US3652184A - After-guide-blading of an axial compressor - Google Patents

After-guide-blading of an axial compressor Download PDF

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US3652184A
US3652184A US36752A US3652184DA US3652184A US 3652184 A US3652184 A US 3652184A US 36752 A US36752 A US 36752A US 3652184D A US3652184D A US 3652184DA US 3652184 A US3652184 A US 3652184A
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compressor
assembly according
ring means
profile ring
walls
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US36752A
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Oswald Conrad
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Daimler Benz AG
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Daimler Benz AG
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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
    • F04D21/00—Pump involving supersonic speed of pumped fluids
    • 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
    • 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
    • F04D29/544—Blade shapes

Definitions

  • the present invention relates to an afterguide blading or guide baffles of an axial compressor, especially of a supersonic axial compressor.
  • the flow is very strongly deflected and decelerated in the afterguide blading which may lead to a separation of the flow, i.e., the flow becoming nonlaminar. This is true to a particularly large extent for supersonic axial compressors in which the deceleration is partially initiated by a perpendicular compression shock.
  • the danger of flow separation can be reduced in a known manner in that downstream of the place at which the compression shock occurs, the deceleration is separated from the deflection.
  • This separation can be attained in that one causes the housing walls to converge within the deflection area in such a manner that the flow is only deflected whereas one decelerates the deflected flow without substantial change in direction by the divergence of the walls downstream of the blading.
  • the bend or brake occurring at the transition of the converging into the diverging wall parts thereby effects a separation of the flow the more, the thicker the boundary layer of the arriving flow medium. Since the boundary layer at the housing wall and at the hub are already relatively thick as a result of the interactions in the rotor and of the compression shock in the afterguide blading, the flow separates with certainty at the bend or brake.
  • the present invention is concerned with the task to eliminate the danger of such a separation of the flow.
  • This is achieved according to the present invention in that between the walls delimiting the afterguideblading, a profile ring is arranged coaxial thereto and projecting partially into the guide blading, whose profile height increases in a first section and decreases again in a second section.
  • the cross-sectional configuration is thereby influenced in this area in such a manner that a break-free design and guidance of the walls is possible.
  • a separation of the flow is avoided thereby whereas a significant boundary layer cannot form on the relatively short, first section of the profile ring which is disposed in front or upstream ofthe maximum profile height thereof.
  • a favorable influencing of the flow can be achieved according to the present invention in that the profile ringprojects into the guide blading so far that the transition from the first to the second section thereof is located within the area of the discharge plane of the guide blades.
  • a particularly advantageous type of construction of the present invention for supersonic axial compressors resides in that the profile of the profile ring is a quadrangle with a long diagonal parallel to the compressor axis and with a short diagonal perpendicular thereto.
  • Another object of the present invention resides in an afterguide blading of an axial compressor which effectively eliminates the danger of flow separation.
  • a further object of the present invention resides in a guide blading of the type described above which not only avoids the danger of flow separation but permits a configuration of the inner and outer walls within the area of the guide blades that is substantially free of any bends or breaks.
  • a still further object of the present invention resides in an axial compressor provided with afterguide blading in the discharge from the compressor blades, which precludes the formation of a significant boundary layer.
  • FIG. 1 is a longitudinal cross-sectional view through a part of a supersonic axial compressor
  • the first section 22 of the i FIG. 2 is a plan view, unfolded into the plane of the drawing, of the illustrated rotor and afterguide blading of HO. 1;
  • F IG. 3 is a'view, on an enlarged scale, illustrating the profile of the profile ring of the afterguide blading.
  • a rotor 12 consisting of a hub disk 13 and of compressor blades 14 is conventionally supported within the housing 11 of a supersonic axial compressor.
  • An afterguide blading 15 adjoins the rotor 12 which is arranged in the annular channel 16 between an outer wall 17 and an inner wall 18 of the housing 1 l.
  • the afterguide blading 15 consists of guide blades 19 and of a profile ring 20 arranged approximately in the center of the annular channel 16 and coaxial thereto.
  • the thickness of the profile generally designated by reference numeral 21 of the profile ring 20 thereby increases in a first section 22 and again decreases in a second section 23. More particularly, the
  • profile 21 has the configuration of a quandrangle with a long diagonal 24 parallel to the compressor axis and a short profile'ring 20 thereby projects so far into the guide blades l9 that the.transition 26 into the second section 23 is disposed within the area of the discharge plane of the guide blades 19.
  • the profile ring 20 leads by reason of its profile 21 which produces a convergent-divergent cross-sectional configuration of the annular channel 16, as well as by reason of its arrangement in relation to the guide blades 19 to a separation of the deceleration from the deflection of the flow in the afterguide, blading 15.
  • the flow within the area of the second, tapering section 23 of the profile ring 20 is decelerated whereas the danger of separation of the flow is strongly reduced at the outer wall 17 as well as at the inner wall 18 which can be constructed without bend or break by reason of the arrangement of the profile ring 20.
  • the walls delimiting the afterguide blading have, in the illustrated embodiment, a cylindrical configuration. However, they may also be slightly inclined or curved without changing the effect of the profile ring of the present invention.
  • the profile ring may, in addition to cross-sectional shapes that are defined by straight lines, for example, such as shown by the illustrated embodiment,.or by a rhombic configuration, also possess a drop-shaped cross-sectional shape.
  • the profiles may thereby also-be constructed asymmetrical to the longitudinal axis thereof. Instead of the center position illustrated in the described embodiment, it is also possible to arrange the profile ring 20 closer to the inner wall or to the outer wall of the annular channel 16.
  • An afterguide blading assembly for an axial compressor with a longitudinally extending axis of the type having a rotatable compressor means and inner and outer housing walls for delimiting a space immediately behind the compressor means for the passage of air and other gases leaving the compressor; said assembly comprising a plurality of radially extending guide blades arranged in said space between said inner and outer walls for smoothing the flow of the output from the compressor means, said guide blades including leading edges disposed at the forward ends thereof for initially contacting said compressor output and trailing edges disposed axially rearwardly of said leading edges for the egress of said compressor output from the guide blades, and an annular profile ring means arranged between said inner and outer walls in substantially coaxial alignment with said walls, said ring means including a first forward section having an increasing cross section in the axial direction of the flow of the compressor output and a second rearward section having a decreasing cross section in the axial direction of said flow, the most forward end of said ring means being disposed between the trailing and leading edges of said guide blade
  • cross-sectional shape of the profile ring means is substantially a quadrangle with a relatively long diagonal substantially parallel to the compressor axis and a relatively short diagonal in said transition plane substantially perpendicular to said compressor axis.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Geometry (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

