US3632221A - Gas turbine engine cooling system incorporating a vortex shaft valve - Google Patents

Gas turbine engine cooling system incorporating a vortex shaft valve Download PDF

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Publication number
US3632221A
US3632221A US60320A US3632221DA US3632221A US 3632221 A US3632221 A US 3632221A US 60320 A US60320 A US 60320A US 3632221D A US3632221D A US 3632221DA US 3632221 A US3632221 A US 3632221A
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Prior art keywords
compressor
turbine engine
gas turbine
vortex
recited
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US60320A
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English (en)
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Donald E Uehling
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General Electric Co
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General Electric Co
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/16—Cooling of plants characterised by cooling medium
    • F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • 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
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10S—TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10S415/00—Rotary kinetic fluid motors or pumps
    • Y10S415/914—Device to control boundary layer

Definitions

  • a plurality of small nozzles are positioned within this chamber with their outlets directed opposite the direction of [56] References Cited rotation of the split disc. Control flow is directed through UNITED STATES PATENTS these nozzles to counteract a natural vortex which is formed within the chamber due to rotation of the disc. The strength of l fgg z the vortex field is then utilized to control the amount of cool- 2'79l 5/1957 wheagey i 415 6 ing flow which passes through the chamber and thus ultimate- 2:9l0, 268 10/1959 Davies et al. 4l5/DIG. 1 to the turbine Swim iNVENT OR. DONALD E.
  • This invention relates generally to a gas turbine engine cooling system and, more particularly, to such a system which incorporates a vortex device to regulate the amount of cooling flow delivered to the turbine section of the engine.
  • Gas turbine engines of the type generally referred to herein normally comprise a compressor, a primary combustion system, a turbine, a tailpipe (possibly including an augmenter combustion system), and a variable area exhaust nozzle.
  • Turbofan engines which include additional fan stages and fan turbine drives, utilize gas generators as described in the previous sentence to drive the fan section.
  • air enters an inlet and is compressed within the compressor, ignited along with metered fuel in the primary combustion system to generate a high-energy gas system, performs work while expanding through the turbine, and exits through the variable area nozzle.
  • the high energy associated with the gas exiting from the exhaust nozzle provides forward thrust to an aircraft powered by such an engine.
  • the problem generally associated with increasing the gas temperature of the primary airstream generated in the combustion system described above is that the components located downstream of the combustion system are incapable of withstanding such high temperatures.
  • the components seeing the highest temperatures, other than the combustion system components themselves, are the turbine vanes and blades located immediately downstream of the combustors.
  • this cooling air is taken from the primary airflow engine compressor somewhere along the axial length thereof and is piped to the turbine section.
  • the cooling air is passed through the turbine blades through openings provided in a face of a rotor which supports the turbine blade and which is exposed to a cooling chamber pressurized by compressor discharge air.
  • the turbine wheel(s) or disc(s) may also be cooled by flowing the cooling air on the way to the turbine blades through radial passageways in the turbine wheel.
  • engine gas temperatures are increased in order to improve gas turbine engine performance levels.
  • the peak temperatures previously mentioned are normally not seen during an entire engine flight mission. That is, as engine thrust levels are varied by the pilot, the gas temperatures within the engine also vary.
  • turbine cooling air is normally bleed air extracted from the compressor. In order to optimize engine efficiency, however, it is necessary to minimize the amount of cooling air which is extracted from the primary airflow during engine operation. Thus, it is highly desirable to have the capability to vary the amount of cooling air delivered to the turbine and turbine disc as the temperatures thereof vary.
  • an aircraft gas turbine engine with at least one compressor disc which is separated or split to provide a passageway for the flow of cooling air from the circumference to the bore thereof.
  • Located within the passageway and connected to the compressor disc are a plurality of air jets having outlets directed in a direction opposite the direction of rotation of the compressor disc. These air jets are fluidically connected to a source of higher pressure air than that associated with the passageway.
  • Valve means are provided to vary the amount of pressurized air delivered to the air jets.
  • the air jets are utilized to vary the strength of a natural vortex which tends to be generated in the cooling flow within the cooling passageway as the compressor disc rotates. The strength of I this vortex field is then utilized to control the amount of cooling flow which passes through the passageway, and thus the amount of flow which passes to the turbine section located downstream of the compressor.
  • FIG. 1 is a graphical plot of turbine inlet temperature versus engine rotor speed
  • FIG. 2 is a schematic view, in axial cross section, of a portion of a gas turbine engine constructed in accordance with this invention
  • FIG. 3 is an enlarged view of portions of FIG. 2;
  • FIG. 4 is a view taken along lines 4-4 of FIG. 3;
  • FIG. 5 is an axial cross-sectional view of an alternative embodiment.
  • FIG. 1 a plot of turbine inlet temperature versus engine speed for a typical gas turbine engine or turbofan engine is shown. This plot shows that the temperature seen by a turbine components varies greatly with engine speed.
