US3632223A - Turbine engine having multistage compressor with interstage bleed air system - Google Patents

Turbine engine having multistage compressor with interstage bleed air system Download PDF

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Publication number
US3632223A
US3632223A US862197A US3632223DA US3632223A US 3632223 A US3632223 A US 3632223A US 862197 A US862197 A US 862197A US 3632223D A US3632223D A US 3632223DA US 3632223 A US3632223 A US 3632223A
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United States
Prior art keywords
bleed
casing
flow path
compressor
passageway
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Expired - Lifetime
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US862197A
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English (en)
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Thomas L Hampton
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General Electric Co
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General Electric Co
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Publication date
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Publication of US3632223A publication Critical patent/US3632223A/en
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    • 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
    • F04D29/522—Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps

Definitions

  • the present invention relates to improvements in gas turbine engines having axial flow compressors and, more particularly, to improved interstage bleed air systems therefor.
  • the invention herein described was made in the course of or under a contract, or a subcontract thereunder, with the United States Department of the Air Force.
  • the primary function of the compressor of a gas turbine engine is to pressurize air so that the energy level of the hot gas stream generated by the engine may be increased, This pressurized air is also utilized for other purposes, such as cooling the hot sections of the engine and pressurizing lubrication sumps for rotating parts.
  • the object of the present invention is to improve such bleed air systems and overcome or minimize the limitations of prior systems which have caused relatively high energy losses in the bled air and increased weight and length of the engine in incorporating a bleed system.
  • annular bleed flow passageway in the casing of the compressor with the entrance to this passageway being defined by an annular lip disposed between a row of stator vanes and a row of rotor blades at the discharge of the stator vanes.
  • the bleed passageway is a diffuser to reduce the velocity of the bleed air so that energy losses in its transmission to a point of utilization will be minimized.
  • the compressor casing be formed in sections joined by radial flanges adjacent the downstream end of the first rotor blade which is surrounded by the bleed air passageway.
  • This arrangement facilities provision of a replaceable insert to form the entrance lip and initial portion of the bleed passageway.
  • the insert which is relatively inexpensive, may be readily replaced in restoring the compressor to its full operating capability.
  • FIG. 1 is a simplified illustration, partially in section, of a gas turbine engine embodying the present invention
  • FIG. 2 is an enlarged, detailed section of the present interstage bleed system
  • FIG. 3 is a section, taken on line III-III, in FIG. 2.
  • the engine seen in FIG. 1 comprises an outer casing which may be compositely formed by several sections. Air enters an inlet at one end of the casing 10 and is pressurized by a multistage axial flow compressor 12. The pressurized air supports combustion of fuel in a combustor 14 to generate a hot gas stream. A portion of the energy of the hot gas stream is used to drive a turbine 16. The turbine rotor 18 is connected by a shaft 20 to the compressor rotor 22. The remaining energy of the hot gas stream may be converted to a propulsive force by being discharged from a nozzle 24.
  • the flow path for bleed air from the compressor to its point of utilization may take many forms which are oftentimes dictated by the type of gas turbine engine and design philosophy. For that reason, broken line 26 is employed to represent the flow path of the bleed to a point of utilization in cooling a portion of the turbine 16.
  • the schematically represented flow path begins at a point after the bleed air has been diverted from the compressor by structure embodying the present invention, which will now be described with reference to FIGS. 2 and 3.