An after-guide blading of an axial compressor, particularly of a supersonic axial compressor, in which a profile ring is provided between the walls defining the after-guide blading which partially projects into the guide blades; the cross-sectional configuration of the profile ring in a plane parallel to and passing through the compressor axis is such that it increases in a first section and decreases again in a second section.

Description

United States Patent Conrad Mar. 28, 1972 [54] AFTER-GUIDE-BLADING OF AN AXIAL 2,923,125 2/1960 Rainbow ..415/79 COMPRESSOR 2,720,750 10/1955 2,989,843 6/1961 [721 Invent Oswald schmlden Germany 2,704,089 5/1955 Woodworth ..415/209 [73] Assrgnee: Daimler-Benz Aktiengesellschaft, Germany FOREIGN PATENTS OR APPLICATIONS [22] May 1970 945,470 11/1948 France ..415/181 [21] Appl. No.: 36,752 546,754 3/1932 Germany ..4l6/l96 Primary Examiner-Henry F. Raduazo [30] Foreign Application Priority Data Attorneycraig Antonem & Hm
May 17, 1969 Germany ..P 19 25 172.0
[57] ABSTRACT F r An after-guide blading of an axial compressor, particularly of l 58] Field 41 6/194 l 196 5/181 209 a supersonic axial compressor, in which a profile ring is pro- 207 1 60/269 vided between the walls defining the after-guide blading which partially projects into the guide blades; the cross-sectional [56] References Cited configuration of the profile ring in a plane parallel to and passing through the compressor axis is such that it increases in UNITED STATES PATENTS a first section and decreases again in a second section. 2,538,739 H1951 Troller ..4l5/209 19 Claims, 3 Drawing Figures PATENTEDmze I972 3,652,184
I II) I5 I6 FIG. 2 I f FIG. 3 26 INVENTOR 22 23 OSWALD CONRAD ATTORNFYS AFTER-GUIDE-BLADING OF AN AXIAL COMPRESSOR The present invention relates to an afterguide blading or guide baffles of an axial compressor, especially of a supersonic axial compressor. The flow is very strongly deflected and decelerated in the afterguide blading which may lead to a separation of the flow, i.e., the flow becoming nonlaminar. This is true to a particularly large extent for supersonic axial compressors in which the deceleration is partially initiated by a perpendicular compression shock. The danger of flow separation can be reduced in a known manner in that downstream of the place at which the compression shock occurs, the deceleration is separated from the deflection. This separation can be attained in that one causes the housing walls to converge within the deflection area in such a manner that the flow is only deflected whereas one decelerates the deflected flow without substantial change in direction by the divergence of the walls downstream of the blading. The bend or brake occurring at the transition of the converging into the diverging wall parts thereby effects a separation of the flow the more, the thicker the boundary layer of the arriving flow medium. Since the boundary layer at the housing wall and at the hub are already relatively thick as a result of the interactions in the rotor and of the compression shock in the afterguide blading, the flow separates with certainty at the bend or brake.
The present invention is concerned with the task to eliminate the danger of such a separation of the flow. This is achieved according to the present invention in that between the walls delimiting the afterguideblading, a profile ring is arranged coaxial thereto and projecting partially into the guide blading, whose profile height increases in a first section and decreases again in a second section. The cross-sectional configuration is thereby influenced in this area in such a manner that a break-free design and guidance of the walls is possible. A separation of the flow is avoided thereby whereas a significant boundary layer cannot form on the relatively short, first section of the profile ring which is disposed in front or upstream ofthe maximum profile height thereof.
A favorable influencing of the flow can be achieved according to the present invention in that the profile ringprojects into the guide blading so far that the transition from the first to the second section thereof is located within the area of the discharge plane of the guide blades.
A particularly advantageous type of construction of the present invention for supersonic axial compressors resides in that the profile of the profile ring is a quadrangle with a long diagonal parallel to the compressor axis and with a short diagonal perpendicular thereto.
Accordingly, it is an object of the present invention to provide a guide bladin g of an axial compressor,especially of a supersonic axial compressor which avoids by simple means the aforementioned shortcomings and drawbacks encountered in the prior art.