  • the amount of cooling flow delivered to the turbine components must be capable of maintaining the components at an acceptable temperature when turbine inlet temperature is at its maximum (generally at maximum engine speed). This amount of cooling flow is much higher than that required, however, when the turbine inlet temperature has substantially decreased (at lower engine speeds). Therefore, in order to maintain engine efficiency it is desirable to provide some means for varying the amount of cooling flow delivered to the turbine.
  • FIG. 2 wherein a gas turbine engine 10 is shown to include an outer casing 12 in which air is pressurized by a compressor 14 for delivery to a combustor 16.
  • a gas turbine engine 10 is shown to include an outer casing 12 in which air is pressurized by a compressor 14 for delivery to a combustor 16.
  • fuel is injected into the combustor 16 through a plurality of fuel nozzles 18 (only one of which is shown), and the resultant fuel/air mixture is ignited to generate a high-energy gas stream.
  • the high-energy gas stream which is designated by the numeral 20, then passes to a turbine 22 through a stationary inlet nozzle 24 which is positioned at the downstream end of the combustor 16.
  • the high-energy gas stream then passes through an engine exhaust nozzle (not shown) to provide a propulsive force to the engine 10.
  • the outer bounds of the annular flow path are defined by the outer engine casing 12 which has circumferentially mounted stator vanes 30 extending into the gas stream between each row of rotating blades 28.
  • the annular flow path continues through an outlet guide vane 32 positioned downstream of the compressor 14. Just aft of the outlet guide vane 32 a snout assembly 34, which forms the upstream end of the combustor l6, breaks the annular flow path into cooling passageways 36 and 38 and diffuser passageway 40. Thus, a portion of the compressor discharge air flows around the combustor 16 to cool the same, while the remaining air enters the combustor 16 through the diffuser passage 40 and is ignited therein to form the high-energy gas stream 20.
  • the high-energy gas stream thus provided next flows through an annular flow path for the turbine 22, which flow path is defined by the circumferences of a pair of discs 42 which are interconnected for rotation by a heat shield 44 and an annular torque member 46.
  • Circumferentially mounted, hollow turbine blades 45 extend from the discs 42 into the annular flow path thus described.
  • the compressor discs 26 and the turbine discs 42 are interconnected for rotation by opposing conical rotor elements 52, 54, and a tubular shaft 56, which cooperate to form what is normally referred to as the engine rotor.
  • the engine rotor is joumaled for rotation by means which are not shown as they form no part of the present invention but which would include thrust bearings, frame members, etc. These members would be lubricated in any manner known in the art.
  • the temperature of the high-energy gas stream 20 is sufficiently high that it is desirable to provide cooling air to certain components of the turbine section.
  • one or more of the compressor discs 26 is separated into two separated disc portions 62 in order to provide a cooling air passageway 64 therebetween.
  • the cooling passageway 64 is fluidically connected to the compressor annular flow path by means of a plurality of radial holes 66 bored through the circumference of the compressor disc.
  • the cooling passageway 64 thus delivers a certain percentage of the gas flow from the compressor annular flow path to a plenum chamber 68 (FIG. 2) formed partially by the conical rotor element 52, the tubular shaft 56, and the circumferential portions of the compressor discs 26.
  • the cooling air next flows into a second plenum chamber 70 through a plurality of airholes 72 positioned within the conical rotor element 52.
  • the plenum chamber 70 is formed by the conical rotor elements 52 and 54 and by a generally cylindrical shaped flow member 74 which is supported on its opposite ends by the outlet guide vanes 32 and the turbine inlet nozzle 24.
  • the wall member 74 further defines the cooling passage 38 previously discussed.
  • the cooling air which is generally designated by the numerals 76 next passes into a third plenum chamber 78 formed by the conical rotor element 54, the tubular shaft 56, and a rotor support element 80 connected to the downstream end of one of the turbine discs 42. Air flowing from the plenum chamber 70 into the plenum chamber 78 does so through one or more airholes 82 positioned in the conical rotor element 54.
  • the cooling air 76 next flows between the turbine discs 42, thus cooling the outer surfaces thereof.
  • the air continues to flow through an opening 84 fonned in the annular torque member 46 and thus into a chamber 88 formed by the annular torque member 46 and the turbine heat shield 44.
  • the cooling air is then delivered to the interior portions of the hollow turbine blades 45 by means of radial passageways 88 extending through the turbine discs 42.
  • the cooling air enters the radial passageways 88 through holes 90 formed in the turbine heat shield 44. After this cooling air has performed its function of cooling the turbine blades, it is allowed to mix with the highenergy gas stream 20 prior to its flowing through the exhaust nozzle (not shown) as is well known in the art.
  • the principal object of the present invention is to provide a simple. and reliable means for varying the amount of cooling flow which is delivered to the turbine blades.
  • a plurality of small jet nozzles 92 are positioned within the passageway 64 along the inner wall of one of the separated disc portions 62.
  • the jet nozzles 92 are positioned along the separated disc portion 62 in such a manner that outlets 94 associated with each jet nozzle 92 are adapted to deliver pressurized air in a direction opposite the direction of rotation of the separated disc portion 62.
  • Each jet nozzle 92 is connected with a source of pressurized air by means of a tube 96 which is connected to the outside of the separated disc portion 62 as clearly shown in FIGS. 2 and 3.
  • the tubing 96 is, in turn, connected to a control valve 98, which may be attached to the tubular shaft 56 for rotation therewith.
  • the control valve 98 varies the amount of pressurized air delivered to the small jet nozzles 92 as will presently be discussed.
  • the control valve 98 is connected to a source of pressurized air by a second tubing 100.
  • a second tubing 100 is connected to the compressor rotor stage which includes the separated disc portions 62.