  • the engine casing surrounding the compressor includes semicylindrical sections 30 and 32 which are joined along a longitudinal split line (not shown) and by bolts 34 extending through radial fianges 36.
  • Stator vanes 38 are mounted in a circumferential row on the casing section 30.
  • the compressor flow path is reduced in area intermediate the vanes 38 and the next successive row of blades 40 mounted on the compressor rotor 22.
  • the reduction in compressor flow area is accom-' plished with an annular lip 39 which defines the entrance a bleed air flow passageway 41.
  • the passageway 41 is essentially annular, being interrupted only by longitudinal ribs 42 which connect the inner and outer walls of the passageway 41. It will further be noted that the passageway 41 increases in cross section from its inlet, at lip 39, to its outlet, at the downstream end of the casing section 32, in order that the bleed will be diffused as it travels along this passageway.
  • the lip 39 and the initial portion of passageway 41 are formed by curved inserts 43 which are detachably mounted on the casing sections 30 as by the illustrated lug and groove arrangement wherein slots (not shown) in the groove provide for installation and removal after the fashion of a bayonet connection.
  • the initial portion of the passageway 41 surrounds the rotating row of blades 40.
  • the described inserts facilitate repair of the casing in the event the blades 40 rub against the inner wall of the casing.
  • the inserts 43 may simply be replaced in the same fashion as the inserts which surround other rows of rotor blades.
  • the low included angle of the annular lip 39 minimizes flow losses and permits the highvelocity air to be diverted from the compressor at a low turning angle.
  • the outer surface of the bleed passageway 41 is a smoothly curved continuation of the inner surface of the casing section 30. Further, by disposing the lip 39 downstream of the row of stator vanes 38, the bleed air enters the bleed passageway 41 in an essentially axial direction. All of this minimizes losses in the velocity head of the air being diverted. This means that the bleed air is able to be utilized at essentially the total pressure of the compressor stage from which it is diverted.
  • the diffuser action of the passageway 41 built into the compressor sections 32, effectively drops the velocity of the air so that it may pass through the flow path 26 to a point of utilization with a minimum energy loss.
  • a gas turbine engine including a multistage, axial flow compressor having a flow path of progressively decreasing area comprising,
  • an outer casing having axially spaced, circumferential rows of inwardly projecting stator vanes mounted thereon and through which the flow path area is maintained at least as small at the exit as at the entrance thereto,
  • a rotor having circumferential rows of blades projecting therefrom and respectively spaced upstream of the vane rows
  • annular lip spaced inwardly from the outer bounds of flow path through the vanes and also inwardly of the outer ends of said one vane row and defining an upstream facing entrance to said bleed passageway, said lip being relatively sharp and having a low included angle whereby the divergence of said bleed passageway from the compressor flow path is at an relatively low angle, the inner surface of said lip also defining the outer bounds of the flow path past the next succeeding blade row the area of which is proportionately reduced by approximately the area of the entrance to said bleed passageway, and
  • the bleed passageway increases in crosssectional area along its length to diffuse the bleed air to a lower velocity.
  • the inner and outer portions of the casing which define the opposite sides of the bleed passageway are interconnected by longitudinal ribs.
  • the casing is split into sections along a plane normal to the engine axis and disposed at the downstream end of the first row of rotor blades which the bleed passageway surrounds,
  • a replaceable insert mounted on the upstream casing section and providing said annular lip and the initial portion of said bleed passageway
  • said insert facilitating repair of said casing in the event of a rotor blade rub thereon.
  • the bleed passageway increases in crosssectional area along its length to diffuse the bleed air to a lower velocity and the inner and outer portions of the casing which define the opposite sides of the bleed passageway are interconnected by longitudinal ribs.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US862197A 1969-09-30 1969-09-30 Turbine engine having multistage compressor with interstage bleed air system Expired - Lifetime US3632223A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US86219769A 1969-09-30 1969-09-30