Another object of the present invention resides in an afterguide blading of an axial compressor which effectively eliminates the danger of flow separation.
A further object of the present invention resides in a guide blading of the type described above which not only avoids the danger of flow separation but permits a configuration of the inner and outer walls within the area of the guide blades that is substantially free of any bends or breaks.
A still further object of the present invention resides in an axial compressor provided with afterguide blading in the discharge from the compressor blades, which precludes the formation of a significant boundary layer.
These and further objects, features and advantages of the present invention will become more obvious from the following description when taken in connection with the accompanying drawing which shows, for purposes of illustration only, one embodiment in accordance with the present invention, and wherein:
FIG. 1 is a longitudinal cross-sectional view through a part of a supersonic axial compressor;
. diagonal perpendicular thereto. The first section 22 of the i FIG. 2 is a plan view, unfolded into the plane of the drawing, of the illustrated rotor and afterguide blading of HO. 1; and
F IG. 3 is a'view, on an enlarged scale, illustrating the profile of the profile ring of the afterguide blading.
Referring now to the drawing wherein like reference numerals are used throughout the various views to designate like parts, and more particularly to FIG. 1, a rotor 12 consisting of a hub disk 13 and of compressor blades 14 is conventionally supported within the housing 11 of a supersonic axial compressor. An afterguide blading 15 adjoins the rotor 12 which is arranged in the annular channel 16 between an outer wall 17 and an inner wall 18 of the housing 1 l.
The afterguide blading 15 consists of guide blades 19 and of a profile ring 20 arranged approximately in the center of the annular channel 16 and coaxial thereto. The thickness of the profile generally designated by reference numeral 21 of the profile ring 20 thereby increases in a first section 22 and again decreases in a second section 23. More particularly, the
profile 21 has the configuration of a quandrangle with a long diagonal 24 parallel to the compressor axis and a short profile'ring 20 thereby projects so far into the guide blades l9 that the.transition 26 into the second section 23 is disposed within the area of the discharge plane of the guide blades 19.
The profile ring 20 according to the present invention leads by reason of its profile 21 which produces a convergent-divergent cross-sectional configuration of the annular channel 16, as well as by reason of its arrangement in relation to the guide blades 19 to a separation of the deceleration from the deflection of the flow in the afterguide, blading 15. As a result thereof, the flow within the area of the second, tapering section 23 of the profile ring 20 is decelerated whereas the danger of separation of the flow is strongly reduced at the outer wall 17 as well as at the inner wall 18 which can be constructed without bend or break by reason of the arrangement of the profile ring 20.
The walls delimiting the afterguide blading have, in the illustrated embodiment, a cylindrical configuration. However, they may also be slightly inclined or curved without changing the effect of the profile ring of the present invention. The profile ring may, in addition to cross-sectional shapes that are defined by straight lines, for example, such as shown by the illustrated embodiment,.or by a rhombic configuration, also possess a drop-shaped cross-sectional shape. The profiles may thereby also-be constructed asymmetrical to the longitudinal axis thereof. Instead of the center position illustrated in the described embodiment, it is also possible to arrange the profile ring 20 closer to the inner wall or to the outer wall of the annular channel 16.
Thus, while I have shown and described only one embodiment in accordance with the present invention, it is understood that the same is not limited thereto but is susceptible of numerous changes and modifications as known to those skilled in the art, and I therefore do not wish to be limited to the details shown and described herein but intend to cover all such changes and modifications as are encompassed by the scope of the appended claims.
I claim:
1. An afterguide blading assembly for an axial compressor with a longitudinally extending axis of the type having a rotatable compressor means and inner and outer housing walls for delimiting a space immediately behind the compressor means for the passage of air and other gases leaving the compressor; said assembly comprising a plurality of radially extending guide blades arranged in said space between said inner and outer walls for smoothing the flow of the output from the compressor means, said guide blades including leading edges disposed at the forward ends thereof for initially contacting said compressor output and trailing edges disposed axially rearwardly of said leading edges for the egress of said compressor output from the guide blades, and an annular profile ring means arranged between said inner and outer walls in substantially coaxial alignment with said walls, said ring means including a first forward section having an increasing cross section in the axial direction of the flow of the compressor output and a second rearward section having a decreasing cross section in the axial direction of said flow, the most forward end of said ring means being disposed between the trailing and leading edges of said guide blades.
2. An assembly according to claim 1, characterized in that the cross-sectional shape of the profile ring means is substantially a quadrangle with a relatively long diagonal substantially parallel to the compressor axis and a relatively short diagonal substantially perpendicular thereto.
3. An assembly according to claim 1, characterized in that said profile ring means is of symmetrical construction.
4. An assembly according to claim 1, characterized in that said profile ring is of asymmetrical construction.
5. An assembly according to claim 1, characterized in that said profile ring means is arranged approximately halfway between said walls.
6. An assembly according to claim 1, characterized in that said profile ring means is arranged closer to one of said two walls than to the other wall.
7. An assembly according to claim 1, characterized in that the compressor is a supersonic axial compressor.
8. An assembly according to claim 1, characterized in that said first and second sections of said ring means have a common transition plane where the maximum cross section of said ring means occurs and in that said transition plane is axially aligned with the plane containing the trailing edges of said guide blades.
9. An assembly according to claim 8, characterized in that the cross-sectional shape of the profile ring means is substantially a quadrangle with a relatively long diagonal substantially parallel to the compressor axis and a relatively short diagonal in said transition plane substantially perpendicular to said compressor axis.
10. An assembly according to claim 8, characterized in that said profile ring means is of symmetrical construction with respect to the longitudinally extending axes therethrough which are parallel to the longitudinal axis of said compressor.
11. An assembly according to claim 8, characterized in that said profile ring means is of asymmetrical construction.
12. An assembly according to claim 9, characterized in that said profile ring means is arranged approximately halfway between said walls.
13. An assembly according to claim 8, characterized in that said profile ring means is arranged closer to one of said two walls than to the other wall.
14. An assembly according to claim 8, characterized in that the compressor is a supersonic axial compressor.
15. An assembly according to claim 1, characterized in that said inner and outer walls are substantially parallel to one another and to the longitudinal axis of the compressor throughout the space containing the guide blades and the ring means.
16. An assembly according to claim 15, characterized in that said first and second sections of said ring means have a common transition plane where the maximum cross section of said ring means occurs and in that said transition plane is axially aligned with the plane containing the trailing edges of said guide blades.
17. An assembly according to claim 16, characterized in;
that the cross-sectional shape of the profile ring means is substantially a quadrangle with a relatively long diagonal substantially parallel to the compressor axis and a relatively short diagonal in said transition plane substantially perpendicular to said compressor axis.
18. An assembly according to claim 15, characterized in that said guide blades are fixed to both said inner and outer walls.
19. An assembly according tovclaim 15, characterized in that the cross-sectional shape of the profile ring means is substantially a quadrangle with a relatively long diagonal substantially parallel to the compressor axis and a relatively short diagonal substantially psrpe ndi cular thereto.

