EP0487242A1 - Abblastruktur eines Verdichters - Google Patents

Abblastruktur eines Verdichters Download PDF

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Publication number
EP0487242A1
EP0487242A1 EP19910310455 EP91310455A EP0487242A1 EP 0487242 A1 EP0487242 A1 EP 0487242A1 EP 19910310455 EP19910310455 EP 19910310455 EP 91310455 A EP91310455 A EP 91310455A EP 0487242 A1 EP0487242 A1 EP 0487242A1
Authority
EP
European Patent Office
Prior art keywords
compressor
air
slot
pressure
bleed
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
EP19910310455
Other languages
English (en)
French (fr)
Other versions
EP0487242B1 (de
Inventor
William Francis Mcgreehan
Andrew John Lammas
Bradley Willis Fintel
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.)
General Electric Co
Original Assignee
General Electric Co
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Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Publication of EP0487242A1 publication Critical patent/EP0487242A1/de
Application granted granted Critical
Publication of EP0487242B1 publication Critical patent/EP0487242B1/de
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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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
    • 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
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/542—Bladed diffusers
    • 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

  • This invention relates to simplified bleed ex traction slots for gas turbine engines and, more particularly, to a specially configured bleed extraction slot for efficiently converting core air to bleed air with a minimum loss in bleed air velocity and pressure.
  • an aircraft gas turbine engine include within its compressor, a structure which permits bleeding or diversion of high pressure air from a stage, such as the 5th stage of the compressor to provide high pressure air for cooling purposes and for operation of airframe accessories, engine accessories, or engine or aircraft de-icing systems.
  • a structure which permits the bleeding of even higher pressure air from the discharge of the compressor to provide pressurized air for cooling downstream turbine components Both interstage bleed and the compressor discharge bleeding are accomplished by flowpath mechanisms which interfere with the normal airflow patterns in the compressor. Further, the casing or bleed structure adds complexity to the assembly of such an engine.
  • the axial location or stage at which air is bled from the compressor is determined by the pressure required to drive the specific system intended to be serviced by that air. In most instances, it is desirable to achieve the highest possible source pressure to also ensure a high delivery pressure. For this reason, prior systems have extracted air from the latter stages of the compressor and more particularly, engines having these systems have been designed to extract high pressure air from the 5th stage of the compressor for low pressure turbine cooling and turbine thermal clearance control. However, bleeding air from the earliest possible stage of the compressor generally increases compressor efficiency by reducing the amount of work invested in the extracted air. Therefore, it is desirable to achieve the highest possible system supply pressure from the earliest and lowest pressure stage of the compressor. The resulting temperature of the cooling air is also lower and hence more effective.
  • U.S. Patent 4,711,084 to Brockett for an ejector-assisted compressor bleed which discloses a bleed aperture 17 in Fig. 2 having rounded hole edges.
  • U.S. Patent 3,108,767 to Eltis, et al., for a bypass gas turbine engine with an air bleed means in Fig. 3 discloses a duct 19 which is attached to the compressor through a series of chopped holes.
  • U.S. Patent 3,898,799 to Pollert, et al., for a device for bleeding off compressor air in a turbine jet engine, in Fig. 2 discloses a compressor orifice marked with the arrow K.
  • Patent 3,777,489 to Johnston, et al. discloses a combustor casing having a concentric air bleed structure which includes a series of conical arms 62, 64, and 66 situated in the low velocity area of the diffuser with the bled air structure making a turn of approximately 180°.
  • U.S. Patent 4,344,282 to Anders is directed to a compressor bleed system which includes a locking strap 12 which seals a series of bleed ports 8.
  • a high pressure compressor bleed air extraction slot structure comprising a compressor outer band having a bleed air portion positioned proximate a rearward and preferably interstage section of a compressor.
  • a diffusing slot can be disposed in the compressor outer casing and can comprise an articulated or punched-out portion of the outer band articulated at an angle approximately 10-20 degrees from a band baseline whereby the diffusion coefficient of the bleed valve is improved.
  • the articulated angle is 15 degrees and the exit velocity V2 is less than the baseline velocity V1 while the exit pressure P2 is greater than the baseline pressure P1.
  • a gas turbine engine 10 is shown in major cross section to include a fan rotor 12, and a core engine rotor 14.
  • the fan rotor 12 includes a plurality of fan blades 16 mounted for rotation on a disk 20.
  • the fan rotor 12 also includes a low pressure or fan turbine 22 which drives the fan disk 20 in a well known manner.
  • the core engine rotor 14 includes a compressor 24 and a high power or high pressure turbine 26 which drives the compressor 24.
  • the core engine also includes a combustion system 28.
  • Air entering the inlet 30 is compressed by means of the rotation of fan blades 16 and thereafter is split into two flow streams, a bypass stream 34 flowing in a bypass passageway 35, and a core engine stream 36 flowing in a core passageway 37.
