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US3897168A - Turbomachine extraction flow guide vanes - Google Patents

Turbomachine extraction flow guide vanes Download PDF

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Publication number
US3897168A
US3897168A US448310A US44831074A US3897168A US 3897168 A US3897168 A US 3897168A US 448310 A US448310 A US 448310A US 44831074 A US44831074 A US 44831074A US 3897168 A US3897168 A US 3897168A
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United States
Prior art keywords
disc
channel shaped
rotor disc
shaped members
generally
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Expired - Lifetime
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US448310A
Inventor
David J Amos
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CBS Corp
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Westinghouse Electric Corp
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Publication date
Application filed by Westinghouse Electric Corp filed Critical Westinghouse Electric Corp
Priority to US448310A priority Critical patent/US3897168A/en
Priority to CA221,035A priority patent/CA986417A/en
Priority to IT20900/75A priority patent/IT1033402B/en
Priority to JP2571675A priority patent/JPS5327842B2/ja
Application granted granted Critical
Publication of US3897168A publication Critical patent/US3897168A/en
Anticipated expiration legal-status Critical
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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
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • 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

  • a gas turbine compressor comprising a rotor disc has continuous flow bleed vanes attached thereto.
  • the bleed vanes are channel shaped members attached to [52] US. Cl 415/144; 415/115 the downstream side of the rotor disc near its periph Cle y and en jacent es he bleed vanes [58] held of Search 416/95 f 415/ i rect the compressible fluid radially inwardly between 15/1 the rotor disc and a coolant guide disc attached to the downstream side of the rotor disc.
  • This invention relates to compressors for gas turbines, and more particularly to such structures adapted for continuously bleeding air from the compressor.
  • Prior Art Gas turbine engines often include a compressor structure which permits extraction or bleeding of the compressible fluid, usually air, from between two of the downstream stages of the compressor, or from the last stage of the compressor, to provide pressurized air for cooling purposes or for accessory use.
  • the compressible fluid usually air
  • turbornachine extraction or bleed systems for the compressors require minimization of pressure losses within the system.
  • Present designs employ valves and ports in annular arraysdisposed radially outwardly of the blades. These arrays are often complicated mechanical systems. They are heavy and they increase the dimensions of the equipment. Plenum structures must be accommodated for these bleed systems. Special .du cting, bridging and external piping must also be employed with many of the prior art extraction systems.
  • the applicant has invented a novel structure in which a final or intermediate stage continuous bleeding Of a compressor is maintained internally, while minimizing the complexity of the machine, minimizing the total pressure losses, and overcoming the objections to the prior art.
  • the present invention provides a simple, yet effective structure for extracting fluid from the flow path of a compressor. This is accomplished by providing channel shaped members disposed on the downstream side of the rotor disc.
  • the channel shaped members have leg portions directed upstream and the leg portions are formed about the steeples on the rotor disc.
  • the channel shaped members duct the compressed fluid radially inwardly between the rotor disc and a coolant guide disc. The fluid is then channeled to portions of the turbine for cooling purposes.
  • FIG. 1 is a perspective view of a portion of the downstream side of a compressor rotor disc showing an'extraction vane mounted thereon;
  • FIG. 2 is a perspective view of a portion of the upstream side of a compressor rotor disc showing the extraction vane mounted therewith;
  • FIG. 3 is a view directed radially inwardly on a compressor disc showing a few of the blades, and the extraction vanes mounted therewith;
  • FIG. 4 is an alternative embodiment of the extraction vane shown in FIG. 1;
  • FIG. 5 is a perspective view of a portion of a compressor rotor disc and a guide disc downstream thereof;
  • FIG. 6 is a side sectional view of a last stage compressor rotor disc having extraction vanes and a guide disc mounted thereon;
  • FIG. 7 is a side sectional view of an intermediate stage compressor rotor disc having extraction vanes and a guide disc mounted thereon.
  • FIG. 1 there is shown a compressor rotor disc 10.
  • Compressors are usually comprised of a series of these rotor discs mounted axially together and disposed in a housing.
  • the discs have airfoil shaped blades. not shown, mounted thereon to cause compression upon the fluid flowing therepast.
  • the rotor disc 10 shown in FIG. 1 for this example, is a last stage disc.
  • the rotor disc 10 has a plurality of grooves l2 disposed on its periphery.
  • the grooves 12 receive root portions of airfoil blades, not shown.
  • a steeple 14 is disposed between adjacent grooves 12.
  • the fluid which maybe air or hedrawn into an air bleed vane 16 disposed about the downstream side of the steeple vl4.
  • the air bleed vane 16 is a generally channel shaped member having leg portions 18 and 18', formed around the walls of the grooves 12 and directed upstream.
