US3588268A - Dump bleed system for the compressor of a gas turbine engine - Google Patents

Dump bleed system for the compressor of a gas turbine engine Download PDF

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US3588268A
US3588268A US862196A US3588268DA US3588268A US 3588268 A US3588268 A US 3588268A US 862196 A US862196 A US 862196A US 3588268D A US3588268D A US 3588268DA US 3588268 A US3588268 A US 3588268A
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compressor
ring
bleed
gas turbine
casing
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US862196A
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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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    • 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

  • the casing of the axial flow compressor is Int. Cl ..F 04d 27/02, provided with a circumferential bleed port of unique configu- F04d 27/00 ration. This port is selectively closed by a ring having tapered Field of Search 415/144, sealing surfaces engaged on opposite sides of the port 145, 146 discharge.
  • the present invention relates to improvements in gas turbine engines and, more particularly, to improved compressor bleed means for such engines.
  • the compressor of a gas turbine engine pressurizes air to increase the energy of a hot gas stream which is generated by burning fuel in the pressurized air stream. Part of the gas stream energy is employed to drive a turbine while the remainder is converted to an output, as by being discharged from a propulsive nozzle.
  • the turbine drives the compressor rotor at a rate dependent primarily upon the rate of fuel combustion.
  • the compressor is aerodynamically optimized for a given rate of compressor rotation as well as for other parameters such as flight speed, which will be encountered at the design point of the engine where maximum performance is desired.
  • the airflow through one portion of a multistage axial flow compressor is greater than can be accepted by a subsequent stage without causing an undue pressure rise across one or more stages of the compressor.
  • Such undue pressure rise, or increase in back pressure can result in the phenomenon of stall wherein airflow breaks down and there is a resultant loss of engine power.
  • the problem described is one which occurs primarily at low rotor speeds.
  • Bleed dump systems are selectively employed during engine operation at off design conditions. At other conditions, and particularly at design point operation, the bleed system must be completely closed since leakage of even small amounts of air from the compressor can have serious effects on overall engine efficiency.
  • Prior dump bleed systems have suffered one or more of several shortcomings. Some have been complicated. Some have been heavy, not only because of the mechanism employed but because of-increased engine length necessary for incorporation. Some have had problems in preventing leakage of compressor air when bleed flow is shut off.
  • the object of the present invention is to provide an improved dump bleed system for compressors of gas turbine engines and in so doing, to overcome or minimize the short comings of prior systems noted above.
  • the above ends are attained by providing a circumferential bleed port in the casing of a gas turbine engine compressor.
  • the casing has, on its exterior, tapered sealing surfaces on opposite sides of the circumferential port.
  • a ring having its inner surface correspondingly tapered and engageable with said sealing surfaces shutsoff bleed air in the closed position of the ring.
  • Means are provided for displacing the ring axially out of register with the port to bleed air from the compressor.
  • Additional features include forming the upstream entrance edge to the port on a gentle curvature.
  • the ring is lightweight, being of U-shaped cross section. The legs of this U-shaped section are generally aligned with the sealing surfaces of the casing in the closed position of the ring. This enables high sealing pressures to be exerted by the wedging action of the tapered surfaces. Guide means are also provided for assuring roundness of the lightweight sealing ring.
  • FIG. 1 is a simplified view of a gas turbine engine incorporating the present invention
  • FIG. 2 is a longitudinal section of the present bleed system
  • FIG. 3 is a view of the bleed system taken from the exterior of the engine.
  • the engine seen in FIG. 1 comprises an outer casing 10 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.
  • a dump bleed system As previously discussed, the function of a dump bleed system is to reduce the pressure at a selected point or station along the length of a compressor in order to prevent stall'dur ing operation of the compressor at off design point conditions.
  • the present system indicated at 25 bleeds air from the compressor between its fourth and fifth stages, as will be seen from FIG. 1.
  • FIGS. 2 and 3 show the system in greater detail.
  • the engine casing comprises a compressor casing'section 26 which is formed by semicylindrical shells joined at a longitudinal split line, not shown.
  • the stator vanes 28 are mounted in circumferential rows on and project inwardly from the casing section 26. Rotor blades 30 project from the rotor 22 between the rows of the vanes 28.
  • a circumferential port 32 is provided in the casing section 26 immediately downstream of the fourth stage stator vanes 28.
  • the entrance to this port is gently curved, at its upstream end, from the stator blades, to minimize turning losses when air is bled from the compressor.
  • the downstream entrance to this port is preferably formed as a relatively sharp edge, as shown in FIG. 2.
  • Ribs 34 span the circumferential port to give structural integrity to the casing section 26. The width of these ribs is minimized by locally increasing the thickness of the easing in the area of the circumferential port 32.
  • the described circumferential port enables bleeding a maximum amount of air from the compressor in a minimum of axial length and with a minimum impedance to its discharge flow.
  • a sealing ring 36 is formed by segments which are joined with bolts 38.
  • the ring is moved in an axial direction by three actuators 40 which are equiangularly spaced around the casing (two actuators are seen in FIG. I).
  • the actuators are mounted on a flange 42 which projects radially from the casing see section 26.
  • the rods 44 of these actuators are connected to the ring 36 by pins 46.
  • the ring is preferably U-shaped in cross section, as seen in FIG. 2, in order to minimize weight. With such a lightweight construction there is a tendency for the ring to deflect from an exact cylindrical shape.
  • a plurality of pads 48 project from the lower portions of the rear leg of the U-shaped ring and ride on iongitudinal ribs 50 formed on the outer surface of the casing sections.
  • the pads 48 and ribs 50 are provided at and intermediate the actuators at fairly close intervals.
  • the inner surface of the ring is tapered, as will be apparent from FIG. 2.
  • the outer surface of the casing on opposite sides of the circumferential port 32 is correspondingly tapered.
  • the tapered surfaces provide a wedging action which assures a positive seal in spite of any tendency of the ring or casing to deflect from an exact cylindrical shape.
  • the legs of the U-shaped ring generally overlie the axially spaced sealing surfaces so that substantial wedging pressures may be taken with a lightweight structure.
  • the ring 36 is displaced between its open and closed positions by the actuators 40 which would be controlled by any suitable means known to those skilled in the art.
  • the short axial length of the circumferential port enables rapid response in either initiating bleed flow or shutting it off. Movement of the ring to an inter mediate position enables partial bleed flow, if required, to be accurately controlled.
  • a gas turbine engine including axial flow compressor means and a dump bleed system comprising:
  • a casing defining the outer bounds of the flow path of air through said compressor means
  • said casing having, on its exterior, tapered sealing surfaces on opposite sides of said circumferential port,
  • a ring encompassing said casing and having an inner surface formed on a taper corresponding to that of the sealing surfaces and engageable therewith in the closed position of said ring in which there is not bleed flow
  • the upstream entrance edge to said circumferential port is gently curved to minimize flow losses when air is bled through said port.
  • said ring is U-shaped in cross section and the legs thereof generally overlie, respectively, the sealing surfaces on said casing when the ring is in its closed position.
  • longitudinal ribs span the circumferential port to transmit structural loadings between the portions of the casing on opposite sides of the port.
  • the ring has circumferentially spaced pads extending therefrom and the casing has correspondingly spaced, longitudinal ribs on which the pads ride and maintain the ring in essentially a cylindrical shape.
  • the means for axially displacing the ring comprise a plurality of actuators mounted on the exterior of said casing with their rods connected to said ring, said actuators being fewer in number than said ribs and pads.