  • the pressurized air which enters the core engine passageway 37 is further pressurized by means of the compressor 24 and is thereafter mixed and ignited along with high energy fuel in the combustion system 28.
  • This highly energized gas stream then flows through the high pressure turbine 26 to drive the compressor 24 and thereafter through the low pressure turbine 22 to drive the fan rotor 12 and disk 20.
  • the pressurized air flowing through the bypass passageway 35 is either mixed with the core engine exhaust system stream by means of a suitable mixer (not shown) or is allowed to exhaust to ambient conditions as a relatively low velocity, low pressure stream surrounding the core engine exhaust. In either case, the core engine stream 36 exhaust and fan bypass stream 34 exhaust provide a propulsive force for an aircraft powered by the turbofan engine 10.
  • a diffusing port or hole 40 comprises an orifice 42 located in line with an outer band 44 of the engine cowling or casing 32.
  • the compressor casing structure 44 provides an annular orifice 42 immediately upstream of one of the intermediate stages of the rotor blades 38 for bleeding inner stage air from the interior of the compressor 24.
  • the compressor 24 includes a rotor 14 having a number of rotor stages 40 which carry a plurality of rotor blades 38.
  • the compressor 24 further includes a casing structure 32 which defines the outer bounds of the compressor flowpath and includes mounting provisions for a plurality of stator vanes 46 aligned in individual stages between each stage of rotor blades 38.
  • the outer band 44 includes a diffuser slot 62 comprising a punched-out and articulated portion 64 articulated at an angle of between 10 and 20 degrees and preferably 15 degrees measured from a baseline 60 of the outer band 44.
  • FIG. 3A a comparison of the prior annular 5th stage orifice 42 is shown in Fig. 3A in relation to the present articulated 4th stage diffuser slot 62 in accordance with the present invention, shown in Fig. 3B.
  • the annular orifice 42 induces a swirling airflow 50 which substantially restricts the opening of orifice 42 and reduces the discharge coefficient C d associated with the orifice.
  • the annular orifice 42 requires the exiting air to alter its velocity by approximately 90 degrees with a concommitant energy reduction.
  • a diffusing slot 62 in accordance with the present invention which is shown in Fig. 3B, includes an articulate portion 64 which expands the volume of a lateral cavity 54 of the compressor vane to cause the cavity to immediately capture diffuser air and minimally change the velocity and energy level of the captured air.
  • the volume of the lateral cavity 54 is considered to be the volume between the casing baseline 60 and the articulated member 64.
  • the swirl pattern established by this slot 62 occurs closely adjacent the slot's surfaces 44 and 64 and thus introduces a minimal obstruction to the air flowpath. Accordingly, the pressure drop associated with the slot 62 is minimized, the discharge coefficient, C d , associated with this slot is maximized and the energy level of air passing through the diffuser is maintained.
  • the efficient energy conversion achieved by this slot produces air at a higher pressure than that previously achieved. Accordingly, the slot 62 can be applied to an earlier or lower pressure stage of the compressor and yet still supply air of a pressure equivalent to that previously derived from a later stage.
  • the bleed slot 62 of the present invention provides a means to recover and convert a portion of the gas steam dynamic pressure into a manifold static pressure rise.
  • the angled recessed surface of the articulated portion 64 acts as a diffuser to decelerate the air as it passes through the outer band opening thereby reducing the irreversible losses in energy.
  • the invention can be characterized based on test data which shows clearly that a higher C d is achieved for the diffusing slot 62 compared to a standard orifice 42 each having the same cross-sectional area. More particularly, in a typical 9-stage compressor, the prior orifice 42 when applied to the 5th stage could achieve a discharge pressure of 132 psia (16/in2) at temperature of 1207°R (Rankine). In contrast, the present invention, when applied to the 4th stage of the same compressor, can achieve a discharge pressure of 118 psia at temperature of 1089°R; thus, improving the efficiency of the engine.
  • the diffuser extraction slot 62 of the present invention allows a portion of the gas flowpath velocity pressure to be recovered as usable manifold static pressure.
  • This higher pressurized flow allows the bleed extraction point to be relocated at least one stage forward in the compressor and represents an overall increase in efficiency and engine performance which can be reflected in lower specific fuel consumption.
  • the extraction of air earlier in the compressor provides a lower temperature source for turbine cooling systems.
  • the size and location of the diffuser slot can be changed to reflect the pressure drop and flow requirements of the system(s) that the bleed slots supplies.
  • the shape of the orifice can be changed such that the pressure gradient across the opening can be minimized to insure a high pressure flow.
  • the bleed diffuser slot construction of the present invention can be adapted to fit a number of gas turbine engines as described herein.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP91310455A 1990-11-19 1991-11-13 Abblastruktur eines Verdichters Expired - Lifetime EP0487242B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US61567690A 1990-11-19 1990-11-19
US615676 1990-11-19

