US3898799A - Device for bleeding-off compressor air in turbine jet engine - Google Patents

Device for bleeding-off compressor air in turbine jet engine Download PDF

Info

Publication number
US3898799A
US3898799A US401544A US40154473A US3898799A US 3898799 A US3898799 A US 3898799A US 401544 A US401544 A US 401544A US 40154473 A US40154473 A US 40154473A US 3898799 A US3898799 A US 3898799A
Authority
US
United States
Prior art keywords
ring
actuating
internal ring
rings
external
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
US401544A
Other languages
English (en)
Inventor
Wolfgang Pollert
Eckhard Kraft
Gregor Pennig
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.)
MTU Aero Engines GmbH
Original Assignee
MTU Motoren und Turbinen Union Muenchen GmbH
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 MTU Motoren und Turbinen Union Muenchen GmbH filed Critical MTU Motoren und Turbinen Union Muenchen GmbH
Application granted granted Critical
Publication of US3898799A publication Critical patent/US3898799A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
    • F04D27/02—Surge control
    • F04D27/0207—Surge control by bleeding, bypassing or recycling fluids
    • F04D27/0215—Arrangements therefor, e.g. bleed or by-pass valves
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
    • F04D27/02—Surge control
    • F04D27/0207—Surge control by bleeding, bypassing or recycling fluids
    • F04D27/023—Details or means for fluid extraction