Publications (1)

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US3632223A true US3632223A (en) 1972-01-04

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US (1) US3632223A (fr)
BE (1) BE752584A (fr)
DE (1) DE2031612A1 (fr)
FR (1) FR2060529A5 (fr)
GB (1) GB1310401A (fr)

Cited By (36)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3730639A (en) * 1970-07-17 1973-05-01 Secr Defence Fan or compressor for a gas turbine engine
US3735593A (en) * 1970-02-11 1973-05-29 Mini Of Aviat Supply In Her Br Ducted fans as used in gas turbine engines of the type known as fan-jets
US3826084A (en) * 1970-04-28 1974-07-30 United Aircraft Corp Turbine coolant flow system
US3966355A (en) * 1975-06-24 1976-06-29 Westinghouse Electric Corporation Steam turbine extraction system
US3976394A (en) * 1975-07-18 1976-08-24 The United States Of America As Represented By The Secretary Of The Air Force Interstage bleed assembly
US4155681A (en) * 1977-02-14 1979-05-22 General Electric Company Manifold protection system
US4156344A (en) * 1976-12-27 1979-05-29 The Boeing Company Inlet guide vane bleed system
US4512715A (en) * 1980-07-22 1985-04-23 Electric Power Research Institute, Inc. Method and means for recapturing coolant in a gas turbine
US4979587A (en) * 1989-08-01 1990-12-25 The Boeing Company Jet engine noise suppressor
US5160241A (en) * 1991-09-09 1992-11-03 General Electric Company Multi-port air channeling assembly
US5327716A (en) * 1992-06-10 1994-07-12 General Electric Company System and method for tailoring rotor tip bleed air
US5351478A (en) * 1992-05-29 1994-10-04 General Electric Company Compressor casing assembly
US5531565A (en) * 1993-08-10 1996-07-02 Abb Management Ag Appliance for extracting secondary air from an axial compressor
EP0877149A3 (fr) * 1997-05-07 2000-02-02 Rolls-Royce Plc Refroidissement du carter d'une turbine à gaz
US6109868A (en) * 1998-12-07 2000-08-29 General Electric Company Reduced-length high flow interstage air extraction
WO1999051866A3 (fr) * 1998-02-26 2001-07-19 Allison Advanced Dev Co Systeme de purge de paroi d'extremite de compresseur
US6438941B1 (en) 2001-04-26 2002-08-27 General Electric Company Bifurcated splitter for variable bleed flow
US6622475B2 (en) * 2001-04-12 2003-09-23 Snecma Moteurs Bleed system driven in simplified manner for a turbojet or turboprop engine
US20030223863A1 (en) * 2002-05-31 2003-12-04 Mitsubishi Heavy Industries, Ltd. Gas turbine compressor and clearance controlling method therefor
US20050141991A1 (en) * 2001-10-17 2005-06-30 Frutschi Hans U. Method for conditioning a compressor airflow and device therefor
US20090000306A1 (en) * 2006-09-14 2009-01-01 Damle Sachin V Stator assembly including bleed ports for turbine engine compressor
US20090196739A1 (en) * 2006-05-31 2009-08-06 Naoki Tsuchiya Axial flow fluid device
US20090297335A1 (en) * 2007-10-30 2009-12-03 Apostolos Pavlos Karafillis Asymmetric flow extraction system
US8307943B2 (en) 2010-07-29 2012-11-13 General Electric Company High pressure drop muffling system
US8430202B1 (en) 2011-12-28 2013-04-30 General Electric Company Compact high-pressure exhaust muffling devices
US8511096B1 (en) 2012-04-17 2013-08-20 General Electric Company High bleed flow muffling system
US8550208B1 (en) 2012-04-23 2013-10-08 General Electric Company High pressure muffling devices
US20140338360A1 (en) * 2012-09-21 2014-11-20 United Technologies Corporation Bleed port ribs for turbomachine case
US9399951B2 (en) 2012-04-17 2016-07-26 General Electric Company Modular louver system
US9689315B2 (en) * 2015-02-13 2017-06-27 Hamilton Sundstrand Corporation Full-area bleed valves
US20180313364A1 (en) * 2017-04-27 2018-11-01 General Electric Company Compressor apparatus with bleed slot including turning vanes
US20180313276A1 (en) * 2017-04-27 2018-11-01 General Electric Company Compressor apparatus with bleed slot and supplemental flange
US20190145420A1 (en) * 2017-11-13 2019-05-16 United Technologies Corporation Gas turbine engine with mid-compressor bleed
US20200149474A1 (en) * 2018-11-09 2020-05-14 United Technologies Corporation Internal heat exchanger system to cool gas turbine engine components
US20230332539A1 (en) * 2022-04-13 2023-10-19 General Electric Company Compressor bleed air channels having a pattern of vortex generators
US12146423B2 (en) 2023-01-11 2024-11-19 General Electric Company Compressor bleed pressure recovery

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3945759A (en) * 1974-10-29 1976-03-23 General Electric Company Bleed air manifold
GB2192229B (en) * 1986-07-04 1990-05-02 Rolls Royce Plc A compressor and air bleed system
GB2251031B (en) * 1990-12-19 1995-01-18 Rolls Royce Plc Cooling air pick up
GB201015029D0 (en) 2010-09-10 2010-10-20 Rolls Royce Plc Gas turbine engine

