US3775023A - Multistage axial flow compressor - Google Patents

Multistage axial flow compressor Download PDF

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US3775023A
US3775023A US00116010A US3775023DA US3775023A US 3775023 A US3775023 A US 3775023A US 00116010 A US00116010 A US 00116010A US 3775023D A US3775023D A US 3775023DA US 3775023 A US3775023 A US 3775023A
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stages
stator
blades
rotor
vanes
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Expired - Lifetime
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US00116010A
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J Davis
T Ivsan
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TDY Industries LLC
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Teledyne Industries Inc
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE-DISPLACEMENT PUMPS
    • F04D19/00Axial-flow pumps
    • F04D19/02Multi-stage pumps
    • F04D19/028Layout of fluid flow through the stages
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12Blades
    • F01D5/14Form or construction
    • F01D5/141Shape, i.e. outer, aerodynamic form
    • F01D5/142Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F01D5/143Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00Details, component parts, or accessories
    • F04D29/40Casings; Connections of working fluid
    • F04D29/52Casings; Connections of working fluid for axial pumps
    • F04D29/522Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/60Efficient propulsion technologies, e.g. for aircraft

Definitions

  • ABSTRACT [52] CL 415/199 f 415/218 A multistage axial flow compressor with stacked 51 1111. C1. F04d 19/00, F04d 3/00 stages in an axial .each inch'ffing a [58] Field of Search 5 415/190 191 192 rotor and a stator. The stages include vanes of identi- 4l5ll93 218 cal vane geometry throughout the stator stages and i 172 the blades are of identical blade geometry throughout the several rotor stages. The blade and vane heights decrease from stage to stage in the direction of exiting [56] References Cited fluld flow from the compressor.
  • This invention relates generally to axial flow compressor machines of the kind having a plurality of consecutive stages of mutually cooperating relatively rotating blading carried by two generally cylindrical members.
  • the two members are arranged one within the other and are of different outer and inner diameters to provide an annular passage therebetween.
  • the fluid acted upon by the blading is progressively compressed so that the sectional area of the passage will change progressively along its length.
  • One of such cylindrical members may be stationary and the other rotated.
  • the two members may both be rotated in opposite directions to provide a contour rotational arrangement. It will be understood that the invention includes within its scope arrangements where either member may be fixed and the other rotated.
  • Conventional multiple stage axial compressors are generally comprised of rotors and stators having differently shaped blades from stage to stage. As the air in the compressor is pressurized, the flow area to each successive stage must be rendered in order to prevent excessive separation which itself tends to cause premature surge and low efficiency.
  • Compressors known to the prior art generally accomplish the decrease in flow area by adjusting the hub radius for the individual blades in each different stage. This dimensioning of flow area has in turn made it necessary to design the rotors and the stators of different stages with different blade sizes and shapes.
  • the geometry of the blades which is typically described by hub and tip radius, camber, thickness and chord length, must be optimized and custom designed for each stage individually based on inlet conditions and flow area requirements.
  • the present invention has particular application to air or gas compressors for internal combustion or gas turbine plants primarily for the propulsion of aircraft or other like craft.
  • the present invention involves the incorporation in multiple stage axial compressors of substantially identical stages with respect to both the blades and the vanes. Since all the respective rotor and stator blades are of the same'air foil section, they can be manufactured from the same master. This makes possible a relatively inexpensive compressor that is easy to design, manufacture and fabricate. In addition, the inventory of spare parts is greatly minimized.
  • the use of a constant diameter hub on the rotor has been found to maximize the axial exit blade height, provide a reasonable radius ratio centrifugal and reduce the number of blades and vanes required.
  • FIG. 1 is a longitudinal half-sectional view taken parallel to the centerline of the rotor
  • FIG. 2 is a sectional view of the apparatus of FIG. 1 taken along the section lines 2--2;
  • FIG. 3 is a perspective fragmentary view of the compressor stator of FIG. 1;
  • FIG. 4 is a perspective view of the compressor rotor of FIG. 1 with parts broken away.
