GB2027811A - A gas turbine engine having means for bleeding compressor air - Google Patents

A gas turbine engine having means for bleeding compressor air Download PDF

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Publication number
GB2027811A
GB2027811A GB7927363A GB7927363A GB2027811A GB 2027811 A GB2027811 A GB 2027811A GB 7927363 A GB7927363 A GB 7927363A GB 7927363 A GB7927363 A GB 7927363A GB 2027811 A GB2027811 A GB 2027811A
Authority
GB
United Kingdom
Prior art keywords
compressor
flow
air
gas turbine
guide vane
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.)
Granted
Application number
GB7927363A
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GB2027811B (en
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
Publication of GB2027811A publication Critical patent/GB2027811A/en
Application granted granted Critical
Publication of GB2027811B publication Critical patent/GB2027811B/en
Expired legal-status Critical Current

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Classifications

    • 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
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F02C6/04—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output
    • F02C6/06—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
    • F02C6/08—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
    • 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
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/10—Final actuators
    • F01D17/12—Final actuators arranged in stator parts
    • F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
    • F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
    • F01D17/162—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
    • 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
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/56—Fluid-guiding means, e.g. diffusers adjustable
    • F04D29/563—Fluid-guiding means, e.g. diffusers adjustable specially adapted for elastic fluid pumps

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

A gas turbine engine comprises an axial-flow compressor having at least one variable guide vane assembly (2), the tips of the guide vanes 3 being disposed adjacent the surface 13 of the rotor, the guide vanes being hollow and being pivotably mounted in the outer casing of the compressor and being arranged so that in use compressor air may be bled from adjacent the rotor surface 13 through the guide vanes (3) and discharged through ducting (7, 15, 16) for subsequent use elsewhere. <IMAGE>

Description

SPECIFICATION A gas turbine engine having means for bleeding compressor air This invention relates to a gas turbine engine having at least one variable axial-flow compressor guide vane cascase and means for bleeding compressor air.
Axial-flow compressors and combined axialflow centrifugal compressors are frequently fitted with variable guide vane cascades in the front stages to adapt them for service at heavily fluctuating operating condition. It is especially with compressors in gas turbine engines that air must be bled from this compressor area--coming as it does with variable guide vane cascades-for subsequent use both internaliy and externally of the engine.
It is known to bleed the air from the compressor area fitted with variable guide vane cascades from the radially outer portion of the flow duct, normally via holes in the casing wall leading to a collector chamber and from there to a duct. While such air bleeding provisions are advantageous in their simplicity, there is the disadvantage that the centrifugal action of the rotor wheels upstream of the bleed point increases the foreign-body content (e.g. dust, sand, water, hail) in the outer flow portion, which may cause trouble during subsequent use. Furthermore, bleeding the air at the radially-outer portion of the flow duct will neither affect nor eliminate the boundary layer at the hub with its considerable impact on efficiency and operating action of the compressor.
An object of the invention is substantially to overcome the above disadvantages in a simple manner and substantially to eliminate the boundary layer adjacent the hub at the radially-inner wall of the flow duct.
The invention provides a gas turbine engine comprising at least one variable axial-flow compressor guide vane cascade having a hub arranged in a flow duct and a plurality of hollow guide vanes pivotably mounted in an outer casing of the compressor or engine only and being arranged so that in use compressor air may be bled from adjacent the hub through the hollow guide vanes and via ducting, for subsequent use elsewhere.
Preferably each guide vane has a turntable and pivot pin through which the bled air flows from the hollow guide vane.
With regard to gas turbine engine guide vanes pivotally supported in the outer casing of the compressor of the engine reference is here made to application DE-PS 10 33 677.
An embodiment of the invention will now be described with reference to the accompanying drawing, which is an elevation view of a combined axial-flow centrifugal compressor of a gas turbine engine.
In the drawing the final axial-flow stage of a compressor has rotor blades 1, and an associated variable guide vane cascade 2 has guide vanes 3. A centrifugal compressor 4 rotates with the portions of the axial-flow section of the compressor.
In this arrangement the aero foils of the guide vanes 3 are a fabricated construction brazed or welded to a cup-shaped turntable 5 likewise made of deep drawn sheet. The assembly of aero foil and turntable 5 is in turn brazed or welded to an associated pivot pin 6 having in its radially-inner portion a cylindrical cavity 7 which communicates with an inner cavity 9 of the associted guide vane 3 through a further cavity 8 of the turntable 5.
The radially-outer, solid portion of the pivot pin 6 serves the function of positioning the guide vane 6 relative to compressor outer casing 1 0. Sealing members 11 and 1 2 are inserted between the pivot pin 6 and adjacent portions of the compressor outer casing 1 0.
The axial-flow section of the compressor has a flow duct 14 having an inner peripheral surface 1 3. Air is bled from the duct 1 4 through the ports of the guide vanes 3 facing the inner surface 13, into the respective guide vane cavities 9, and from there the air flows through to the respective cavities 8 and 7 into drilled passages 1 5 connected with the cavities 7. The passages 1 5 are arranged at right angles to the axis of rotation D of the guide vanes 3. These drilled passages 1 5 are followed axially by casing passageways 1 6 issuing into a collector chamber 1 7 arranged between portions of the compressor outer casing 10 and a cover plate 1 8 of the centrifugal compressor.The bled air is ducted from the collector chamber 1 7 to a consumer through a duct (not shown).
This above-described construction ensures that the bled air is relatively free from foreign matter. Furthermore, the boundary layer at the surface 1 3 is syphoned away at least partially and this will benefit the downstream areas of the compressor in terms of efficiency and operating action.
The gas turbine engine of the present invention may have a straight axial-flow compressor, having one or several variable guide vane cascades, in which the air is bled at one or several stages of the axial-flow compressor.
1. A gas turbine engine comprising at least one variable axial-flow compressor guide vane cascade having a hub arranged in a flow duct and a plurality of hollow guide vanes pivotably mounted in an outer casing of the compressor or engine only and being arranged so that in use compressor air may be bled from adjacent the hub through the hollow guide vanes and via ducting for subsequent use elsewhere.
2. A gas turbine engine as claimed in claim 1, wherein each guide vane has a
**WARNING** end of DESC field may overlap start of CLMS **.