Definitions

  • No: 401,544 ranged compressors.
  • the apparatus includes a first annular chamber communicated directly with compressor air from one of the compressors and a second annular chamber communicated directly with atmo- [30] Foreign Application Priority Data Sept. 27. 1972 Germany alone/2017, etc.
  • Actuating means including a radially extending actuating shaft and a pair of pivotally connected levers intercon necting the internal ring and the actuating shaft are provided for moving the internal ring with respect to References cued the external ring.
  • the internal ring is provided with UNITED STATES PATENTS guide slots and the external ring is provided with guide rollers for guiding relative movement of the internal and external rings such that the internal ring simultaneously moves both circumferentially and axially during adjustments thereof.
  • a condition may occur where, e.g., the speed and consequently also the flow supplied by a first compressor is excessive relative to the speed of a subsequent second compressor.
  • the present invention contemplates providing a relatively simple, robust and reliable device for compressor air bleedoff.
  • this device is easy to actuate, permits the bleed-off of a relatively large amount of compressor air and provides a tight seal at the compressor bleed point(s) when no bleed-off is desired.
  • control of the bleed air flow is provided for adaptation to the prevailing operating conditions.
  • the present invention further contemplates providing an internal ring adjustable both axially and tangentially within a first annular chamber supplied with compressor air actuation of this internal ring uncovers a plurality of openings of an associated external ring and thus allows compressor air to flow into a second annular chamber located above or radially outwardly of the external ring and communicating with the bypass duct of the engine or with atmosphere.
  • the present invention avoids the disadvantages of contemplated arrangements with only one and consequently relatively large compressor bleed air opening to be opened or closed so that the bleed sectional area on the one hand is relatively limited and on the other hand an uneven load distribution occurs over the circumference of the compressor casing section concerned.
  • the existence of only one and thus relatively large compressor bleed air opening results in a relatively large differential pressure between the chambers to be sealed and consequently in a relatively large actuating force for opening or closing the compressor bleed air opening.
  • the motions of a ring, movable both tangentially and radially are superimposed to one another for opening an annular slot in the compressor casing section for compressor air bleed-off.
  • this ring When moved as described this ring is highly susceptible to tilting within the compressor casing, unless this ring is moved in an absolutely synchronous motion over its complete circumference.
  • the radial motion occurring when the sliding ring is actuated generally excludes the employment of this arrangement for compressor casings of relatively small diameter. In addition, this makes the requirement of an absolutely tight seal at the bleed opening questionable.
  • the present invention also eliminates the disadvantages of this type of arrangement.
  • FIG. 1 is a schematic and fragmentary sectional view of the top half of a turbine jet engine taken along the longitudinal center axis which illustrates the association of the device according to the invention relative to said engine;
  • FIG. 2 is a longitudinal section of the device according to the invention arranged between a first and a second compressor, shown partially and within the bypass duct area only;
  • FIG. 3 is a view ofa section of the device in direction A of FIG. 2, and
  • FIG. 4 is a sectional view along line III-III of FIG. 3.
  • the turbine jet engine shown schematically in FIG. 1 comprises low pressure compressor 1, intermediate pressure compressor 2, high pressure compressor 3, annular combustion chamber 5 coaxial in this case, for example, with the longitudinal center axis 4, followed by high pressure turbine 6, intermediate pressure turbine 7 and low pressure turbine 8 arranged one behind the other.
  • Low pressure compressor 1 and low pressure turbine 8 are connected by a common shaft 9.
  • Intermediate pressure compressor 2 and intermediate pressure turbine 7 are connected to each other by means of a hollow shaft 10 coaxially enclosing shaft 9.
  • Another hollow shaft 11 enclosing hollow shaft 10 connects high pressure compressor 3 with high pressure turbine 6.
  • bypass air flow supplied by low pressure compressor 1 enters bypass duct 12 arranged coaxially with the longitudinal center axis 4 of the engine and after joining the engine exhaust gases it flows into afterburner jet pipe 13'.
  • intermediate pressure compressor 2 and high pressure compressor 3 for example, the device for compressor air bleed-off explained in more detail in conjunction with FIGS. 2 and 3 is located on compressor section 13, arranged coaxially with the center longitudinal axis 4.
  • a portion of the air supplied by intermediate pressure compressor 2 can be bled off by the apparatus of the present invention either into bypass duct 12 of the engine (arrow 14) or via hollow struts 16, passing through bypass duct 12, to atmosphere as indicated by dotted arrow 15.
  • intermediate pressure compressor 2 In a turbine jet engine according to FIG. intermediate pressure compressor 2, generally indicated by guide vanes and rotor blades 17 and 18, respectively, in FIG. 2, supplies compressed air in the direction of arrows K into a first annular chamber 19.
  • an internal ring 20 Within this annular chamber 19 an internal ring 20, adjustable axially and tangentially is provided, which when actuated uncovers openings 21 of associated external ring 22 and thus allows compressor air to flow into a second annular chamber 23, arrangedabove or radially outward of external ring 22 and communicating via additional openings 24 with bypass duct 12 of the jet engine.
  • This chamber 23 can alternatively be communicated with atmosphere by an arrangement such as l5, 16 of FIG. 1.
  • Ring 20 and the seal rings 25, 26, are slidable in circumferential (tangential) and axial (parallel to center line 4) to open positions where openings 21 of external ring 22 are in communication with chamber 19 (open position would be with ring 20 and seal ring 25 and 26 moved leftward of position in FIG. 2).
  • External ring 22 is supported elastically in radial direction at its free end (left end in FIG. 2) and is provided with a seal relative to the first annular chamber 19 through seal ring 27.
  • the right end of ring 22 is threadedly connected with a radially inwardly projecting rib of the housing.
  • Inner ring 20 is actuated by means of actuating shaft 28, passing through bypass duct 12.
  • the torque or rotational movement of the shaft 28 is transmitted to inner ring 20 by at least one actuating lever 29 and intermediate lever 30 pivotably connected to lever 29.
  • actuating lever 29 and intermediate lever 30 pivotably connected to lever 29.
  • spherical bearings 31, 32 are provided.
  • ring 20 For insuring a combined motion of internal ring 20 in axial and circumferential direction, ring 20 is provided with guide slots 33 (FIG. 3), inclined in accordance with the desired direction of rotation. A plurality of said guide slots, e.g. three or six are preferably spaced uniformly over the circumference of ring 20.
  • FIG. 3 includes a dashline showing the internal ring 20 in blocking relationship to the openings 21 in external ring 22. As can be seen in FIG. 3, a large number of openings 21 are provided on external ring 22.
  • the actuating shaft 28 is forcefully rotated by means not shown in the desired positions corresponding to the desired compressed air bleed-off flow.