  • FIG. 1 shows the detail of a multiple stage compressor incorporating the present invention. Included in the compressor are an outer stator 10 and an inner rotor 12. The rotor 12 and the stator 10 are arranged typically within a stationary cylinder or casing 14 which is shown in part in FIG. 3 hereinafter.
  • the stator 10 includes three successive rings or stages each of which contains a plurality of uniformly spaced, circumferentially arranged vanes 16a, 16b and 160, which vanes are affixed in pre-machined slots by welding, brazing or other like methods to the cylinder 18 of the stator 10.
  • the cylinder 18 may be of one piece construction or may include a plurality of cylindrical rings in which the rings are separately assembled with the vanes 16a, b or c, and finally fixed together along the joints 19. The detail of this type of construction is shown in greater detail in FIG. 3 hereinafter.
  • each of the cylindrical rings 18a is formed with a first inwardly sloping surface 19a, an intermediate surface 19b of constant diameter and a second inwardly sloping surface 19c.
  • the blades 16 are centered in the rings 18a and are therefore fixed to the surfaces 19b.
  • stator 10 has interspersed between them a plurality of rotor stages including blades 20a, b and c.
  • the rotor blades in a manner similar to the stator vanes are arranged in rings of gradually decreasing height in the direction of exiting fluid flow through the compressor.
  • the fluid flow direction through the compressor and toward the exit provided by centrifugal stage 22 is shown by arrows A.
  • the rotor blades 20a, b and c are fixed to a rotor hub 13, which itself is of a constant diameter cylindrical shape, by means similar to that used to fix the stator vanes in place.
  • the arrangement of the rotor 12 within the stator 10 typically includes within the compressor casing 14 a bearing and support mechanism, not shown, to provide for rotative movement of the rotor 12 within the stator 10.
  • the several blade carrying rings are preferably fabricated separately and then joined along a plurality of seams 15.
  • FIG. 2 taken along the lines 2--2 of FIG. 1 shows the form of the several stator vanes 16a, b and c and of the several rotor blades 20a, b and c.
  • the cross section of individual ones of the respective blades and vanes illustrates that the blades 20a, b and c are of an identical air foil configuration, particularly with respect to their blade geometry, which in more detail may be described by camber, thickness and chord length.
  • the vanes 16a, b and c are likewise of the same air foil configuration, which may be more specifically described with respect to camber thickness and chord length.
  • the differences reside mainly in the height differences as between the several successive rotor 12 and stator 10 stages.
  • FIG. 3 shows the left hand or intake end of the compressor and the several successive stages of stator vanes 16a, b and c, which are assembled in the three successive rings comprising the stator 10. It will be seen that in their assembled positions, the various stator vanes are all given a like orientation with respect to the air passage formed between the rotor and stator.
  • FIG. 4 shows the several successive stages provided for the rotor 12 and the manner in which the vanes a, b and c all have a like air foil configuration throughout the three stages with only a height difference existing between the three individual stages.
  • the rotor 12 is shown to have a constant diameter for the hub 13 throughout the several stages.
  • An axial flow compressor including a multistage stator housing, a multiple stage rotor mounted for rotation coaxially within said housing and having a constant diameter hub, each of said rotor stages having a plurality of blades arranged circumferentially about said hub, each of said stator stages having a plurality of vanes arranged circumferentially within said housing, all of said blades and vanes in each of said stages having an identical hub radius, camber, thickness and chord length but having a gradually diminishing height in the direction of exiting fluid flow and all of the blades in each different rotor stage having a tip radius sloping in the direction of exiting fluid flow to provide optimum flow whereby all of the blades of the various stages are of the same blade geometry except for height and tip radius and therefore can be manufactured from the same master, each of said stator stages comprising a ring member with an inner surface defined by an inwardly sloped section, an intermediate section of constant diameter and a second inwardly sloped section whereby the stator stages combine to provide a passage which reduce

Abstract

A multistage axial flow compressor with stacked stages in an axial direction, each stage including a rotor and a stator. The stages include vanes of identical vane geometry throughout the stator stages and the blades are of identical blade geometry throughout the several rotor stages. The blade and vane heights decrease from stage to stage in the direction of exiting fluid flow from the compressor. A constant hub diameter is maintained throughout the several rotor stages.