  • an air passageway consisting of one or more holes 102 is provided in the circumference of one of the downstream compressor discs 26. Should multiple holes 102 be provided, the control flow therefrom could be manifolded thereby balancing circumferential extraction pressure. (As is generally known in the art, the pressure of the gas stream increases the air flows downstream in an axial flow compressor.
  • air from the tubing 96 could be delivered through a manifold formed within one of the separated disc portions 62 for delivery to the jet nozzles 92 or, in the alternative, each radial row of jet nozzles 92 could be provided with a separate tube 96 and control valve 98.
  • the compressor discs 26 and thus the separated disc portions 62 form part of the compressor rotor. That is, the compressor discs 26 rotate in order to provide rotation of the compressor blades 28. Because of this rotation, the cooling air flowing through the cooling passageway 64 tends to form a natural vortex within the passageway 64. In order to overcome this vortex, prior art devices were equipped with radially extending paddles or vanes which prevented the formation of the vortex. The present device not only eliminates the need for the radial vanes, but also utilizes the natural vortex to control the amount of cooling flow which flows through the passageway 64 and thus to control the amount of cooling air which is delivered to the turbine discs 42 and to the interiors of the turbine blades 45, as required.
  • Modulation of the amount of cooling air delivered to the turbine blades is accomplished in the following manner.
  • control air flows through the airholes 102, through the tubing 100, through the control valve 98, through the tubing 96, and thus out of the small jet nozzles 92.
  • an aerodynamic bafile is fonned within the cooling passageway 64 to preclude formation of the natural vortex in much the same manner as radially extending vanes would do.
  • the number of the jet nozzles 92, their radial position, and the magnitude of the angle 9 (FIG. 4) would depend upon the design speed and the size of the separated disc portions 62. The number and position would best be determined by analytical and experimental development.
  • control valve 98 responds to centrifugal force such that maximum control flow occurs at maximum compressor rotor speed.
  • maximum control flow passes through the jet nozzles 92 thus assuring prevention of vortex buildup within the passageway 64. Cooling flow through the radial holes 66 is thus unimpeded by a vortex, and maximum flow through the radial holes 66 occurs.
  • the actual amount of cooling flow would, of course, depend on the size and number of the radial holes 66, the pressure drop across the holes 66, and normal losses associated with such a flow system.
  • the control valve 98 allows less control flow therethrough; and thus the flow through the jet nozzles 92 decreases.
  • a partial vortex forms within the passageway 64 and the pressure buildup associated with such a vortex results in decreased cooling flow through the radial holes 66. That is, as the vortex builds up within the passageway 64, the pressure drop across the radial holes 66 decreases and thus the cooling flow through the radial holes 66 decreases.
  • control valve 98 While any type valve could be utilized for the control valve 98, as previously discussed, a desirable type valve would be a centrifugally actuated type. In any case, the control valve 98 would be designed to fall open, so that in the event of failure, maximum cooling flow is provided to the turbine.
  • control valve or valves 98 could be positioned at the periphery of one of the compressor discs 26.
  • the control flow line(s) 100 runs along the outer portion of the disc 26 to a manifold 104 positioned either externally or internally of one of the separated disc portions 62'. The air then flows from the manifold 104 to the jet nozzles 92 to prevent formation of a vortex within the passageway 64 as described in conjunction with FIGS. 2 through 4.
  • a rotating member including a compressor rotor having a plurality of compressor blades projecting into said gas stream and a turbine rotor having a plurality of turbine blades projecting from turbine discs into said gas stream, said turbine blades having passageways for cooling purposes, and said compressor rotor including means for forming a vortex chamber,
  • first passageway means for providing a flow path from said gas stream to said vortex chamber
  • a gas turbine engine as recited in claim 1 further characterized in that said compressor rotor includes a plurality of discs for supporting said compressor blades, and at least one of said compressor discs comprises separated disc portions defining said vortex chamber.
  • a gas turbine engine as recited in claim 4 wherein said means for delivering control fluid comprising tubing fluidically connected to a source of pressurized fluid and valve means for varying the amount of pressurized fluid flowing through said tubing.
  • valve means are centrifugally controlled.
  • valve means permit maximum fluid passage through said tubing at maximum rotational speed of said rotating member.
  • a gas turbine engine as recited in claim 9 further characterized in that said rotating member includes a tubular shaft, a first conical rotor element connected on one end to said tubular shaft and on its opposite end to said compressor discs, and a second conical rotor element connected on one end to said tubular shaft and on its opposite end to said turbine rotor.
  • said second passageway means comprises one or more airholes positioned in said first conical rotor element and one or more airholes positioned within said second conical rotor element, whereby cooling air is capable of passing from said vortex chamber into said first plenum chamber, from said first plenum chamber into the said second plenum chamber, from said second chamber into said third plenum chamber, and from said third plenum chamber into said turbine blade passageways.
  • passageway means for providing a flow path from said vortex chamber to said turbine blade passageways
  • fluidic means for controlling the strength of a natural vortex formed within said vortex chamber due to rotation of said rotating members, said fluidic means comprising a plurality of jet nozzles positioned within said vortex chamber, and means for providing control fluid to said jet nozzles.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US60320A 1970-08-03 1970-08-03 Gas turbine engine cooling system incorporating a vortex shaft valve Expired - Lifetime US3632221A (en)