  • the control of this actuating shaft 28 could be responsive to sensed pressures within the engine so as to assure that undesirable compressor surge is avoided.
  • a further preferred embodiment of the present invention includes thread-type spiral shaped guide slots in place of slots 33. (see FIG. 3)
  • Apparatus for bleeding off compressor air in a turbine engine of the type having a longitudinal center line and a plurality of mechanically independent compressors', and apparatus comprising:
  • an external ring positioned downstream of said internal ring with respect to air flow through said turbine engine, positioned adjacent and externally of said internal ring with respect to said center line, and including a plurality of openings
  • bypass means a second chamber communicated with bypass means of the engine which bypass means accommodates air flow therein in bypassing relationship to an engine combustion chamber arranged downstream of said compressors,
  • each of said chambers are annular chambers which extend around said longitudinal centerline. of the engine which center line forms the axes of rotation for compressor wheels, and wherein each of said rings are annular rings which extend about said centerline.
  • Apparatus according to claim 2 wherein said actuating means includes means for moving said internal ring in both axial and circumferential directions with respect to said centerline.
  • Apparatus according to claim 3 further comprising two annular seal rings on said internal ring which sealingly engage said external ring to form a third annular chamber between said internal and external rings which is sealed relative to said first chamber when said internal ring is in said closed positions.
  • said actuating means includes: a rotatable actuating shaft extending through a bypass duct of said engine in a radial direction with respect to said centerline, an actuating lever movable with said actuating shaft and an intermediate lever pivotally connected to both said actuating lever and said internal ring.
  • said actuating means includes: a rotatable actuating shaft extending through a bypass duct of said engine in a radial direction with respect to said centerline, an actuating lever movable with said actuating shaft and an intermediate lever pivotally connected to both said actuating lever and said internal ring.
  • said actuating means includes: a rotatable actuating shaft extending through a bypass duct of said engine in a radial direction with respect to said centerline, an actuating lever movable with said actuating shaft and an intermediate lever pivotally connected to both said actuating lever and said internal ring.
  • Apparatus according to claim 8 wherein said actuating and intermediate levers are connected to one another by a first spherical bearing, and wherein said intermediate lever and said internal ring are connected to one another by a second spherical bearing.
  • said actuating means includes: a plurality of circumferentially spaced guide slots in said internal ring which are inclined in accordance with the direction of movement of said internal ring, and rollers mounted on said external ring and engageable in said guide slots for accomodating and guiding movement of said internal ring with respect to said external ring over a predetermined path.
  • said actuating means includes: a plurality of circumferentially spaced guide slots in said internal ring which are inclined in accordance with the direction of movement of said internal ring, and rollers mounted on said external ring and engageable in said guide slots for accommodating and guiding movement of said internal ring with respect to said external ring over a predetermined path.
  • said actuating means includes: a plurality of circumferentially spaced guide slots in said internal ring which are inclined in accordance with the direction of movement of said internal ring. and rollers mounted on said external ring and engageable in said guide slots for accommodating and guiding movement of said internal ring with respect to said external ring over a predetermined path.
  • said actuating means includes: a plurality of circumferentially spaced guide slots in said internal ring which are inclined in accordance with the direction of movement of said internal ring. and rollers mounted on said external ring and engageable in said guide slots for accommodating and guiding movement of said internal ring with respect to said external ring over a predetermined path.
  • said actuating means includes: a plurality of circumferentially spaced guide slots in said internal ring which are inclined in accordance with the direction of movement of said internal ring, and rollers mounted on said external ring and engageable in said guide slots for accommodating and guiding movement of said internal ring with respect to said external ring over a predetermined path.
  • rollers are adjustable by means of eccentric pins interconnecting them with the external ring.
  • rollers are adjustable by means of eccentric pins interconnecting them with the external ring.
  • rollers are adjustable by means of eccentric pins interconnecting them with the external ring.
  • actuating means includes guide slots in one of said internal and external rings and rollers mounted in the other of said internal and external rings for engagement in said guide slots to forcibly guide relative movement of said external and internal rings.
  • each of said chambers are annular chambers which extend around said longitudinal centerline of the engine which center line forms the axes of rotation for the compressor wheels, and wherein each of said rings are annular rings which extend about said centerline.
  • actuating means includes means for moving said internal ring in both axial and circumferential directions with respect to said centerline.
  • said actuating means includes; a rotatable actuating shaft extending through a bypass duct of said engine in a radial direction with respect to said centerline, an actuating lever movable with said actuating shaft and an intermediate lever pivotally connected to both said actuating lever and said internal ring.
  • Apparatus according to claim 26 wherein said actuating and intermediate levers are connected to one another by a first spherical bearing, and wherein said intermediate lever and said internal ring are connected to one another by a second spherical bearing.
  • bypass means a second chamber communicated with bypass means of the engine which bypass means accommodates air flow therein in bypassing relationship to an engine combustion chamber arranged downstream of said compressor means,
  • actuating means includes guide slots in one of said first and second rings and rollers mounted in the other of said first and second rings for engagement in said guide slots to forcibly guide relative movement of said second and first rings.
  • Apparatus according to claim 35 further comprising two annular seal rings on said first ring which sealingly engage said second ring to form a third annular chamber between said first and second rings which is sealed relative to said first chamber when said first ring is in said closed positions.
  • actuating means includes: a rotatable actuating shaft extending through a bypass duct of said engine in a radial direction with respect to said centerline, an actuating lever movable with said actuating shaft and an intermediate lever pivotally connected to both said actuating lever and said first ring.
  • said actuating means includes: a rotatable actuating shaft extending through a bypass duct of said engine in a ra dial direction with respect to said centerline, an actuating lever movable with said actuating shaft and an intermediate lever pivotally connected to both said actuating lever and said first ring.
  • Apparatus according to claim 1 further comprising annular seal rings movable with and provided on said internal ring which seal rings engage said external ring.