Description

United States Patent 11 1 Davis et a1. [45] N 27, 1973 [54] MULTISTAGE AXIAL FLOW COMPRESSOR 2,622,790 12/1952 McLeod 415/218 2,706,451 4/1955 Ortiz et a1 415/190 [75] Inventors- 9' Davis Lambemlle 3,012,308 12/1961 Zech et a1 29/156.813 Thomas Toledo, 3,241,493 3/1966 Freg; 415/191 Ohi 3 3,493,169 2/1970 Abild e181. 415 143 Assignee: Teledyne Industries, Inc. Los 2,151,699 3/1939 Hemer 415/219 Angeles, Calif. Primary Examiner-Henry F. Raduazo 22 Filed. Feb. 17, 1971 Ammey Hauke etaL [21] Appl. No.: 116,010
{ [57] ABSTRACT [52] CL 415/199 f 415/218 A multistage axial flow compressor with stacked 51 1111. C1. F04d 19/00, F04d 3/00 stages in an axial .each inch'ffing a [58] Field of Search 5 415/190 191 192 rotor and a stator. The stages include vanes of identi- 4l5ll93 218 cal vane geometry throughout the stator stages and i 172 the blades are of identical blade geometry throughout the several rotor stages. The blade and vane heights decrease from stage to stage in the direction of exiting [56] References Cited fluld flow from the compressor. A constant hub diam- UNITED STATES T eter is maintained throughout the several rotor stages. 2,540,968 2/1951 Thomas .1. ..415/170R 3,112,866 12/1963 Fortescue 415/181 1 Claim, 4 Drawing Figures /Ja. /9C a. m2
03 1 l: I l 1 11 I l 1' 11 0 i?: 3 4 204 l 11111 1 1 1| .30
I muv 27 ms 3.775.023
INVENTORS JAMES V. DAVIS BY THOMAS IVSAN #M, $5M m ATTO RN EYS l MULTISTAGE AXIAL FLOW COMPRESSOR BACKGROUND OF THE INVENTION This invention relates generally to axial flow compressor machines of the kind having a plurality of consecutive stages of mutually cooperating relatively rotating blading carried by two generally cylindrical members. The two members are arranged one within the other and are of different outer and inner diameters to provide an annular passage therebetween. The fluid acted upon by the blading is progressively compressed so that the sectional area of the passage will change progressively along its length. One of such cylindrical members may be stationary and the other rotated. It is generally more convenient to rotate the inner one with the blading carried by the outer member then constituting a guide blading. Alternately, the two members may both be rotated in opposite directions to provide a contour rotational arrangement. It will be understood that the invention includes within its scope arrangements where either member may be fixed and the other rotated.
Conventional multiple stage axial compressors are generally comprised of rotors and stators having differently shaped blades from stage to stage. As the air in the compressor is pressurized, the flow area to each successive stage must be rendered in order to prevent excessive separation which itself tends to cause premature surge and low efficiency. Compressors known to the prior art generally accomplish the decrease in flow area by adjusting the hub radius for the individual blades in each different stage. This dimensioning of flow area has in turn made it necessary to design the rotors and the stators of different stages with different blade sizes and shapes. The geometry of the blades which is typically described by hub and tip radius, camber, thickness and chord length, must be optimized and custom designed for each stage individually based on inlet conditions and flow area requirements.
SUMMARY OF THE INVENTION A The present invention has particular application to air or gas compressors for internal combustion or gas turbine plants primarily for the propulsion of aircraft or other like craft. The present invention involves the incorporation in multiple stage axial compressors of substantially identical stages with respect to both the blades and the vanes. Since all the respective rotor and stator blades are of the same'air foil section, they can be manufactured from the same master. This makes possible a relatively inexpensive compressor that is easy to design, manufacture and fabricate. In addition, the inventory of spare parts is greatly minimized. The use of a constant diameter hub on the rotor has been found to maximize the axial exit blade height, provide a reasonable radius ratio centrifugal and reduce the number of blades and vanes required.
BRIEF DESCRIPTION OF THE DRAWINGS The accompanying drawings illustrate the present invention wherein like numerals refer to like parts throughout the several views and wherein:
FIG. 1 is a longitudinal half-sectional view taken parallel to the centerline of the rotor;
FIG. 2 is a sectional view of the apparatus of FIG. 1 taken along the section lines 2--2;
FIG. 3 is a perspective fragmentary view of the compressor stator of FIG. 1; and
FIG. 4 is a perspective view of the compressor rotor of FIG. 1 with parts broken away.
DETAILED DESCRIPTION FIG. 1 shows the detail of a multiple stage compressor incorporating the present invention. Included in the compressor are an outer stator 10 and an inner rotor 12. The rotor 12 and the stator 10 are arranged typically within a stationary cylinder or casing 14 which is shown in part in FIG. 3 hereinafter. The stator 10 includes three successive rings or stages each of which contains a plurality of uniformly spaced, circumferentially arranged vanes 16a, 16b and 160, which vanes are affixed in pre-machined slots by welding, brazing or other like methods to the cylinder 18 of the stator 10. The cylinder 18 may be of one piece construction or may include a plurality of cylindrical rings in which the rings are separately assembled with the vanes 16a, b or c, and finally fixed together along the joints 19. The detail of this type of construction is shown in greater detail in FIG. 3 hereinafter.