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BE (1) BE766536A (enrdf_load_stackoverflow)
CA (1) CA951921A (enrdf_load_stackoverflow)
DE (1) DE2121069A1 (enrdf_load_stackoverflow)
FR (1) FR2101178B1 (enrdf_load_stackoverflow)
GB (1) GB1345910A (enrdf_load_stackoverflow)

Cited By (25)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3897168A (en) * 1974-03-05 1975-07-29 Westinghouse Electric Corp Turbomachine extraction flow guide vanes
US4213738A (en) * 1978-02-21 1980-07-22 General Motors Corporation Cooling air control valve
US4217755A (en) * 1978-12-04 1980-08-19 General Motors Corporation Cooling air control valve
US4397471A (en) * 1981-09-02 1983-08-09 General Electric Company Rotary pressure seal structure and method for reducing thermal stresses therein
US4787820A (en) * 1987-01-14 1988-11-29 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Turbine plant compressor disc with centripetal accelerator for the induction of turbine cooling air
US4793772A (en) * 1986-11-14 1988-12-27 Mtu Motoren-Und Turbinen-Union Munchen Gmbh Method and apparatus for cooling a high pressure compressor of a gas turbine engine
US4808073A (en) * 1986-11-14 1989-02-28 Mtu Motoren- Und Turbinen- Union Munchen Gmbh Method and apparatus for cooling a high pressure compressor of a gas turbine engine
US4919590A (en) * 1987-07-18 1990-04-24 Rolls-Royce Plc Compressor and air bleed arrangement
US5316437A (en) * 1993-02-19 1994-05-31 General Electric Company Gas turbine engine structural frame assembly having a thermally actuated valve for modulating a flow of hot gases through the frame hub
US20030101730A1 (en) * 2001-12-05 2003-06-05 Stefan Hein Vortex reducer in the high-pressure compressor of a gas turbine
US6672072B1 (en) 1998-08-17 2004-01-06 General Electric Company Pressure boosted compressor cooling system
EP1566531A1 (de) * 2004-02-19 2005-08-24 Siemens Aktiengesellschaft Gasturbine mit einem gegen Auskühlen geschützten Verdichtergehäuse und Verfahren zum Betrieb einer Gasturbine
US20070258813A1 (en) * 2004-09-01 2007-11-08 Mtu Aero Engines Gmbh Rotor for a Power Plant
US20100326046A1 (en) * 2008-03-22 2010-12-30 Schirtzinger Gary A Valve system for a gas turbine engine
WO2013029010A1 (en) * 2011-08-24 2013-02-28 Qwtip Llc Water treatment system and method
US8623212B2 (en) 2010-08-24 2014-01-07 Qwtip Llc Water treatment and revitalization system and method
US9091173B2 (en) 2012-05-31 2015-07-28 United Technologies Corporation Turbine coolant supply system
US9469553B2 (en) 2011-08-24 2016-10-18 Qwtip, Llc Retrofit attachments for water treatment systems
US9605663B2 (en) 2010-08-24 2017-03-28 Qwtip Llc System and method for separating fluids and creating magnetic fields
US9617917B2 (en) 2013-07-31 2017-04-11 General Electric Company Flow control assembly and methods of assembling the same
US20170191419A1 (en) * 2015-12-30 2017-07-06 General Electric Company System and method of reducing post-shutdown engine temperatures
US9714176B2 (en) 2012-02-28 2017-07-25 Qwtip Llc Desalination and/or gas production system and method
US9878636B2 (en) 2012-02-29 2018-01-30 Qwtip Llc Levitation and distribution system and method
US10463993B2 (en) 2011-08-24 2019-11-05 Qwtip Llc Water treatment system and water
US10790723B2 (en) 2010-08-24 2020-09-29 Qwtip Llc Disk-pack turbine

Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2491549B1 (fr) * 1980-10-08 1985-07-05 Snecma Dispositif de refroidissement d'une turbine a gaz, par prelevement d'air au niveau du compresseur
EP0267478B1 (de) * 1986-11-14 1991-12-18 Mtu Motoren- Und Turbinen-Union MàœNchen Gmbh Gasturbinenstrahltriebwerk mit einem Hochdruckverdichter
DE19632038A1 (de) * 1996-08-08 1998-02-12 Asea Brown Boveri Vorrichtung zur Abscheidung von Staubpartikeln
US7331763B2 (en) * 2005-12-20 2008-02-19 General Electric Company Turbine disk
US10822100B2 (en) * 2017-06-26 2020-11-03 General Electric Company Hybrid electric propulsion system for an aircraft

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2618433A (en) * 1948-06-23 1952-11-18 Curtiss Wright Corp Means for bleeding air from compressors
US2787440A (en) * 1953-05-21 1957-04-02 Westinghouse Electric Corp Turbine apparatus
US2791091A (en) * 1950-05-15 1957-05-07 Gen Motors Corp Power plant cooling and thrust balancing systems
US2910268A (en) * 1951-10-10 1959-10-27 Rolls Royce Axial flow fluid machines
US3043561A (en) * 1958-12-29 1962-07-10 Gen Electric Turbine rotor ventilation system
NL6802034A (enrdf_load_stackoverflow) * 1968-01-10 1969-07-14