Claims (5)

**WARNING** start of CLMS field may overlap end of DESC **. SPECIFICATION A gas turbine engine having means for bleeding compressor air This invention relates to a gas turbine engine having at least one variable axial-flow compressor guide vane cascase and means for bleeding compressor air. Axial-flow compressors and combined axialflow centrifugal compressors are frequently fitted with variable guide vane cascades in the front stages to adapt them for service at heavily fluctuating operating condition. It is especially with compressors in gas turbine engines that air must be bled from this compressor area--coming as it does with variable guide vane cascades-for subsequent use both internaliy and externally of the engine. It is known to bleed the air from the compressor area fitted with variable guide vane cascades from the radially outer portion of the flow duct, normally via holes in the casing wall leading to a collector chamber and from there to a duct. While such air bleeding provisions are advantageous in their simplicity, there is the disadvantage that the centrifugal action of the rotor wheels upstream of the bleed point increases the foreign-body content (e.g. dust, sand, water, hail) in the outer flow portion, which may cause trouble during subsequent use. Furthermore, bleeding the air at the radially-outer portion of the flow duct will neither affect nor eliminate the boundary layer at the hub with its considerable impact on efficiency and operating action of the compressor. An object of the invention is substantially to overcome the above disadvantages in a simple manner and substantially to eliminate the boundary layer adjacent the hub at the radially-inner wall of the flow duct. The invention provides a gas turbine engine comprising at least one variable axial-flow compressor guide vane cascade having a hub arranged in a flow duct and a plurality of hollow guide vanes pivotably mounted in an outer casing of the compressor or engine only and being arranged so that in use compressor air may be bled from adjacent the hub through the hollow guide vanes and via ducting, for subsequent use elsewhere. Preferably each guide vane has a turntable and pivot pin through which the bled air flows from the hollow guide vane. With regard to gas turbine engine guide vanes pivotally supported in the outer casing of the compressor of the engine reference is here made to application DE-PS 10 33 677. An embodiment of the invention will now be described with reference to the accompanying drawing, which is an elevation view of a combined axial-flow centrifugal compressor of a gas turbine engine. In the drawing the final axial-flow stage of a compressor has rotor blades 1, and an associated variable guide vane cascade 2 has guide vanes 3. A centrifugal compressor 4 rotates with the portions of the axial-flow section of the compressor. In this arrangement the aero foils of the guide vanes 3 are a fabricated construction brazed or welded to a cup-shaped turntable 5 likewise made of deep drawn sheet. The assembly of aero foil and turntable 5 is in turn brazed or welded to an associated pivot pin 6 having in its radially-inner portion a cylindrical cavity 7 which communicates with an inner cavity 9 of the associted guide vane 3 through a further cavity 8 of the turntable 5. The radially-outer, solid portion of the pivot pin 6 serves the function of positioning the guide vane 6 relative to compressor outer casing 1 0. Sealing members 11 and 1 2 are inserted between the pivot pin 6 and adjacent portions of the compressor outer casing 1 0. The axial-flow section of the compressor has a flow duct 14 having an inner peripheral surface 1 3. Air is bled from the duct 1 4 through the ports of the guide vanes 3 facing the inner surface 13, into the respective guide vane cavities 9, and from there the air flows through to the respective cavities 8 and 7 into drilled passages 1 5 connected with the cavities 7. The passages 1 5 are arranged at right angles to the axis of rotation D of the guide vanes 3. These drilled passages 1 5 are followed axially by casing passageways 1 6 issuing into a collector chamber 1 7 arranged between portions of the compressor outer casing 10 and a cover plate 1 8 of the centrifugal compressor.The bled air is ducted from the collector chamber 1 7 to a consumer through a duct (not shown). This above-described construction ensures that the bled air is relatively free from foreign matter. Furthermore, the boundary layer at the surface 1 3 is syphoned away at least partially and this will benefit the downstream areas of the compressor in terms of efficiency and operating action. The gas turbine engine of the present invention may have a straight axial-flow compressor, having one or several variable guide vane cascades, in which the air is bled at one or several stages of the axial-flow compressor. CLAIMS
1. A gas turbine engine comprising at least one variable axial-flow compressor guide vane cascade having a hub arranged in a flow duct and a plurality of hollow guide vanes pivotably mounted in an outer casing of the compressor or engine only and being arranged so that in use compressor air may be bled from adjacent the hub through the hollow guide vanes and via ducting for subsequent use elsewhere.
2. A gas turbine engine as claimed in claim 1, wherein each guide vane has a turntable and pivot pin through which the bled air flows from the hollow guide vane.
3. A gas turbine engine as claimed in claim 2, wherein each pivot pin is provided with at least one drilled hole which extends at right angles to the axis of rotation of the pivot pin and connects a cavity of the pivot pin with a passageway in the casing, the passageway issuing into a collector chamber enclosed by casing portions and communicating with means for conveying the bled air for subsequent use elsewhere.
4. A gas turbine engine as claimed in claim 1, 2 or 3 and comprising a combined axial-flow centrifugal compressor.
5. A gas turbine engine substantially as herein described with reference to the accompanying drawing.
GB7927363A 1978-08-09 1979-08-06 Gasturbine engine having means for bleeding Expired GB2027811B (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE19782834822 DE2834822C2 (en) 1978-08-09 1978-08-09 Device for the extraction of compressor air in gas turbine engines