Abstract

A DUMP BLEED SYSTEM FOR THE COMPRESSOR OF A GAS TURBINE ENGINE. THE CASING OF THE AXIAL FLOW COMPRESSOR IS PROVIDED WITH A CIRCUMFERENTIAL BLEED PORT OF UNIQUE CONFIGURATION.

THIS PORT IS SELECTIVELY CLOSED BY A RING HAVING TAPERED SEALING SURFACES ENGAGED ON OPPOSITE SIDES OF THE PORT DISCHARGE.

Description

Unite States Patent Thomas L. Hampton Loveland, Ohio Sept. 30, 1969 June 28, 1971 General Electric Company Inventor App]. No. Filed Patented Assignee Primary Examiner-Henry F. Raduazo Att0meys- Derek P. Lawrence, E. S. Lee, [11, Lee H. Sachs,
Frank I... Neuhauser, Oscar B. Waddell and Joseph B.
DUMP BLEED SYSTEM FOR THE COMPRESSOR Forman OF A GAS TURBINE ENGXNE 6 claimssnmwmg Figs ABSTRACT: A dump bleed system for the compressor of a US. Cl 415/145 gas turbine engine. The casing of the axial flow compressor is Int. Cl ..F 04d 27/02, provided with a circumferential bleed port of unique configu- F04d 27/00 ration. This port is selectively closed by a ring having tapered Field of Search 415/144, sealing surfaces engaged on opposite sides of the port 145, 146 discharge.
DUMP IiLEED SYSTEM FOR THE COMPRESSOR OF A GAS TURBINE ENGINE The present invention relates to improvements in gas turbine engines and, more particularly, to improved compressor bleed means for such engines.
The compressor of a gas turbine engine pressurizes air to increase the energy of a hot gas stream which is generated by burning fuel in the pressurized air stream. Part of the gas stream energy is employed to drive a turbine while the remainder is converted to an output, as by being discharged from a propulsive nozzle. The turbine, in turn, drives the compressor rotor at a rate dependent primarily upon the rate of fuel combustion.
The compressor is aerodynamically optimized for a given rate of compressor rotation as well as for other parameters such as flight speed, which will be encountered at the design point of the engine where maximum performance is desired. In many instances, when operating at conditions other than at the design point, the airflow through one portion of a multistage axial flow compressor is greater than can be accepted by a subsequent stage without causing an undue pressure rise across one or more stages of the compressor. Such undue pressure rise, or increase in back pressure, can result in the phenomenon of stall wherein airflow breaks down and there is a resultant loss of engine power. The problem described is one which occurs primarily at low rotor speeds.
One solution to this problem which has been used for some time is to bleed air and dump it overboard from the compressor casing. Bleed dump systems are selectively employed during engine operation at off design conditions. At other conditions, and particularly at design point operation, the bleed system must be completely closed since leakage of even small amounts of air from the compressor can have serious effects on overall engine efficiency.
Prior dump bleed systems have suffered one or more of several shortcomings. Some have been complicated. Some have been heavy, not only because of the mechanism employed but because of-increased engine length necessary for incorporation. Some have had problems in preventing leakage of compressor air when bleed flow is shut off.
The object of the present invention is to provide an improved dump bleed system for compressors of gas turbine engines and in so doing, to overcome or minimize the short comings of prior systems noted above.
The above ends are attained by providing a circumferential bleed port in the casing of a gas turbine engine compressor. The casing has, on its exterior, tapered sealing surfaces on opposite sides of the circumferential port. A ring having its inner surface correspondingly tapered and engageable with said sealing surfaces shutsoff bleed air in the closed position of the ring. Means are provided for displacing the ring axially out of register with the port to bleed air from the compressor.
Additional features include forming the upstream entrance edge to the port on a gentle curvature. The ring is lightweight, being of U-shaped cross section. The legs of this U-shaped section are generally aligned with the sealing surfaces of the casing in the closed position of the ring. This enables high sealing pressures to be exerted by the wedging action of the tapered surfaces. Guide means are also provided for assuring roundness of the lightweight sealing ring.
The above and other related objects and features of the invention will be apparent from a reading of the following description of the disclosure found in the accompanying drawings and the novelty thereof pointed out in the appended claims.
In the drawings:
FIG. 1 is a simplified view of a gas turbine engine incorporating the present invention;
FIG. 2 is a longitudinal section of the present bleed system; and
FIG. 3 is a view of the bleed system taken from the exterior of the engine.
The engine seen in FIG. 1 comprises an outer casing 10 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 preceding describes a typical gas turbine engine to set forth the environment of the present bleed system which is applicable to a wide variety of engine configurations, as will be apparent from the following detailed description of this system. I
As previously discussed, the function of a dump bleed system is to reduce the pressure at a selected point or station along the length of a compressor in order to prevent stall'dur ing operation of the compressor at off design point conditions. The present system, indicated at 25 bleeds air from the compressor between its fourth and fifth stages, as will be seen from FIG. 1.
FIGS. 2 and 3 show the system in greater detail. The engine casing comprises a compressor casing'section 26 which is formed by semicylindrical shells joined at a longitudinal split line, not shown. The stator vanes 28 are mounted in circumferential rows on and project inwardly from the casing section 26. Rotor blades 30 project from the rotor 22 between the rows of the vanes 28.