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US25253A (en) * 1859-08-30 Faucet
US2314058A (en) * 1941-06-23 1943-03-16 Edward A Stalker Pump
US2693904A (en) * 1950-11-14 1954-11-09 A V Roe Canada Ltd Air bleed for compressors
US2738921A (en) * 1950-11-22 1956-03-20 United Aircraft Corp Boundary layer control apparatus for compressors
GB869576A (en) * 1958-05-15 1961-05-31 Rolls Royce Improvements in compressors for gas turbine engines
US3142438A (en) * 1961-04-21 1964-07-28 Rolls Royce Multi-stage axial compressor

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US25253A (en) * 1859-08-30 Faucet
US2314058A (en) * 1941-06-23 1943-03-16 Edward A Stalker Pump
US2693904A (en) * 1950-11-14 1954-11-09 A V Roe Canada Ltd Air bleed for compressors
US2738921A (en) * 1950-11-22 1956-03-20 United Aircraft Corp Boundary layer control apparatus for compressors
GB869576A (en) * 1958-05-15 1961-05-31 Rolls Royce Improvements in compressors for gas turbine engines
US3142438A (en) * 1961-04-21 1964-07-28 Rolls Royce Multi-stage axial compressor

Cited By (51)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3735593A (en) * 1970-02-11 1973-05-29 Mini Of Aviat Supply In Her Br Ducted fans as used in gas turbine engines of the type known as fan-jets
US3826084A (en) * 1970-04-28 1974-07-30 United Aircraft Corp Turbine coolant flow system
US3730639A (en) * 1970-07-17 1973-05-01 Secr Defence Fan or compressor for a gas turbine engine
US3966355A (en) * 1975-06-24 1976-06-29 Westinghouse Electric Corporation Steam turbine extraction system
US3976394A (en) * 1975-07-18 1976-08-24 The United States Of America As Represented By The Secretary Of The Air Force Interstage bleed assembly
US4156344A (en) * 1976-12-27 1979-05-29 The Boeing Company Inlet guide vane bleed system
US4155681A (en) * 1977-02-14 1979-05-22 General Electric Company Manifold protection system
US4512715A (en) * 1980-07-22 1985-04-23 Electric Power Research Institute, Inc. Method and means for recapturing coolant in a gas turbine
US4979587A (en) * 1989-08-01 1990-12-25 The Boeing Company Jet engine noise suppressor
US5160241A (en) * 1991-09-09 1992-11-03 General Electric Company Multi-port air channeling assembly
US5351478A (en) * 1992-05-29 1994-10-04 General Electric Company Compressor casing assembly
US5327716A (en) * 1992-06-10 1994-07-12 General Electric Company System and method for tailoring rotor tip bleed air
US5531565A (en) * 1993-08-10 1996-07-02 Abb Management Ag Appliance for extracting secondary air from an axial compressor
EP0877149A3 (fr) * 1997-05-07 2000-02-02 Rolls-Royce Plc Refroidissement du carter d'une turbine à gaz
US6089821A (en) * 1997-05-07 2000-07-18 Rolls-Royce Plc Gas turbine engine cooling apparatus
WO1999051866A3 (fr) * 1998-02-26 2001-07-19 Allison Advanced Dev Co Systeme de purge de paroi d'extremite de compresseur
US6428271B1 (en) 1998-02-26 2002-08-06 Allison Advanced Development Company Compressor endwall bleed system
EP1147291A4 (fr) * 1998-02-26 2003-03-26 Allison Advanced Dev Co Systeme de purge de paroi d'extremite de compresseur
US6109868A (en) * 1998-12-07 2000-08-29 General Electric Company Reduced-length high flow interstage air extraction
US6622475B2 (en) * 2001-04-12 2003-09-23 Snecma Moteurs Bleed system driven in simplified manner for a turbojet or turboprop engine
US6438941B1 (en) 2001-04-26 2002-08-27 General Electric Company Bifurcated splitter for variable bleed flow