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2618433A (en) * 1948-06-23 1952-11-18 Curtiss Wright Corp Means for bleeding air from compressors
US2791091A (en) * 1950-05-15 1957-05-07 Gen Motors Corp Power plant cooling and thrust balancing systems
US2910268A (en) * 1951-10-10 1959-10-27 Rolls Royce Axial flow fluid machines
US2787440A (en) * 1953-05-21 1957-04-02 Westinghouse Electric Corp Turbine apparatus
US3043561A (en) * 1958-12-29 1962-07-10 Gen Electric Turbine rotor ventilation system
NL6802034A (enrdf_load_stackoverflow) * 1968-01-10 1969-07-14

Cited By (56)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3897168A (en) * 1974-03-05 1975-07-29 Westinghouse Electric Corp Turbomachine extraction flow guide vanes
US4213738A (en) * 1978-02-21 1980-07-22 General Motors Corporation Cooling air control valve
US4217755A (en) * 1978-12-04 1980-08-19 General Motors Corporation Cooling air control valve
US4397471A (en) * 1981-09-02 1983-08-09 General Electric Company Rotary pressure seal structure and method for reducing thermal stresses therein
US4793772A (en) * 1986-11-14 1988-12-27 Mtu Motoren-Und Turbinen-Union Munchen Gmbh Method and apparatus for cooling a high pressure compressor of a gas turbine engine
US4808073A (en) * 1986-11-14 1989-02-28 Mtu Motoren- Und Turbinen- Union Munchen Gmbh Method and apparatus for cooling a high pressure compressor of a gas turbine engine
US4787820A (en) * 1987-01-14 1988-11-29 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Turbine plant compressor disc with centripetal accelerator for the induction of turbine cooling air
US4919590A (en) * 1987-07-18 1990-04-24 Rolls-Royce Plc Compressor and air bleed arrangement
US5316437A (en) * 1993-02-19 1994-05-31 General Electric Company Gas turbine engine structural frame assembly having a thermally actuated valve for modulating a flow of hot gases through the frame hub
US6672072B1 (en) 1998-08-17 2004-01-06 General Electric Company Pressure boosted compressor cooling system
US7159402B2 (en) * 2001-12-05 2007-01-09 Rolls-Royce Deutschland Ltd & Co Kg Vortex reducer in the high-pressure compressor of a gas turbine
US20030101730A1 (en) * 2001-12-05 2003-06-05 Stefan Hein Vortex reducer in the high-pressure compressor of a gas turbine
US20070289286A1 (en) * 2004-02-18 2007-12-20 Holger Bauer Gas Turbine With a Compressor Housing Which is Protected Against Cooling Down and Method for Operating a Gas Turbine
US8336315B2 (en) 2004-02-18 2012-12-25 Siemens Aktiengesellschaft Gas turbine with a compressor housing which is protected against cooling down and method for operating a gas turbine
EP1566531A1 (de) * 2004-02-19 2005-08-24 Siemens Aktiengesellschaft Gasturbine mit einem gegen Auskühlen geschützten Verdichtergehäuse und Verfahren zum Betrieb einer Gasturbine
US7828514B2 (en) * 2004-09-01 2010-11-09 Mtu Aero Engines Gmbh Rotor for an engine
US20070258813A1 (en) * 2004-09-01 2007-11-08 Mtu Aero Engines Gmbh Rotor for a Power Plant
US8578716B2 (en) 2008-03-22 2013-11-12 United Technologies Corporation Valve system for a gas turbine engine
US20100326046A1 (en) * 2008-03-22 2010-12-30 Schirtzinger Gary A Valve system for a gas turbine engine
US8636910B2 (en) 2010-08-24 2014-01-28 Qwtip Llc Water treatment and revitalization system and method
US8623212B2 (en) 2010-08-24 2014-01-07 Qwtip Llc Water treatment and revitalization system and method
US12392330B2 (en) 2010-08-24 2025-08-19 Qwtip Llc Waveform disks and a system using the waveform disks
US12345242B2 (en) 2010-08-24 2025-07-01 Qwtip, Llc Fluid processing system with a disk-pack turbine
US12270387B2 (en) 2010-08-24 2025-04-08 Qwtip Llc Waveform disks for use in a disk-pack turbine
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Also Published As

Publication number Publication date
FR2101178A1 (enrdf_load_stackoverflow) 1972-03-31
BE766536A (fr) 1971-09-16
FR2101178B1 (enrdf_load_stackoverflow) 1974-04-05
DE2121069A1 (de) 1972-02-10
CA951921A (en) 1974-07-30
GB1345910A (en) 1974-02-06

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