 Thomas Wilfred Airfoil conditioning means
US2746671A (en) * 1950-04-14 1956-05-22 United Aircraft Corp Compressor deicing and thrust balancing arrangement
US2787049A (en) * 1952-05-23 1957-04-02 Stalkcr Dev Company Process of fabricating blades for turbines, compressors and the like
US2873944A (en) * 1952-09-10 1959-02-17 Gen Motors Corp Turbine blade cooling
US2847185A (en) * 1953-04-13 1958-08-12 Rolls Royce Hollow blading with means to supply fluid thereinto for turbines or compressors
US2859011A (en) * 1953-07-27 1958-11-04 Gen Motors Corp Turbine bucket and liner
US2851216A (en) * 1954-01-13 1958-09-09 Schwarzkopf Dev Co Device adapted for respiration cooling and process of making same

Cited By (24)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3191908A (en) * 1961-05-02 1965-06-29 Rolls Royce Blades for fluid flow machines
US3319593A (en) * 1962-12-24 1967-05-16 Papst Hermann Boundary layer control
US3346235A (en) * 1963-12-23 1967-10-10 Papst Hermann Boundary layer control
US3240468A (en) * 1964-12-28 1966-03-15 Curtiss Wright Corp Transpiration cooled blades for turbines, compressors, and the like
US3301527A (en) * 1965-05-03 1967-01-31 Gen Electric Turbine diaphragm structure
US3373970A (en) * 1965-12-11 1968-03-19 Daimler Benz Ag Gas turbine blade
US3384346A (en) * 1966-02-01 1968-05-21 Rolls Royce Aerofoil shaped blade for a fluid flow machine such as a gas turbine engine
US3540810A (en) * 1966-03-17 1970-11-17 Gen Electric Slanted partition for hollow airfoil vane insert
US3574481A (en) * 1968-05-09 1971-04-13 James A Pyne Jr Variable area cooled airfoil construction for gas turbines
US3700348A (en) * 1968-08-13 1972-10-24 Gen Electric Turbomachinery blade structure
US3715170A (en) * 1970-12-11 1973-02-06 Gen Electric Cooled turbine blade
US3767322A (en) * 1971-07-30 1973-10-23 Westinghouse Electric Corp Internal cooling for turbine vanes
JPS5134925B1 (en) * 1971-07-30 1976-09-29
US3806276A (en) * 1972-08-30 1974-04-23 Gen Motors Corp Cooled turbine blade
US3806275A (en) * 1972-08-30 1974-04-23 Gen Motors Corp Cooled airfoil
US4021139A (en) * 1974-11-08 1977-05-03 Brown Boveri Sulzer Turbomachinery, Ltd. Gas turbine guide vane
US4105364A (en) * 1975-12-20 1978-08-08 Rolls-Royce Limited Vane for a gas turbine engine having means for impingement cooling thereof
US4359310A (en) * 1979-12-12 1982-11-16 Bbc Brown, Boveri & Company Limited Cooled wall
US6237344B1 (en) 1998-07-20 2001-05-29 General Electric Company Dimpled impingement baffle
EP1039096A3 (en) * 1999-03-22 2003-03-05 General Electric Company Turbine nozzle
US20140321965A1 (en) * 2013-04-24 2014-10-30 Honeywell International Inc. Turbine nozzles and methods of manufacturing the same
US9719362B2 (en) * 2013-04-24 2017-08-01 Honeywell International Inc. Turbine nozzles and methods of manufacturing the same
US20150322860A1 (en) * 2014-05-07 2015-11-12 United Technologies Corporation Variable vane segment
US10066549B2 (en) * 2014-05-07 2018-09-04 United Technologies Corporation Variable vane segment

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DE1092255B (en) 1960-11-03
FR1177035A (en) 1959-04-20
GB853328A (en) 1960-11-02

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