Landscapes

  • Engineering & Computer Science (AREA)
  • Life Sciences & Earth Sciences (AREA)
  • Sustainable Development (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Taps Or Cocks (AREA)
  • Supercharger (AREA)
US401544A 1972-09-27 1973-09-27 Device for bleeding-off compressor air in turbine jet engine Expired - Lifetime US3898799A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE2247400A DE2247400C2 (de) 1972-09-27 1972-09-27 Vorrichtung zum Abblasen von verdichteter Luft aus einem Verdichter eines Gasturbinenstrahltriebwerks

Publications (1)

Publication Number Publication Date
US3898799A true US3898799A (en) 1975-08-12

Family

ID=5857538

Family Applications (1)

Application Number Title Priority Date Filing Date
US401544A Expired - Lifetime US3898799A (en) 1972-09-27 1973-09-27 Device for bleeding-off compressor air in turbine jet engine

Country Status (5)

Country Link
US (1) US3898799A (de)
DE (1) DE2247400C2 (de)
FR (1) FR2209044B1 (de)
GB (1) GB1445415A (de)
IT (1) IT993227B (de)

Cited By (26)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4055946A (en) * 1976-03-29 1977-11-01 United Technologies Corporation Twin-spool gas turbine power plant with means to spill compressor interstage airflow
US4086761A (en) * 1976-04-26 1978-05-02 The Boeing Company Stator bypass system for turbofan engine
US4232513A (en) * 1977-10-19 1980-11-11 Rolls-Royce Limited Pressure relief panel for aircraft powerplant
US4390318A (en) * 1977-09-10 1983-06-28 Mtu Motoren-Und Turbinen-Union Muenche Gmbh Apparatus for operating shut-off members in gas turbine engines, particularly in turbojet engines
US4463552A (en) * 1981-12-14 1984-08-07 United Technologies Corporation Combined surge bleed and dust removal system for a fan-jet engine
US4546605A (en) * 1983-12-16 1985-10-15 United Technologies Corporation Heat exchange system
US4698964A (en) * 1985-09-06 1987-10-13 The Boeing Company Automatic deflector for a jet engine bleed air exhaust system
EP0298015A2 (de) * 1987-06-29 1989-01-04 United Technologies Corporation Aufbau eines Leitradventils für eine Rotormaschine
US5044153A (en) * 1988-12-15 1991-09-03 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Turbojet compressor blow off valves with water collecting and discharge means
EP0487242A1 (de) * 1990-11-19 1992-05-27 General Electric Company Abblastruktur eines Verdichters
US5136840A (en) * 1982-09-30 1992-08-11 General Electric Company Gas turbine engine actuation system
US5177914A (en) * 1989-06-19 1993-01-12 Hilmer Elwyn P Vertical section building construction
US5209633A (en) * 1990-11-19 1993-05-11 General Electric Company High pressure compressor flowpath bleed valve extraction slot
US5211003A (en) * 1992-02-05 1993-05-18 General Electric Company Diffuser clean air bleed assembly
US5231825A (en) * 1990-04-09 1993-08-03 General Electric Company Method for compressor air extraction
DE4216033A1 (de) * 1992-03-12 1993-09-16 Bmw Rolls Royce Gmbh Leitschaufel-verstellvorrichtung
US5287697A (en) * 1992-01-02 1994-02-22 General Electric Company Variable area bypass injector seal
US5380151A (en) * 1993-10-13 1995-01-10 Pratt & Whitney Canada, Inc. Axially opening cylindrical bleed valve
US5477673A (en) * 1994-08-10 1995-12-26 Pratt & Whitney Canada Inc. Handling bleed valve
US6048171A (en) * 1997-09-09 2000-04-11 United Technologies Corporation Bleed valve system
US6122905A (en) * 1998-02-13 2000-09-26 Pratt & Whitney Canada Corp. Compressor bleed valve
EP0936357A3 (de) * 1998-02-13 2001-01-10 Pratt & Whitney Canada Corp. Verdichterabblasventil
US6398491B1 (en) * 1999-02-24 2002-06-04 Alstom (Switzerland) Ltd Multistage turbocompressor
US20040050071A1 (en) * 2002-09-13 2004-03-18 Bachelder Kenneth Alan Methods and apparatus for supporting variable bypass valve systems
US20110103949A1 (en) * 2009-11-05 2011-05-05 General Electric Company Extraction Cavity Wing Seal
US20130170966A1 (en) * 2012-01-04 2013-07-04 General Electric Company Turbine cooling system