As can best be seen in FIG. 1 each of the cylindrical rings 18a is formed with a first inwardly sloping surface 19a, an intermediate surface 19b of constant diameter and a second inwardly sloping surface 19c. The blades 16 are centered in the rings 18a and are therefore fixed to the surfaces 19b.
It will be seen that the several stages of stator 10 have interspersed between them a plurality of rotor stages including blades 20a, b and c. The rotor blades in a manner similar to the stator vanes are arranged in rings of gradually decreasing height in the direction of exiting fluid flow through the compressor. The fluid flow direction through the compressor and toward the exit provided by centrifugal stage 22 is shown by arrows A. The rotor blades 20a, b and c are fixed to a rotor hub 13, which itself is of a constant diameter cylindrical shape, by means similar to that used to fix the stator vanes in place. The arrangement of the rotor 12 within the stator 10 typically includes within the compressor casing 14 a bearing and support mechanism, not shown, to provide for rotative movement of the rotor 12 within the stator 10. The several blade carrying rings are preferably fabricated separately and then joined along a plurality of seams 15.
FIG. 2 taken along the lines 2--2 of FIG. 1 shows the form of the several stator vanes 16a, b and c and of the several rotor blades 20a, b and c. The cross section of individual ones of the respective blades and vanes illustrates that the blades 20a, b and c are of an identical air foil configuration, particularly with respect to their blade geometry, which in more detail may be described by camber, thickness and chord length. The vanes 16a, b and c are likewise of the same air foil configuration, which may be more specifically described with respect to camber thickness and chord length. The differences reside mainly in the height differences as between the several successive rotor 12 and stator 10 stages.
FIG. 3 shows the left hand or intake end of the compressor and the several successive stages of stator vanes 16a, b and c, which are assembled in the three successive rings comprising the stator 10. It will be seen that in their assembled positions, the various stator vanes are all given a like orientation with respect to the air passage formed between the rotor and stator.
FIG. 4 shows the several successive stages provided for the rotor 12 and the manner in which the vanes a, b and c all have a like air foil configuration throughout the three stages with only a height difference existing between the three individual stages. The rotor 12 is shown to have a constant diameter for the hub 13 throughout the several stages.
It has been found that the optimum arrangement for pressure ratio for the lowest number of stages, for example in the range of 9: 1-1 2:1 pressure ratio, was provided by an axial compressor having three axial stages and one final centrifugal stage. It has further been found that the compressor according to the present invention is of particular advantage in that the constant hub flow path included serves to greatly reduce the number of blades and vanes required throughout the several successive stages.
it will thus be seen that there has been provided by the present invention a new and improved axial compressor. While one preferred embodiment of the present invention has been shown above, it will be understood that it is not intended that it will be limited to the exact form of the construction set forth since various changes may be made still within the spirit and scope of the invention.
We claim:
1. An axial flow compressor including a multistage stator housing, a multiple stage rotor mounted for rotation coaxially within said housing and having a constant diameter hub, each of said rotor stages having a plurality of blades arranged circumferentially about said hub, each of said stator stages having a plurality of vanes arranged circumferentially within said housing, all of said blades and vanes in each of said stages having an identical hub radius, camber, thickness and chord length but having a gradually diminishing height in the direction of exiting fluid flow and all of the blades in each different rotor stage having a tip radius sloping in the direction of exiting fluid flow to provide optimum flow whereby all of the blades of the various stages are of the same blade geometry except for height and tip radius and therefore can be manufactured from the same master, each of said stator stages comprising a ring member with an inner surface defined by an inwardly sloped section, an intermediate section of constant diameter and a second inwardly sloped section whereby the stator stages combine to provide a passage which reduces in cross section from entrance to exit.