  • the bleed vane 16 is also comprised of a radially directed generally U- shaped member 20.
  • the U-shaped member 20 has a wall portion 22 whose radially innermost portion gradually curves away from the steeple 14, as shown in FIG. 1. This permits the securing of the bleed vane disc by the utilization of an annularly shaped guide disc 24 in addition to the securing of a bleed vane 16 by the use of leg portions 18 and 18'.
  • the guide disc 24 is shown further in FIGS. 5, 6 and 7.
  • the bleed vane 16 has tabs 26 and 26' that are bent around the upstream side of the steeple 14, as shown in FIG. 2.
  • the tabs 26 and 26, are the end portions of the leg portions 18 and 18, respectively.
  • the guide disc 24 is disposed downstream of and adjacent to the bleed vanes 16 as indicated by the dotted lines in FIG. 3.
  • a plurality of airfiol blades 28 are shown disposed in the grooves 12 on the disc 10. The sides of the blade roots are shown wedging the leg portions 18 and 18', close to the steeples 14.
  • the fluid as shown by vector V1, in FIG. 3, enters the bleed vane 16 and is turned to the axial direction, indicated by vector V2, simultaneously as it is directed radially inwardly.
  • the tangential velocity, indicated by the vector V0 is reduced to zero relative to the rotor disc 10, as the fluid flow enters the bleed system.
  • FIG. 4 A slight modification of a portion of the bleed vane 16, is shown in FIG. 4.
  • the U-shaped channel member 20 has one side 21 that is radially shorter, as indicated by the letter B in FIG. 4, than are the other walls of the channel member 20.
  • the wall 21 of the channel 20 in the direction of disc rotation being radially shorter than the other walls permits an added inducement of fluid flow through the channel member 20.
  • the area between the bleed vane 16 and the guide disc 24 will be of lower pressure than the pressure in the motive fluid flow path so as to draw the fluid therethrough.
  • FIGS. and 6 A portion of the annularly shaped guide disc 24 for a last stage rotor is shown in FIGS. and 6.
  • the guide disc 24 is bolted to the rotor disc by an annular array of bolts 30.
  • a radial array of ducts 32 permits the fluid to flow between the rotor disc 10 and the guide disc 24 from where it will be channeled to critical areas requir ing cooling fluid or pressurized fluid.
  • FIG. 6 Also shown in FIG. 6, is a portion of a compressor housing 34 having annular arrays of stationary blades 36. An additional rotor disc 37 is shown upstream of the disc 10. The curved wall portion 22 of U-shaped channel portion is shown in a close fitting restrained relationship with the guide disc 24.
  • the fluid is drawn through the compressible fluid flow path, indicated by the letter C.
  • the area of the flow path is reduced as the fluid is drawn therethrough.
  • the compressed fluid is ducted to a combustion area of a turbine, not shown, for mixture with fuel to drive the turbine.
  • a portion of the compressed fluid as indicated in this invention is'continuously drawn through bleed vane 16 due to the pressure difference between the fluid flow pathC and the fluid extraction route as indicated by the letter D in FIG. 6.
  • FIG. 7 An alternative arrangement is shown in FIG. 7, wherein a portion of a compressor 40 has an annular array of bleed vanes 16 and a guide disc 42 disposed on an intermediate rotor disc 44, and wherein a last stage rotor disc 46 does not have a bleed system disposed thereon.
  • An array of stationary airfoil shaped blades 48 is disposed between annular arrays of rotating airfoil shaped blades 50 and 52 respectively.
  • the fluid is continuously extracted through the bleed vane 16 as was shown in the prior figures.
  • the fluid in this embodiment follows the path indicated by the arrows E," around connecting bolts 53 into a coolant passageway 51.
  • channel shaped members are disposed on the downstream side of steeples between adjacent rotor blades, and an annularly shaped guid disc is disposed on the downstream side of the rotor disc, between which are disposed an annular array of channel shaped members to help duct the fluid to an area in the turbine for cooling purposes.
  • the guide disc also helps secure the channel shaped member to the rotor disc.
  • An axial flow compressor having a continuous fluid bleed system including:
  • extraction flow guide vane means attached to the side of said rotor disc, said extraction flow guide vane means comprising generally channel shaped members directed radially inwardly, said channel shaped members being attached to the downstream side of said rotor disc, adjacent the periphery of said disc, and disposed between adjacent turbine blades, said generally channel shaped members having generally parallel leg portions, one leg of each channel shaped member being restrained from dislocation by being attached to a wall of one groove on said disc, the other leg of each of said channel shaped members being attached to an adjacent wall of an adjacent groove on said disc, said legs of said channel shaped members being additionally restrained from dislocation by the wedging action of said blades holding said legs between themselves and said walls of said grooves;
  • coolant flow guide disc attached to the downstream side of said rotor, said coolant flow guide disc guiding the extracted fluid generally radially inwardly adjacent said rotor disc.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Life Sciences & Earth Sciences (AREA)
  • Sustainable Development (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