A circumferential port 32 is provided in the casing section 26 immediately downstream of the fourth stage stator vanes 28. The entrance to this port is gently curved, at its upstream end, from the stator blades, to minimize turning losses when air is bled from the compressor. The downstream entrance to this port is preferably formed as a relatively sharp edge, as shown in FIG. 2. Ribs 34 span the circumferential port to give structural integrity to the casing section 26. The width of these ribs is minimized by locally increasing the thickness of the easing in the area of the circumferential port 32. The described circumferential port enables bleeding a maximum amount of air from the compressor in a minimum of axial length and with a minimum impedance to its discharge flow.
A sealing ring 36 is formed by segments which are joined with bolts 38. The ring is moved in an axial direction by three actuators 40 which are equiangularly spaced around the casing (two actuators are seen in FIG. I). The actuators are mounted on a flange 42 which projects radially from the casing see section 26. The rods 44 of these actuators are connected to the ring 36 by pins 46.
The ring is preferably U-shaped in cross section, as seen in FIG. 2, in order to minimize weight. With such a lightweight construction there is a tendency for the ring to deflect from an exact cylindrical shape. In order to maintain a reasonable cylindrical shape, a plurality of pads 48 project from the lower portions of the rear leg of the U-shaped ring and ride on iongitudinal ribs 50 formed on the outer surface of the casing sections. The pads 48 and ribs 50 are provided at and intermediate the actuators at fairly close intervals.
The inner surface of the ring is tapered, as will be apparent from FIG. 2. Likewise, the outer surface of the casing on opposite sides of the circumferential port 32 is correspondingly tapered. This forms uninterrupted sealing surfaces which are engaged by the ring 36 when it is in its closed position, indicated by broken lines in FIG. 2. The tapered surfaces provide a wedging action which assures a positive seal in spite of any tendency of the ring or casing to deflect from an exact cylindrical shape. Further, the legs of the U-shaped ring generally overlie the axially spaced sealing surfaces so that substantial wedging pressures may be taken with a lightweight structure.
Operation of the present bleed system should be apparent from the preceding discussion. The ring 36 is displaced between its open and closed positions by the actuators 40 which would be controlled by any suitable means known to those skilled in the art. The short axial length of the circumferential port enables rapid response in either initiating bleed flow or shutting it off. Movement of the ring to an inter mediate position enables partial bleed flow, if required, to be accurately controlled.
While the present invention has been described as an in terstage dump bleed system for a multistage compressor, it is applicable to provide dump bleed at other locations in axial flow compressors of the single or multiple spool type. The location of the bleed point would be an aerodynamic function of a given compressor design. Further modifications in and departures from the preferred embodiment described will occur to those skilled in the art within the spirit and scope of the present inventive concepts.
Having thus described the invention, what is claimed as novel and desired to be secured by Letters Patent of the United States is:
lclaim:
1. A gas turbine engine including axial flow compressor means and a dump bleed system comprising:,
A casing defining the outer bounds of the flow path of air through said compressor means,
a circumferential bleed port formed in said compressor,
said casing having, on its exterior, tapered sealing surfaces on opposite sides of said circumferential port,
a ring encompassing said casing and having an inner surface formed on a taper corresponding to that of the sealing surfaces and engageable therewith in the closed position of said ring in which there is not bleed flow, and
means for axially displacing said ring between said closed position and an open position in which the ring is moved out of register with said circumferential port, said means forcing said ring to its closed position with the pressure between the sealing surfaces of the casing and ring being increased by the taper of these surfaces.
2. A gas turbine engine as in claim 1 wherein,
the upstream entrance edge to said circumferential port is gently curved to minimize flow losses when air is bled through said port.
3. A gas turbine engine as in claim 1 wherein,
said ring is U-shaped in cross section and the legs thereof generally overlie, respectively, the sealing surfaces on said casing when the ring is in its closed position.
4. A gas turbine engine as in claim 3 wherein,
longitudinal ribs span the circumferential port to transmit structural loadings between the portions of the casing on opposite sides of the port.
5. A gas turbine engine as in claim 3 wherein,
the ring has circumferentially spaced pads extending therefrom and the casing has correspondingly spaced, longitudinal ribs on which the pads ride and maintain the ring in essentially a cylindrical shape.
6. A gas turbine engine as in claim 5 wherein,
the means for axially displacing the ring comprise a plurality of actuators mounted on the exterior of said casing with their rods connected to said ring, said actuators being fewer in number than said ribs and pads.
US862196A 1969-09-30 1969-09-30 Dump bleed system for the compressor of a gas turbine engine Expired - Lifetime US3588268A (en)