Claims (19)

1. An afterguide blading assembly for an axial compressor with a longitudinally extending axis of the type having a rotatable compressor means and inner and outer housing walls for delimiting a space immediately behind the compressor means for the passage of air and other gases leaving the compressor; said assembly comprising a plurality of radially extending guide blades arranged in said space between said inner and outer walls for smoothing the flow of the output from the compressor means, said guide blades including leading edges disposed at the forward ends thereof for initially contacting said compressor output and trailing edges disposed axially rearwardly of said leading edges for the egress of said compressor output from the guide blades, and an annular profile ring means arranged between said inner and outer walls in substantially coaxial alignment with said walls, said ring means including a first forward section having an increasing cross section in the axial direction of the flow of the compressor output and a second rearward section having a decreasing cross section in the axial direction of said flow, the most forward end of said ring means being disposed between the trailing and leading edges of said guide blades.
2. An assembly according to claim 1, characterized in that the cross-sectional shape of the profile ring means is substantially a quadrangle with a relatively long diagonal substantially parallel to the compressor axis and a relatively short diagonal substantially perpendicular thereto.
3. An assembly according to claim 1, characterized in that said profile ring means is of symmetrical construction.
4. An assembly according to claim 1, characterized in that said profile ring means is of asymmetrical construction.
5. An assembly according to claim 1, characterized in that said profile ring means is arranged approximately halfway between said walls.
6. An assembly according to claim 1, characterized in that said profile ring means is arranged closer to one of said two walls than to the other wall.
7. An assembly according to claim 1, characterized in that the compressor is a supersonic axial compressor.
8. An assembly according to claim 1, characterized in that said first and second sections of said ring means have a common transition plane where the maximum cross section of said ring means occurs and in that said transition plane is axially aligned with the plane containing the trailing edges of said guide blades.
9. An assembly according to claim 8, characterized in that the cross-sectional shape of the profile ring means is substantially a quadrangle with a relatively long diagonal substantially parallel to the compressor axis and a relatively short diagonal in said transition plane substantially perpendicular to said compressor axis.
10. An assembly according to claim 8, characterized in that said profile ring means is of symmetrical construction with respect to the longitudinally extending axes therethrough which are parallel to the longitudinal axis of said compressor.
11. An assembly according to claim 8, characterized in that said profile ring means is of asymmetrical construction.
12. An assembly according to claim 9, characterized in that said profile ring means is arranged approximately halfway between said walls.
13. An assembly according to claim 8, characterized in that said profile ring means is arranged closer to one of saId two walls than to the other wall.
14. An assembly according to claim 8, characterized in that the compressor is a supersonic axial compressor.
15. An assembly according to claim 1, characterized in that said inner and outer walls are substantially parallel to one another and to the longitudinal axis of the compressor throughout the space containing the guide blades and the ring means.
16. An assembly according to claim 15, characterized in that said first and second sections of said ring means have a common transition plane where the maximum cross section of said ring means occurs and in that said transition plane is axially aligned with the plane containing the trailing edges of said guide blades.
17. An assembly according to claim 16, characterized in that the cross-sectional shape of the profile ring means is substantially a quadrangle with a relatively long diagonal substantially parallel to the compressor axis and a relatively short diagonal in said transition plane substantially perpendicular to said compressor axis.
18. An assembly according to claim 15, characterized in that said guide blades are fixed to both said inner and outer walls.
19. An assembly according to claim 15, characterized in that the cross-sectional shape of the profile ring means is substantially a quadrangle with a relatively long diagonal substantially parallel to the compressor axis and a relatively short diagonal substantially perpendicular thereto.
US36752A 1969-05-17 1970-05-13 After-guide-blading of an axial compressor Expired - Lifetime US3652184A (en)

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DE19691925172 DE1925172B2 (en) 1969-05-17 1969-05-17 DETECTION GRID OF AN AXIAL COMPRESSOR, IN PARTICULAR OF A SUPERSONIC AXIAL COMPRESSOR