Publications (2)

Publication Number Publication Date
GB2027811A true GB2027811A (en) 1980-02-27
GB2027811B GB2027811B (en) 1982-09-29

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Family Applications (1)

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GB7927363A Expired GB2027811B (en) 1978-08-09 1979-08-06 Gasturbine engine having means for bleeding

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DE (1) DE2834822C2 (en)
FR (1) FR2433106A1 (en)
GB (1) GB2027811B (en)

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0395498A1 (en) * 1989-04-26 1990-10-31 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Variable inlet guide vane with a built-in turntable
US5025629A (en) * 1989-03-20 1991-06-25 Woollenweber William E High pressure ratio turbocharger
EP1388642A2 (en) * 2002-08-06 2004-02-11 AVIO S.p.A. Variable-geometry turbine stator blade, particularly for aircraft engines
WO2007115446A1 (en) * 2006-04-07 2007-10-18 Changzhe Liu A cascade radial-flow compressor
EP1998025A1 (en) * 2007-05-30 2008-12-03 Snecma Compressor with air re-injection
EP2058524A1 (en) * 2007-11-12 2009-05-13 Siemens Aktiengesellschaft Air bleed compressor with variable guide vanes
US7845901B2 (en) * 2005-01-14 2010-12-07 Snecma Air bleed device on a machine stator pivoting blade
US20110110773A1 (en) * 2008-06-25 2011-05-12 Snecma Turbomachine compressor
CN106687666A (en) * 2014-09-12 2017-05-17 通用电气公司 Axi-centrifugal compressor with variable outlet guide vanes
CN113074126A (en) * 2021-04-18 2021-07-06 上海尚实能源科技有限公司 Two-stage axial flow compressor
CN114144573A (en) * 2019-07-24 2022-03-04 赛峰航空发动机公司 Turbomachine rectifier stage with cooling air leakage channels having variable cross-section according to the orientation of the blades