Publications (2)

Publication Number Publication Date
EP0487242A1 true EP0487242A1 (de) 1992-05-27
EP0487242B1 EP0487242B1 (de) 1995-09-20

Family

ID=24466390

Family Applications (1)

Application Number Title Priority Date Filing Date
EP91310455A Expired - Lifetime EP0487242B1 (de) 1990-11-19 1991-11-13 Abblastruktur eines Verdichters

Country Status (4)

Country Link
EP (1) EP0487242B1 (de)
JP (1) JP2513954B2 (de)
CA (1) CA2048829C (de)
DE (1) DE69113209T2 (de)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9145772B2 (en) 2012-01-31 2015-09-29 United Technologies Corporation Compressor disk bleed air scallops
US9260974B2 (en) 2011-12-16 2016-02-16 General Electric Company System and method for active clearance control
EP2803822B1 (de) * 2013-05-13 2019-12-04 Safran Aero Boosters SA Luftentnahmesystem aus einer axialen Turbomaschine
CN113847280A (zh) * 2021-10-10 2021-12-28 中国航发沈阳发动机研究所 一种压气机转子级间引气结构

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9528391B2 (en) 2012-07-17 2016-12-27 United Technologies Corporation Gas turbine engine outer case with contoured bleed boss
JP6000142B2 (ja) * 2013-01-28 2016-09-28 三菱重工業株式会社 回転機械、及びこれを備えているガスタービン
JP6134628B2 (ja) 2013-10-17 2017-05-24 三菱重工業株式会社 軸流式の圧縮機、及びガスタービン
GB201518448D0 (en) * 2015-10-19 2015-12-02 Rolls Royce Compressor

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3108767A (en) * 1960-03-14 1963-10-29 Rolls Royce By-pass gas turbine engine with air bleed means
DE1428216A1 (de) * 1961-04-21 1969-07-31 Rolls Royce Mehrstufen-Axialkompressor
US3777489A (en) * 1972-06-01 1973-12-11 Gen Electric Combustor casing and concentric air bleed structure
US3898799A (en) * 1972-09-27 1975-08-12 Mtu Muenchen Gmbh Device for bleeding-off compressor air in turbine jet engine
US4344282A (en) * 1980-12-16 1982-08-17 United Technologies Corporation Compressor bleed system
US4711084A (en) * 1981-11-05 1987-12-08 Avco Corporation Ejector assisted compressor bleed
EP0374004A1 (de) * 1988-12-15 1990-06-20 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Abblasventil eines Triebwerkverdichters

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS5364112A (en) * 1976-11-19 1978-06-08 Hitachi Ltd Gas turbine compressor
US4546605A (en) * 1983-12-16 1985-10-15 United Technologies Corporation Heat exchange system
JPS6124675U (ja) * 1984-07-17 1986-02-14 日本電気株式会社 耐電圧試験器

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3108767A (en) * 1960-03-14 1963-10-29 Rolls Royce By-pass gas turbine engine with air bleed means
DE1428216A1 (de) * 1961-04-21 1969-07-31 Rolls Royce Mehrstufen-Axialkompressor
US3777489A (en) * 1972-06-01 1973-12-11 Gen Electric Combustor casing and concentric air bleed structure
US3898799A (en) * 1972-09-27 1975-08-12 Mtu Muenchen Gmbh Device for bleeding-off compressor air in turbine jet engine
US4344282A (en) * 1980-12-16 1982-08-17 United Technologies Corporation Compressor bleed system
US4711084A (en) * 1981-11-05 1987-12-08 Avco Corporation Ejector assisted compressor bleed
EP0374004A1 (de) * 1988-12-15 1990-06-20 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Abblasventil eines Triebwerkverdichters

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9260974B2 (en) 2011-12-16 2016-02-16 General Electric Company System and method for active clearance control
US9145772B2 (en) 2012-01-31 2015-09-29 United Technologies Corporation Compressor disk bleed air scallops
EP2803822B1 (de) * 2013-05-13 2019-12-04 Safran Aero Boosters SA Luftentnahmesystem aus einer axialen Turbomaschine
CN113847280A (zh) * 2021-10-10 2021-12-28 中国航发沈阳发动机研究所 一种压气机转子级间引气结构

Also Published As

Publication number Publication date
CA2048829C (en) 2001-12-18
DE69113209D1 (de) 1995-10-26
JP2513954B2 (ja) 1996-07-10
EP0487242B1 (de) 1995-09-20
DE69113209T2 (de) 1996-05-02
JPH04284136A (ja) 1992-10-08
CA2048829A1 (en) 1992-05-20

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