A gas turbine compressor comprising a rotor disc has continuous flow bleed vanes attached thereto. The bleed vanes are channel shaped members attached to the downstream side of the rotor disc near its periphery and between adjacent blades. The bleed vanes direct the compressible fluid radially inwardly between the rotor disc and a coolant guide disc attached to the downstream side of the rotor disc. The bleed vanes minimize the difference in tangential velocities of the extraction flow and the rotor disc speed at the inlet region of the continuous bleed system. The extracted fluid is ducted thereafter to a turbine where it is employed for cooling purposes.

Description

United States Patent Amos July 29, 1975 TURBOMACHINE EXTRACTION FLOW 3,632,221 H1972 Vehling 415/115 GUIDE VANES 3,647,313 3/1972 Koff 415/115 [75] lnventor: David J. Amos, Wallmgford, Pa. Primary Examiner Henry F. Raduazo [73] Assignee: Westinghouse Electric Corporation, Attorney, Agent, or Firm-G. H. Telfer Pittsburgh, Pa.
[22] Filed: Mar. 5, 1974 [57] ABSTRACT [21] A l N 448,310 A gas turbine compressor comprising a rotor disc has continuous flow bleed vanes attached thereto. The bleed vanes are channel shaped members attached to [52] US. Cl 415/144; 415/115 the downstream side of the rotor disc near its periph Cle y and en jacent es he bleed vanes [58] held of Search 416/95 f 415/ i rect the compressible fluid radially inwardly between 15/1 the rotor disc and a coolant guide disc attached to the downstream side of the rotor disc. The bleed vanes [56] References Clted minimize the difference in tangential velocities of the UNITED STATES PATENTS extraction flow and the rotor disc speed at the inlet 2,618,433 11/1952 Loos et a1. 415/115 region of the continuous bleed system. The extracted 2,636,665 4/1953 Lombard 415/115 fluid is ducted thereafter to a turbine where it is em- DaVleS et a1. ployed for cooling purposes 3,031,132 4/1962 Davies 415/115 3,575,522 4/1971 Melenchuk 416/96 5 Claims, 7 Drawing Figures 22 D as 19 3.2 37 30 3, 8 97. 16 8 PATENTED JULEQIQYS SHEET 2 FIG.4
PATENTEU JULZ 9 i975 SHEET SHEET PATENTED JUL 2 9 B75 TURBOMACI'IINE EXTRACTION FLOW GUIDE VANES.
1. Field of the Invention This invention relates to compressors for gas turbines, and more particularly to such structures adapted for continuously bleeding air from the compressor.
2, Prior Art Gas turbine engines often include a compressor structure which permits extraction or bleeding of the compressible fluid, usually air, from between two of the downstream stages of the compressor, or from the last stage of the compressor, to provide pressurized air for cooling purposes or for accessory use.
The turbornachine extraction or bleed systems for the compressors require minimization of pressure losses within the system. Present designs employ valves and ports in annular arraysdisposed radially outwardly of the blades. These arrays are often complicated mechanical systems. They are heavy and they increase the dimensions of the equipment. Plenum structures must be accommodated for these bleed systems. Special .du cting, bridging and external piping must also be employed with many of the prior art extraction systems.
To this end, in accordance with the objects of the invention, the applicant has invented a novel structure in which a final or intermediate stage continuous bleeding Of a compressor is maintained internally, while minimizing the complexity of the machine, minimizing the total pressure losses, and overcoming the objections to the prior art.
I I SUMMARY OF THE INVENTION The present invention provides a simple, yet effective structure for extracting fluid from the flow path of a compressor. This is accomplished by providing channel shaped members disposed on the downstream side of the rotor disc. The channel shaped members have leg portions directed upstream and the leg portions are formed about the steeples on the rotor disc. The channel shaped members duct the compressed fluid radially inwardly between the rotor disc and a coolant guide disc. The fluid is then channeled to portions of the turbine for cooling purposes. The pressure losses of the fluid-are reduced to a minimum and the tangential velocity of the compressed fluid with respect tothe rotor tem.
BRIEF DESCRIPTION OF THE DRAWINGS The invention, along with the objects and advantages thereof will be best understood from the following detailed description taken in conjunction with the accompanying'drawings in which:
FIG. 1 is a perspective view of a portion of the downstream side of a compressor rotor disc showing an'extraction vane mounted thereon;
FIG. 2 is a perspective view of a portion of the upstream side of a compressor rotor disc showing the extraction vane mounted therewith;
FIG. 3 is a view directed radially inwardly on a compressor disc showing a few of the blades, and the extraction vanes mounted therewith;
FIG. 4 is an alternative embodiment of the extraction vane shown in FIG. 1;
FIG. 5 is a perspective view of a portion of a compressor rotor disc and a guide disc downstream thereof;
FIG. 6 is a side sectional view of a last stage compressor rotor disc having extraction vanes and a guide disc mounted thereon; and,
FIG. 7 is a side sectional view of an intermediate stage compressor rotor disc having extraction vanes and a guide disc mounted thereon.
DESCRIPTION OF THE PREFERRED EMBODIMENT Referring to the drawings in detail, and particularly to FIG. 1, there is shown a compressor rotor disc 10. Compressors are usually comprised of a series of these rotor discs mounted axially together and disposed in a housing. The discs have airfoil shaped blades. not shown, mounted thereon to cause compression upon the fluid flowing therepast. The rotor disc 10 shown in FIG. 1 for this example, is a last stage disc. The rotor disc 10 has a plurality of grooves l2 disposed on its periphery. The grooves 12 receive root portions of airfoil blades, not shown. A steeple 14 is disposed between adjacent grooves 12. The fluid, which maybe air or hedrawn into an air bleed vane 16 disposed about the downstream side of the steeple vl4. The air bleed vane 16 is a generally channel shaped member having leg portions 18 and 18', formed around the walls of the grooves 12 and directed upstream. The bleed vane 16 is also comprised of a radially directed generally U- shaped member 20. The U-shaped member 20 has a wall portion 22 whose radially innermost portion gradually curves away from the steeple 14, as shown in FIG. 1. This permits the securing of the bleed vane disc by the utilization of an annularly shaped guide disc 24 in addition to the securing of a bleed vane 16 by the use of leg portions 18 and 18'. The guide disc 24 is shown further in FIGS. 5, 6 and 7.
The bleed vane 16 has tabs 26 and 26' that are bent around the upstream side of the steeple 14, as shown in FIG. 2. The tabs 26 and 26, are the end portions of the leg portions 18 and 18, respectively. When the blades, not shown in this figure, are inserted into their grooves 12, they frictionally engage the leg portions 18 and 18' of the bleed vane 16 and also by a wedging action, help secure the bleed vanes to the rotor disc 10.
The guide disc 24 is disposed downstream of and adjacent to the bleed vanes 16 as indicated by the dotted lines in FIG. 3. A plurality of airfiol blades 28 are shown disposed in the grooves 12 on the disc 10. The sides of the blade roots are shown wedging the leg portions 18 and 18', close to the steeples 14. The fluid, as shown by vector V1, in FIG. 3, enters the bleed vane 16 and is turned to the axial direction, indicated by vector V2, simultaneously as it is directed radially inwardly. The tangential velocity, indicated by the vector V0, is reduced to zero relative to the rotor disc 10, as the fluid flow enters the bleed system.
A slight modification of a portion of the bleed vane 16, is shown in FIG. 4. The U-shaped channel member 20 has one side 21 that is radially shorter, as indicated by the letter B in FIG. 4, than are the other walls of the channel member 20. The wall 21 of the channel 20 in the direction of disc rotation being radially shorter than the other walls permits an added inducement of fluid flow through the channel member 20. Also, in either case the area between the bleed vane 16 and the guide disc 24 will be of lower pressure than the pressure in the motive fluid flow path so as to draw the fluid therethrough.
A portion of the annularly shaped guide disc 24 for a last stage rotor is shown in FIGS. and 6. The guide disc 24 is bolted to the rotor disc by an annular array of bolts 30. A radial array of ducts 32 permits the fluid to flow between the rotor disc 10 and the guide disc 24 from where it will be channeled to critical areas requir ing cooling fluid or pressurized fluid.