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19907907A1 (en) 1999-02-24 2000-08-31 Abb Alstom Power Ch Ag Multi-stage turbo compressor
DE10355240A1 (en) * 2003-11-26 2005-07-07 Rolls-Royce Deutschland Ltd & Co Kg Fluid flow machine with fluid removal
US10794272B2 (en) 2018-02-19 2020-10-06 General Electric Company Axial and centrifugal compressor

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2720356A (en) * 1952-06-12 1955-10-11 John R Erwin Continuous boundary layer control in compressors
GB1075958A (en) * 1966-04-29 1967-07-19 Rolls Royce Gas turbine engine
FR2085470B1 (en) * 1970-04-23 1974-05-03 Snecma

Cited By (25)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5025629A (en) * 1989-03-20 1991-06-25 Woollenweber William E High pressure ratio turbocharger
EP0395498A1 (en) * 1989-04-26 1990-10-31 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Variable inlet guide vane with a built-in turntable
FR2646467A1 (en) * 1989-04-26 1990-11-02 Snecma STATOR VARIABLE STATOR VANE WITH REPLACED CUP
US5039277A (en) * 1989-04-26 1991-08-13 Societe National D'etude Et De Construction De Moteurs D'aviation Variable stator vane with separate guide disk
EP1388642A2 (en) * 2002-08-06 2004-02-11 AVIO S.p.A. Variable-geometry turbine stator blade, particularly for aircraft engines
EP1388642A3 (en) * 2002-08-06 2004-10-06 AVIO S.p.A. Variable-geometry turbine stator blade, particularly for aircraft engines
US6913440B2 (en) 2002-08-06 2005-07-05 Avio S.P.A. Variable-geometry turbine stator blade, particularly for aircraft engines
US7845901B2 (en) * 2005-01-14 2010-12-07 Snecma Air bleed device on a machine stator pivoting blade
WO2007115446A1 (en) * 2006-04-07 2007-10-18 Changzhe Liu A cascade radial-flow compressor
EP1998025A1 (en) * 2007-05-30 2008-12-03 Snecma Compressor with air re-injection
FR2916815A1 (en) * 2007-05-30 2008-12-05 Snecma Sa AIR REINJECTION COMPRESSOR
US8182209B2 (en) 2007-05-30 2012-05-22 Snecma Air reinjection compressor
EP2058524A1 (en) * 2007-11-12 2009-05-13 Siemens Aktiengesellschaft Air bleed compressor with variable guide vanes
WO2009062793A1 (en) * 2007-11-12 2009-05-22 Siemens Aktiengesellschaft Air bleed in compressor with variable guide vanes
US20110110773A1 (en) * 2008-06-25 2011-05-12 Snecma Turbomachine compressor
US8974175B2 (en) * 2008-06-25 2015-03-10 Snecma Turbomachine compressor
CN106687666A (en) * 2014-09-12 2017-05-17 通用电气公司 Axi-centrifugal compressor with variable outlet guide vanes
US20170248156A1 (en) * 2014-09-12 2017-08-31 General Electric Company Axi-centrifugal compressor with variable outlet guide vanes
JP2017527733A (en) * 2014-09-12 2017-09-21 ゼネラル・エレクトリック・カンパニイ Axial flow-centrifugal compressor with variable output guide vanes
CN106687666B (en) * 2014-09-12 2019-07-09 通用电气公司 Axial-flow centrifugal compressor with variable export orientation wheel blade
US10704563B2 (en) * 2014-09-12 2020-07-07 General Electric Company Axi-centrifugal compressor with variable outlet guide vanes
US11448235B2 (en) 2014-09-12 2022-09-20 General Electric Company Axi-centrifugal compressor with variable outlet guide vanes
CN114144573A (en) * 2019-07-24 2022-03-04 赛峰航空发动机公司 Turbomachine rectifier stage with cooling air leakage channels having variable cross-section according to the orientation of the blades
US11952950B2 (en) * 2019-07-24 2024-04-09 Safran Aircraft Engines Sas Axial turbine engine, and rectifier stage with variable orientation vanes for an axial turbine engine
CN113074126A (en) * 2021-04-18 2021-07-06 上海尚实能源科技有限公司 Two-stage axial flow compressor

Also Published As

Publication number Publication date
GB2027811B (en) 1982-09-29
DE2834822C2 (en) 1981-09-17
FR2433106A1 (en) 1980-03-07
DE2834822A1 (en) 1980-02-14
FR2433106B3 (en) 1981-04-17

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PCNP Patent ceased through non-payment of renewal fee