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
IL59497A (en) * 1979-04-23 1984-08-31 Gen Electric Valve actuation system for use on a gas turbine engine

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3030006A (en) * 1958-05-27 1962-04-17 United Aircraft Corp Circumferential bleed valve
US3057541A (en) * 1958-06-03 1962-10-09 United Aircraft Corp Circumferential bleed valve
US3094270A (en) * 1958-08-05 1963-06-18 Rolls Royce Annular valve device
US3095010A (en) * 1959-04-30 1963-06-25 Rolls Royce Gas turbine engine having adjustable guide vanes
US3688504A (en) * 1970-11-27 1972-09-05 Gen Electric Bypass valve control
US3747341A (en) * 1971-01-02 1973-07-24 Dowty Rotol Ltd Fans
US3777489A (en) * 1972-06-01 1973-12-11 Gen Electric Combustor casing and concentric air bleed structure

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3030006A (en) * 1958-05-27 1962-04-17 United Aircraft Corp Circumferential bleed valve
US3057541A (en) * 1958-06-03 1962-10-09 United Aircraft Corp Circumferential bleed valve
US3094270A (en) * 1958-08-05 1963-06-18 Rolls Royce Annular valve device
US3095010A (en) * 1959-04-30 1963-06-25 Rolls Royce Gas turbine engine having adjustable guide vanes
US3688504A (en) * 1970-11-27 1972-09-05 Gen Electric Bypass valve control
US3747341A (en) * 1971-01-02 1973-07-24 Dowty Rotol Ltd Fans
US3777489A (en) * 1972-06-01 1973-12-11 Gen Electric Combustor casing and concentric air bleed structure