Also shown in FIG. 6, is a portion of a compressor housing 34 having annular arrays of stationary blades 36. An additional rotor disc 37 is shown upstream of the disc 10. The curved wall portion 22 of U-shaped channel portion is shown in a close fitting restrained relationship with the guide disc 24.
As is well known in the art, the fluid is drawn through the compressible fluid flow path, indicated by the letter C. The area of the flow path is reduced as the fluid is drawn therethrough. The compressed fluid is ducted to a combustion area of a turbine, not shown, for mixture with fuel to drive the turbine.
A portion of the compressed fluid as indicated in this invention is'continuously drawn through bleed vane 16 due to the pressure difference between the fluid flow pathC and the fluid extraction route as indicated by the letter D in FIG. 6.
An alternative arrangement is shown in FIG. 7, wherein a portion of a compressor 40 has an annular array of bleed vanes 16 and a guide disc 42 disposed on an intermediate rotor disc 44, and wherein a last stage rotor disc 46 does not have a bleed system disposed thereon. An array of stationary airfoil shaped blades 48 is disposed between annular arrays of rotating airfoil shaped blades 50 and 52 respectively. The fluid is continuously extracted through the bleed vane 16 as was shown in the prior figures. The fluid in this embodiment however, follows the path indicated by the arrows E," around connecting bolts 53 into a coolant passageway 51.
From the foregoing description it should now be apparent that a new and useful continuous bleed vane structure has been disclosed for axial flow compressors. in which channel shaped members are disposed on the downstream side of steeples between adjacent rotor blades, and an annularly shaped guid disc is disposed on the downstream side of the rotor disc, between which are disposed an annular array of channel shaped members to help duct the fluid to an area in the turbine for cooling purposes. The guide disc also helps secure the channel shaped member to the rotor disc.
Though the invention has been described with a certain degree of particularity, changes may be made therein without departing from the scope thereof. For example, the invention has been applied to only one of a plurality of rotor discs. Obviously a plurality of such discs could be used simultaneously, or the channel members could be modified to increase the flow rate therethrough.
I claim as my invention:
1. An axial flow compressor having a continuous fluid bleed system including:
at least one rotor disc in said compressor;
a plurality of generally axially directed grooves on the periphery of said rotor disc;
an annular array of radially directed airfoil blades supportively disposed in said grooves;
extraction flow guide vane means attached to the side of said rotor disc, said extraction flow guide vane means comprising generally channel shaped members directed radially inwardly, said channel shaped members being attached to the downstream side of said rotor disc, adjacent the periphery of said disc, and disposed between adjacent turbine blades, said generally channel shaped members having generally parallel leg portions, one leg of each channel shaped member being restrained from dislocation by being attached to a wall of one groove on said disc, the other leg of each of said channel shaped members being attached to an adjacent wall of an adjacent groove on said disc, said legs of said channel shaped members being additionally restrained from dislocation by the wedging action of said blades holding said legs between themselves and said walls of said grooves;
a coolant flow guide disc attached to the downstream side of said rotor, said coolant flow guide disc guiding the extracted fluid generally radially inwardly adjacent said rotor disc.
2. An axial flow compressor arrangement as recited in claim 1, wherein one generally channel shaped memher is disposed between adjacent blades.
3. An axial flow compressor as recited in claim 1, wherein said leg portions of said channel shaped members are attached to said rotor disc by bending the leg portions to conform to the groove wall and rotor disc configurations.
4. An axial flow compressor as recited in claim 3, wherein said leg portions of said generally channel shaped members are directed in the generally upstream direction.
5. An axial flow compressor as recited in claim 4, wherein said generally channel shaped members have downstream wall portions, said downstream wall portions being generally curved to mate with said guide disc in a restraining relationship.