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Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4674951A (en) * 1984-09-06 1987-06-23 Societe Nationale D'Etude et de Construction de Meteur D'Aviation (S.N.E.C.M.A.) Ring structure and compressor blow-off arrangement comprising said ring
US4679982A (en) * 1984-09-06 1987-07-14 Societe Nationale D'etude Et De Construction De Moteur D'aviation "S. N. E. C. M. A." Compressor blow-off arrangement
US5505587A (en) * 1995-01-05 1996-04-09 Northrop Grumman Corporation RAM air turbine generating apparatus
US6048171A (en) * 1997-09-09 2000-04-11 United Technologies Corporation Bleed valve system
US20050008476A1 (en) * 2003-07-07 2005-01-13 Andreas Eleftheriou Inflatable compressor bleed valve system
US20060277919A1 (en) * 2005-02-25 2006-12-14 Volvo Aero Corporation Bleed structure for a bleed passage in a gas turbine engine
US7152913B2 (en) 1999-10-22 2006-12-26 Mack Trucks, Inc. Modular sleeper units
US20120288359A1 (en) * 2011-05-12 2012-11-15 Rolls-Royce Deutschland Ltd & Co Kg Gas-turbine engine with bleed-air tapping device
US20140341708A1 (en) * 2011-12-29 2014-11-20 Rolls-Royce North American Technologies, Inc. Valve for gas turbine engine
US20150027130A1 (en) * 2012-09-28 2015-01-29 United Technologies Corporation Split ring valve
US10393128B2 (en) 2015-05-26 2019-08-27 Pratt & Whitney Canada Corp. Translating gaspath bleed valve
US11619170B1 (en) * 2022-03-07 2023-04-04 Rolls-Royce North American Technologies Inc. Gas turbine engine with radial turbine having modulated fuel cooled cooling air
US20230212989A1 (en) * 2022-01-05 2023-07-06 General Electric Company Bleed valve assemblies