Cited By (35)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4055946A (en) * 1976-03-29 1977-11-01 United Technologies Corporation Twin-spool gas turbine power plant with means to spill compressor interstage airflow
US4086761A (en) * 1976-04-26 1978-05-02 The Boeing Company Stator bypass system for turbofan engine
US4390318A (en) * 1977-09-10 1983-06-28 Mtu Motoren-Und Turbinen-Union Muenche Gmbh Apparatus for operating shut-off members in gas turbine engines, particularly in turbojet engines
US4232513A (en) * 1977-10-19 1980-11-11 Rolls-Royce Limited Pressure relief panel for aircraft powerplant
US4463552A (en) * 1981-12-14 1984-08-07 United Technologies Corporation Combined surge bleed and dust removal system for a fan-jet engine
US5136840A (en) * 1982-09-30 1992-08-11 General Electric Company Gas turbine engine actuation system
US4546605A (en) * 1983-12-16 1985-10-15 United Technologies Corporation Heat exchange system
US4698964A (en) * 1985-09-06 1987-10-13 The Boeing Company Automatic deflector for a jet engine bleed air exhaust system
US4827713A (en) * 1987-06-29 1989-05-09 United Technologies Corporation Stator valve assembly for a rotary machine
EP0298015A3 (en) * 1987-06-29 1989-07-05 United Technologies Corporation Stator valve assembly for a rotary machine
EP0298015A2 (de) * 1987-06-29 1989-01-04 United Technologies Corporation Aufbau eines Leitradventils für eine Rotormaschine
US5044153A (en) * 1988-12-15 1991-09-03 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Turbojet compressor blow off valves with water collecting and discharge means
US5177914A (en) * 1989-06-19 1993-01-12 Hilmer Elwyn P Vertical section building construction
US5231825A (en) * 1990-04-09 1993-08-03 General Electric Company Method for compressor air extraction
US5209633A (en) * 1990-11-19 1993-05-11 General Electric Company High pressure compressor flowpath bleed valve extraction slot
EP0487242A1 (de) * 1990-11-19 1992-05-27 General Electric Company Abblastruktur eines Verdichters
US5287697A (en) * 1992-01-02 1994-02-22 General Electric Company Variable area bypass injector seal
US5343697A (en) * 1992-01-02 1994-09-06 General Electric Company Variable area bypass injector
US5211003A (en) * 1992-02-05 1993-05-18 General Electric Company Diffuser clean air bleed assembly
DE4216033A1 (de) * 1992-03-12 1993-09-16 Bmw Rolls Royce Gmbh Leitschaufel-verstellvorrichtung
US5380151A (en) * 1993-10-13 1995-01-10 Pratt & Whitney Canada, Inc. Axially opening cylindrical bleed valve
US5477673A (en) * 1994-08-10 1995-12-26 Pratt & Whitney Canada Inc. Handling bleed valve
US6048171A (en) * 1997-09-09 2000-04-11 United Technologies Corporation Bleed valve system
EP0936357A3 (de) * 1998-02-13 2001-01-10 Pratt & Whitney Canada Corp. Verdichterabblasventil
US6122905A (en) * 1998-02-13 2000-09-26 Pratt & Whitney Canada Corp. Compressor bleed valve
US6398491B1 (en) * 1999-02-24 2002-06-04 Alstom (Switzerland) Ltd Multistage turbocompressor
US20040050071A1 (en) * 2002-09-13 2004-03-18 Bachelder Kenneth Alan Methods and apparatus for supporting variable bypass valve systems
JP2004108364A (ja) * 2002-09-13 2004-04-08 General Electric Co <Ge> 可変バイパス弁システムを支持するための方法及び装置
US6742324B2 (en) * 2002-09-13 2004-06-01 General Electric Company Methods and apparatus for supporting variable bypass valve systems
EP1398464A3 (de) * 2002-09-13 2006-02-22 General Electric Company Halterung des Bypassventils einer Gasturbine
JP4641713B2 (ja) * 2002-09-13 2011-03-02 ゼネラル・エレクトリック・カンパニイ 可変バイパス弁システムを支持するための方法及び装置
US20110103949A1 (en) * 2009-11-05 2011-05-05 General Electric Company Extraction Cavity Wing Seal
US8388313B2 (en) * 2009-11-05 2013-03-05 General Electric Company Extraction cavity wing seal
US20130170966A1 (en) * 2012-01-04 2013-07-04 General Electric Company Turbine cooling system
CN103195490A (zh) * 2012-01-04 2013-07-10 通用电气公司 涡轮机冷却系统

Also Published As

Publication number Publication date
FR2209044B1 (de) 1980-11-07
IT993227B (it) 1975-09-30
DE2247400B1 (de) 1974-05-16
FR2209044A1 (de) 1974-06-28
GB1445415A (en) 1976-08-11
DE2247400C2 (de) 1975-01-16

Similar Documents

Publication Publication Date Title
US3898799A (en) Device for bleeding-off compressor air in turbine jet engine
US2698711A (en) Compressor air bleed closure
US3638428A (en) Bypass valve mechanism
US3869221A (en) Rotor wheel fan blade adjusting apparatus for turbojet engines and the like
US4785624A (en) Turbine engine blade variable cooling means
US4827713A (en) Stator valve assembly for a rotary machine
US2598176A (en) Sealing device
US5575616A (en) Turbine cooling flow modulation apparatus
EP1888881B1 (de) Turbine mit variabler geometrie
EP0654587A1 (de) Turbine mit variabler Einlassgeometrie
US2630673A (en) Cooling means for variable area nozzles
GB1602767A (en) Turbocharger control
GB2031069A (en) Turbine of exhaust gas turbo- charger
US20040016238A1 (en) Pneumatic compressor bleed valve
US4674951A (en) Ring structure and compressor blow-off arrangement comprising said ring
US2445837A (en) Air-cooled gas turbine
US3379366A (en) Contra-rotating compressors
CN110023592B (zh) 装配有排气系统的双涵道涡轮机
CN1006168B (zh) 蒸汽轮机高压端均压孔和汽封系统
US20060225432A1 (en) Supercharged open cycle gas turbine engine
US3030006A (en) Circumferential bleed valve
GB2319307A (en) Controlling cooling air in a gas turbine engine
US1986435A (en) Turbine engine
GB737473A (en) Turbines and like machines having adjustable guide blades
US4130989A (en) Automotive turbine engine