Claims (5)

1. An axial flow compressor having a continuous fluid bleed system including: at least one rotor disc in said compressor; a plurality of generally axially directed grooves on the periphery of said rotor disc; an annular array of radially directed airfoil blades supportively disposed in said grooves; extraction flow guide vane means attached to the side of said rotor disc, said extraction flow guide vane means comprising generally channel shaped members directed radially inwardly, said channel shaped members being attached to the downstream side of said rotor disc, adjacent the periphery of said disc, and disposed between adjacent turbine blades, said generally channel shaped members having generally parallel leg portions, one leg of each channel shaped member being restrained from dislocation by being attached to a wall of one groove on said disc, the other leg of each of said channel shaped members being attached to an adjacent wall of an adjacent groove on said disc, said legs of said channel shaped members being additionally restrained from disloCation by the wedging action of said blades holding said legs between themselves and said walls of said grooves; a coolant flow guide disc attached to the downstream side of said rotor, said coolant flow guide disc guiding the extracted fluid generally radially inwardly adjacent said rotor disc.
2. An axial flow compressor arrangement as recited in claim 1, wherein one generally channel shaped member is disposed between adjacent blades.
3. An axial flow compressor as recited in claim 1, wherein said leg portions of said channel shaped members are attached to said rotor disc by bending the leg portions to conform to the groove wall and rotor disc configurations.
4. An axial flow compressor as recited in claim 3, wherein said leg portions of said generally channel shaped members are directed in the generally upstream direction.
5. An axial flow compressor as recited in claim 4, wherein said generally channel shaped members have downstream wall portions, said downstream wall portions being generally curved to mate with said guide disc in a restraining relationship.
US448310A 1974-03-05 1974-03-05 Turbomachine extraction flow guide vanes Expired - Lifetime US3897168A (en)

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US448310A US3897168A (en) 1974-03-05 1974-03-05 Turbomachine extraction flow guide vanes
CA221,035A CA986417A (en) 1974-03-05 1975-02-28 Turbomachine extraction flow guide vanes
IT20900/75A IT1033402B (en) 1974-03-05 1975-03-04 AXIAL COMPRESSOR EQUIPPED WITH A CONTINUOUS FLUID EXTRACTION SYSTEM
JP2571675A JPS5327842B2 (en) 1974-03-05 1975-03-04

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Cited By (17)