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE2410139C3 (en) * 1974-03-02 1981-10-15 Klöckner-Humboldt-Deutz AG, 5000 Köln Gas turbine plant
US4086761A (en) * 1976-04-26 1978-05-02 The Boeing Company Stator bypass system for turbofan engine
US7624581B2 (en) * 2005-12-21 2009-12-01 General Electric Company Compact booster bleed turbofan

Cited By (19)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4679982A (en) * 1984-09-06 1987-07-14 Societe Nationale D'etude Et De Construction De Moteur D'aviation "S. N. E. C. M. A." Compressor blow-off arrangement
US4674951A (en) * 1984-09-06 1987-06-23 Societe Nationale D'Etude et de Construction de Meteur D'Aviation (S.N.E.C.M.A.) Ring structure and compressor blow-off arrangement comprising said ring
US5505587A (en) * 1995-01-05 1996-04-09 Northrop Grumman Corporation RAM air turbine generating apparatus
US6048171A (en) * 1997-09-09 2000-04-11 United Technologies Corporation Bleed valve system
US7152913B2 (en) 1999-10-22 2006-12-26 Mack Trucks, Inc. Modular sleeper units
US20050008476A1 (en) * 2003-07-07 2005-01-13 Andreas Eleftheriou Inflatable compressor bleed valve system
US6899513B2 (en) 2003-07-07 2005-05-31 Pratt & Whitney Canada Corp. Inflatable compressor bleed valve system
US20080115504A1 (en) * 2005-02-25 2008-05-22 Volvo Aero Corporation Bleed Structure For A Bleed Passage In A Gas Turbine Engine
US20060277919A1 (en) * 2005-02-25 2006-12-14 Volvo Aero Corporation Bleed structure for a bleed passage in a gas turbine engine
US8484982B2 (en) 2005-02-25 2013-07-16 Volvo Aero Corporation Bleed structure for a bleed passage in a gas turbine engine
US20120288359A1 (en) * 2011-05-12 2012-11-15 Rolls-Royce Deutschland Ltd & Co Kg Gas-turbine engine with bleed-air tapping device
US8944754B2 (en) * 2011-05-12 2015-02-03 Rolls-Royce Deutschland Ltd & Co Kg Gas-turbine engine with bleed-air tapping device
US20140341708A1 (en) * 2011-12-29 2014-11-20 Rolls-Royce North American Technologies, Inc. Valve for gas turbine engine
US9856884B2 (en) * 2011-12-29 2018-01-02 Rolls-Royce North American Technologies Inc. Valve for gas turbine engine
US20150027130A1 (en) * 2012-09-28 2015-01-29 United Technologies Corporation Split ring valve
US9328735B2 (en) * 2012-09-28 2016-05-03 United Technologies Corporation Split ring valve
US10393128B2 (en) 2015-05-26 2019-08-27 Pratt & Whitney Canada Corp. Translating gaspath bleed valve
US20230212989A1 (en) * 2022-01-05 2023-07-06 General Electric Company Bleed valve assemblies
US11619170B1 (en) * 2022-03-07 2023-04-04 Rolls-Royce North American Technologies Inc. Gas turbine engine with radial turbine having modulated fuel cooled cooling air

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BE756363A (en) 1971-03-01
DE2045983A1 (en) 1971-04-15
FR2068282A5 (en) 1971-08-20
GB1319348A (en) 1973-06-06

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