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Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3846039A (en) * 1973-10-23 1974-11-05 Stalker Corp Axial flow compressor
US4116584A (en) * 1973-10-12 1978-09-26 Gutehoffnungshutte Sterkrade Ag Device for extending the working range of axial flow compressors
US5102298A (en) * 1989-09-12 1992-04-07 Asea Brown Boveri Ltd. Axial flow turbine
US5592821A (en) * 1993-06-10 1997-01-14 Societe Nationale D'etude Et De Construction De Moteurs F'aviation S.N.E.C.M.A. Gas turbine engine having an integral guide vane and separator diffuser
EP0837247A3 (en) * 1996-10-21 1998-11-25 United Technologies Corporation Stator assembly for the flow path of a gas turbine engine
US6299410B1 (en) 1997-12-26 2001-10-09 United Technologies Corporation Method and apparatus for damping vibration in turbomachine components
WO2005081979A2 (en) * 2004-02-23 2005-09-09 Revcor, Inc. Fan assembly and method
US20060275110A1 (en) * 2004-06-01 2006-12-07 Volvo Aero Corporation Gas turbine compression system and compressor structure
GB2452297B (en) * 2007-08-30 2010-01-06 Rolls Royce Plc A compressor
US20170184053A1 (en) * 2015-12-23 2017-06-29 Rolls-Royce Plc Gas turbine engine vane splitter

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE546754C (en) * 1928-01-24 1932-03-18 Bruno Franzl Blading for steam and gas turbines with cast single blades
FR945470A (en) * 1947-02-19 1949-05-05 Supersonic flying machine
US2538739A (en) * 1946-03-27 1951-01-16 Joy Mfg Co Housing for fan and motor
US2704089A (en) * 1952-06-09 1955-03-15 Lee R Woodworth Gas turbine diffuser
US2720750A (en) * 1947-11-04 1955-10-18 Helmut R Schelp Revolving fuel injection system for jet engines and gas turbines
US2923125A (en) * 1953-12-30 1960-02-02 Armstrong Siddeley Motors Ltd Ducting structure for by-pass turbojet engines
US2989843A (en) * 1953-07-24 1961-06-27 Curtiss Wright Corp Engine for supersonic flight

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE546754C (en) * 1928-01-24 1932-03-18 Bruno Franzl Blading for steam and gas turbines with cast single blades
US2538739A (en) * 1946-03-27 1951-01-16 Joy Mfg Co Housing for fan and motor
FR945470A (en) * 1947-02-19 1949-05-05 Supersonic flying machine
US2720750A (en) * 1947-11-04 1955-10-18 Helmut R Schelp Revolving fuel injection system for jet engines and gas turbines
US2704089A (en) * 1952-06-09 1955-03-15 Lee R Woodworth Gas turbine diffuser
US2989843A (en) * 1953-07-24 1961-06-27 Curtiss Wright Corp Engine for supersonic flight
US2923125A (en) * 1953-12-30 1960-02-02 Armstrong Siddeley Motors Ltd Ducting structure for by-pass turbojet engines

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4116584A (en) * 1973-10-12 1978-09-26 Gutehoffnungshutte Sterkrade Ag Device for extending the working range of axial flow compressors
US3846039A (en) * 1973-10-23 1974-11-05 Stalker Corp Axial flow compressor
US5102298A (en) * 1989-09-12 1992-04-07 Asea Brown Boveri Ltd. Axial flow turbine
US5592821A (en) * 1993-06-10 1997-01-14 Societe Nationale D'etude Et De Construction De Moteurs F'aviation S.N.E.C.M.A. Gas turbine engine having an integral guide vane and separator diffuser
EP0837247A3 (en) * 1996-10-21 1998-11-25 United Technologies Corporation Stator assembly for the flow path of a gas turbine engine
US6299410B1 (en) 1997-12-26 2001-10-09 United Technologies Corporation Method and apparatus for damping vibration in turbomachine components
WO2005081979A2 (en) * 2004-02-23 2005-09-09 Revcor, Inc. Fan assembly and method
WO2005081979A3 (en) * 2004-02-23 2008-10-09 Revcor Inc Fan assembly and method
US20060275110A1 (en) * 2004-06-01 2006-12-07 Volvo Aero Corporation Gas turbine compression system and compressor structure
US8757965B2 (en) * 2004-06-01 2014-06-24 Volvo Aero Corporation Gas turbine compression system and compressor structure
GB2452297B (en) * 2007-08-30 2010-01-06 Rolls Royce Plc A compressor
US20170184053A1 (en) * 2015-12-23 2017-06-29 Rolls-Royce Plc Gas turbine engine vane splitter

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DE1925172C3 (en) 1978-03-02
DE1925172A1 (en) 1970-11-19

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