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US4231704A (en) * 1977-08-26 1980-11-04 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Cooling fluid bleed for axis of turbine rotor
US4415310A (en) * 1980-10-08 1983-11-15 Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." System for cooling a gas turbine by bleeding air from the compressor
US4787820A (en) * 1987-01-14 1988-11-29 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Turbine plant compressor disc with centripetal accelerator for the induction of turbine cooling air
DE3828834C1 (en) * 1988-08-25 1989-11-02 Mtu Muenchen Gmbh
US5211533A (en) * 1991-10-30 1993-05-18 General Electric Company Flow diverter for turbomachinery seals
WO1998030803A1 (en) 1997-01-13 1998-07-16 Massachusetts Institute Of Technology Counter-rotating compressors with control of boundary layers by fluid removal
WO1998030802A1 (en) * 1997-01-13 1998-07-16 Massachusetts Institute Of Technology Enhancement of turbomachines and compressors by fluid removal
GB2411697A (en) * 2004-03-06 2005-09-07 Rolls Royce Plc Cooling arrangement for rim of turbine disc.
WO2006024273A1 (en) * 2004-09-01 2006-03-09 Mtu Aero Engines Gmbh Rotor for a power plant
FR2930589A1 (en) * 2008-04-24 2009-10-30 Snecma Sa CENTRIFIC AIR COLLECTION IN A COMPRESSOR ROTOR OF A TURBOMACHINE
CN101900132A (en) * 2009-05-28 2010-12-01 通用电气公司 Turbomachine compressor wheel member
US20100300113A1 (en) * 2009-05-27 2010-12-02 Grewal Daljit Singh Anti-vortex device for a gas turbine engine compressor
US20140294597A1 (en) * 2011-10-10 2014-10-02 Snecma Cooling for the retaining dovetail of a turbomachine blade
US20140308133A1 (en) * 2011-11-15 2014-10-16 Snecma Rotor wheel for a turbine engine
US10519976B2 (en) 2017-01-09 2019-12-31 Rolls-Royce Corporation Fluid diodes with ridges to control boundary layer in axial compressor stator vane
CN113847280A (en) * 2021-10-10 2021-12-28 中国航发沈阳发动机研究所 Compressor rotor interstage bleed air structure
CN113898610A (en) * 2021-10-10 2022-01-07 中国航发沈阳发动机研究所 Gas-entraining structure for disk center of rotor disk of compressor

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US2618433A (en) * 1948-06-23 1952-11-18 Curtiss Wright Corp Means for bleeding air from compressors
US2636665A (en) * 1947-03-11 1953-04-28 Rolls Royce Gas turbine engine
US2910268A (en) * 1951-10-10 1959-10-27 Rolls Royce Axial flow fluid machines
US3031132A (en) * 1956-12-19 1962-04-24 Rolls Royce Gas-turbine engine with air tapping means
US3575522A (en) * 1968-08-30 1971-04-20 Gen Motors Corp Turbine cooling
US3632221A (en) * 1970-08-03 1972-01-04 Gen Electric Gas turbine engine cooling system incorporating a vortex shaft valve
US3647313A (en) * 1970-06-01 1972-03-07 Gen Electric Gas turbine engines with compressor rotor cooling

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US2636665A (en) * 1947-03-11 1953-04-28 Rolls Royce Gas turbine engine
US2618433A (en) * 1948-06-23 1952-11-18 Curtiss Wright Corp Means for bleeding air from compressors
US2910268A (en) * 1951-10-10 1959-10-27 Rolls Royce Axial flow fluid machines
US3031132A (en) * 1956-12-19 1962-04-24 Rolls Royce Gas-turbine engine with air tapping means
US3575522A (en) * 1968-08-30 1971-04-20 Gen Motors Corp Turbine cooling
US3647313A (en) * 1970-06-01 1972-03-07 Gen Electric Gas turbine engines with compressor rotor cooling
US3632221A (en) * 1970-08-03 1972-01-04 Gen Electric Gas turbine engine cooling system incorporating a vortex shaft valve

Cited By (29)

* Cited by examiner, † Cited by third party
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US4231704A (en) * 1977-08-26 1980-11-04 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Cooling fluid bleed for axis of turbine rotor
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CA986417A (en) 1976-03-30
JPS50121811A (en) 1975-09-25
IT1033402B (en) 1979-07-10
JPS5327842B2 (en) 1978-08-10

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