Claims (1)

1. An axial flow compressor including a multistage stator housing, a multiple stage rotor mounted for rotation coaxially within said housing and having a constant diameter hub, each of said rotor stages having a plurality of blades arranged circumferentially about said hub, each of said stator stages having a plurality of vanes arranged circumferentially within said housing, all of said blades and vanes in each of said stages having an identical hub radius, camber, thickness and chord length but having a gradually diminishing height in the direction of exiting fluid flow and all of the blades in each different rotor stage having a tip radius sloping in the direction of exiting fluid flow to provide optimum flow whereby all of the blades of the various stages are of the same blade geometry except for height and tip radius and therefore can be manufactured from the same master, each of said stator stages comprising a ring member with an inner surface defined by an inwardly sloped section, an intermediate section of constant diameter and a second inwardly sloped section whereby the stator stages combine to provide a passage which reduces in cross section from entrance to exit.
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Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4132507A (en) * 1977-07-13 1979-01-02 Kabushiki Kaisha Shikutani Blowing apparatus
EP0042044A1 (en) * 1980-06-13 1981-12-23 M.A.N. MASCHINENFABRIK AUGSBURG-NÜRNBERG Aktiengesellschaft Axial-flow compressor with displaced surge limit
US4693669A (en) * 1985-03-29 1987-09-15 Rogers Sr Leroy K Supercharger for automobile engines
EP0675290A2 (en) * 1994-03-28 1995-10-04 Research Institute Of Advanced Material Gas-Generator, Ltd. Axial flow compressor
CN1058774C (en) * 1994-12-14 2000-11-22 株式会社日立制作所 Axial-flow blower with guiding in channel
EP1382797A2 (en) * 2002-07-20 2004-01-21 Rolls-Royce Deutschland Ltd & Co KG Fluid machine flow path with increased stage contraction ratios
US20060182626A1 (en) * 2004-11-04 2006-08-17 Del Valle Bravo Facundo Axial flow supercharger and fluid compression machine
US20080232949A1 (en) * 2004-01-22 2008-09-25 Siemens Aktiengesellschaft Turbomachine Having an Axially Displaceable Rotor
US8869504B1 (en) 2013-11-22 2014-10-28 United Technologies Corporation Geared turbofan engine gearbox arrangement
US9353754B2 (en) 2012-03-13 2016-05-31 Embry-Riddle Aeronautical University, Inc. Multi-stage axial compressor with counter-rotation using accessory drive
US9534608B2 (en) 2012-02-17 2017-01-03 Embry-Riddle Aeronautical University, Inc. Multi-stage axial compressor with counter-rotation
EP3287640A1 (en) * 2016-08-26 2018-02-28 Rolls-Royce Deutschland Ltd & Co KG Fluid flow machine with high performance
US10280934B2 (en) * 2015-09-16 2019-05-07 MTU Aero Engines AG Gas turbine compressor stage
US11655757B2 (en) 2021-07-30 2023-05-23 Rolls-Royce North American Technologies Inc. Modular multistage compressor system for gas turbine engines
US11879386B2 (en) 2022-03-11 2024-01-23 Rolls-Royce North American Technologies Inc. Modular multistage turbine system for gas turbine engines

Citations (8)

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US2151699A (en) * 1938-07-28 1939-03-28 John N Heiner Casing for turbines
US2540968A (en) * 1948-12-23 1951-02-06 Hamilton Thomas Corp Bearing structure for pump shafts
US2622790A (en) * 1946-02-25 1952-12-23 Power Jets Res & Dev Ltd Bladed stator assembly primarily for axial flow compressors
US2706451A (en) * 1948-10-20 1955-04-19 Mayer-Ortiz Carlos Axial flow pump
US3012308A (en) * 1957-08-12 1961-12-12 Joy Mfg Co Method of making blade structures
US3112866A (en) * 1961-07-05 1963-12-03 Gen Dynamics Corp Compressor blade structure
US3241493A (en) * 1964-05-04 1966-03-22 Cascade Corp Pump impeller
US3493169A (en) * 1968-06-03 1970-02-03 United Aircraft Corp Bleed for compressor