US20050141991A1 (en) * 2001-10-17 2005-06-30 Frutschi Hans U. Method for conditioning a compressor airflow and device therefor
US20030223863A1 (en) * 2002-05-31 2003-12-04 Mitsubishi Heavy Industries, Ltd. Gas turbine compressor and clearance controlling method therefor
US6732530B2 (en) * 2002-05-31 2004-05-11 Mitsubishi Heavy Industries, Ltd. Gas turbine compressor and clearance controlling method therefor
US20090196739A1 (en) * 2006-05-31 2009-08-06 Naoki Tsuchiya Axial flow fluid device
US20090000306A1 (en) * 2006-09-14 2009-01-01 Damle Sachin V Stator assembly including bleed ports for turbine engine compressor
US8292567B2 (en) 2006-09-14 2012-10-23 Caterpillar Inc. Stator assembly including bleed ports for turbine engine compressor
US20090297335A1 (en) * 2007-10-30 2009-12-03 Apostolos Pavlos Karafillis Asymmetric flow extraction system
US8388308B2 (en) 2007-10-30 2013-03-05 General Electric Company Asymmetric flow extraction system
US8307943B2 (en) 2010-07-29 2012-11-13 General Electric Company High pressure drop muffling system
US8430202B1 (en) 2011-12-28 2013-04-30 General Electric Company Compact high-pressure exhaust muffling devices
US8511096B1 (en) 2012-04-17 2013-08-20 General Electric Company High bleed flow muffling system
US9399951B2 (en) 2012-04-17 2016-07-26 General Electric Company Modular louver system
US8550208B1 (en) 2012-04-23 2013-10-08 General Electric Company High pressure muffling devices
US20140338360A1 (en) * 2012-09-21 2014-11-20 United Technologies Corporation Bleed port ribs for turbomachine case
US9689315B2 (en) * 2015-02-13 2017-06-27 Hamilton Sundstrand Corporation Full-area bleed valves
US20180313276A1 (en) * 2017-04-27 2018-11-01 General Electric Company Compressor apparatus with bleed slot and supplemental flange
US10934943B2 (en) * 2017-04-27 2021-03-02 General Electric Company Compressor apparatus with bleed slot and supplemental flange
CN108799202A (zh) * 2017-04-27 2018-11-13 通用电气公司 具有包括导流板的排放槽的压缩机设备
CN108799200A (zh) * 2017-04-27 2018-11-13 通用电气公司 具有排放槽和辅助法兰的压缩机设备
US20180313364A1 (en) * 2017-04-27 2018-11-01 General Electric Company Compressor apparatus with bleed slot including turning vanes
US11719168B2 (en) * 2017-04-27 2023-08-08 General Electric Company Compressor apparatus with bleed slot and supplemental flange
CN108799202B (zh) * 2017-04-27 2022-05-17 通用电气公司 具有包括导流板的排放槽的压缩机设备
CN113757172A (zh) * 2017-04-27 2021-12-07 通用电气公司 具有排放槽和辅助法兰的压缩机设备
US20190145420A1 (en) * 2017-11-13 2019-05-16 United Technologies Corporation Gas turbine engine with mid-compressor bleed
US10626879B2 (en) * 2017-11-13 2020-04-21 United Technologies Corporation Gas turbine engine with mid-compressor bleed
US10823069B2 (en) * 2018-11-09 2020-11-03 Raytheon Technologies Corporation Internal heat exchanger system to cool gas turbine engine components
US20200149474A1 (en) * 2018-11-09 2020-05-14 United Technologies Corporation Internal heat exchanger system to cool gas turbine engine components
US20230332539A1 (en) * 2022-04-13 2023-10-19 General Electric Company Compressor bleed air channels having a pattern of vortex generators
US11828226B2 (en) * 2022-04-13 2023-11-28 General Electric Company Compressor bleed air channels having a pattern of vortex generators
US12146423B2 (en) 2023-01-11 2024-11-19 General Electric Company Compressor bleed pressure recovery

Also Published As

Publication number Publication date
FR2060529A5 (fr) 1971-06-18
BE752584A (fr) 1970-12-01
GB1310401A (en) 1973-03-21
DE2031612A1 (de) 1971-04-15

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