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2151699A (en) * 1938-07-28 1939-03-28 John N Heiner Casing for turbines
US2622790A (en) * 1946-02-25 1952-12-23 Power Jets Res & Dev Ltd Bladed stator assembly primarily for axial flow compressors
US2706451A (en) * 1948-10-20 1955-04-19 Mayer-Ortiz Carlos Axial flow pump
US2540968A (en) * 1948-12-23 1951-02-06 Hamilton Thomas Corp Bearing structure for pump shafts
US3012308A (en) * 1957-08-12 1961-12-12 Joy Mfg Co Method of making blade structures
US3112866A (en) * 1961-07-05 1963-12-03 Gen Dynamics Corp Compressor blade structure
US3241493A (en) * 1964-05-04 1966-03-22 Cascade Corp Pump impeller
US3493169A (en) * 1968-06-03 1970-02-03 United Aircraft Corp Bleed for compressor

Cited By (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4132507A (en) * 1977-07-13 1979-01-02 Kabushiki Kaisha Shikutani Blowing apparatus
EP0042044A1 (en) * 1980-06-13 1981-12-23 M.A.N. MASCHINENFABRIK AUGSBURG-NÜRNBERG Aktiengesellschaft Axial-flow compressor with displaced surge limit
US4693669A (en) * 1985-03-29 1987-09-15 Rogers Sr Leroy K Supercharger for automobile engines
EP0675290A2 (en) * 1994-03-28 1995-10-04 Research Institute Of Advanced Material Gas-Generator, Ltd. Axial flow compressor
EP0675290A3 (en) * 1994-03-28 1997-06-25 Res Inst Of Advanced Material Axial flow compressor.
CN1058774C (en) * 1994-12-14 2000-11-22 株式会社日立制作所 Axial-flow blower with guiding in channel
EP1382797A2 (en) * 2002-07-20 2004-01-21 Rolls-Royce Deutschland Ltd & Co KG Fluid machine flow path with increased stage contraction ratios
EP1382797A3 (en) * 2002-07-20 2005-01-12 Rolls-Royce Deutschland Ltd & Co KG Fluid machine flow path with increased stage contraction ratios
US7559741B2 (en) * 2004-01-22 2009-07-14 Siemens Aktiengesellschaft Turbomachine having an axially displaceable rotor
US20080232949A1 (en) * 2004-01-22 2008-09-25 Siemens Aktiengesellschaft Turbomachine Having an Axially Displaceable Rotor
US7478629B2 (en) * 2004-11-04 2009-01-20 Del Valle Bravo Facundo Axial flow supercharger and fluid compression machine
US20060182626A1 (en) * 2004-11-04 2006-08-17 Del Valle Bravo Facundo Axial flow supercharger and fluid compression machine
US9534608B2 (en) 2012-02-17 2017-01-03 Embry-Riddle Aeronautical University, Inc. Multi-stage axial compressor with counter-rotation
US9353754B2 (en) 2012-03-13 2016-05-31 Embry-Riddle Aeronautical University, Inc. Multi-stage axial compressor with counter-rotation using accessory drive
US8869504B1 (en) 2013-11-22 2014-10-28 United Technologies Corporation Geared turbofan engine gearbox arrangement
US9500126B2 (en) 2013-11-22 2016-11-22 United Technologies Corporation Geared turbofan engine gearbox arrangement
US10578018B2 (en) 2013-11-22 2020-03-03 United Technologies Corporation Geared turbofan engine gearbox arrangement
US10280934B2 (en) * 2015-09-16 2019-05-07 MTU Aero Engines AG Gas turbine compressor stage
EP3287640A1 (en) * 2016-08-26 2018-02-28 Rolls-Royce Deutschland Ltd & Co KG Fluid flow machine with high performance
US10378545B2 (en) 2016-08-26 2019-08-13 Rolls-Royce Deutschland Ltd & Co Kg Fluid flow machine with high performance
US11655757B2 (en) 2021-07-30 2023-05-23 Rolls-Royce North American Technologies Inc. Modular multistage compressor system for gas turbine engines
US11879386B2 (en) 2022-03-11 2024-01-23 Rolls-Royce North American Technologies Inc. Modular multistage turbine system for gas turbine engines

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