US2962260A - Sweep back in blading - Google Patents

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Publication number
US2962260A
US2962260A US474806A US47480654A US2962260A US 2962260 A US2962260 A US 2962260A US 474806 A US474806 A US 474806A US 47480654 A US47480654 A US 47480654A US 2962260 A US2962260 A US 2962260A
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angle
blades
row
fluid
point
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US474806A
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John R Foley
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United Technologies Corp
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United Technologies Corp
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    • 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
    • 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
    • F04D29/54Fluid-guiding means, e.g. diffusers
    • F04D29/541Specially adapted for elastic fluid pumps
    • F04D29/542Bladed diffusers
    • F04D29/544Blade shapes
    • 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
    • Y02T50/67Relevant aircraft propulsion technologies
    • Y02T50/673Improving the blades aerodynamics
    • 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
    • Y10TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10STECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10S416/00Fluid reaction surfaces, i.e. impellers
    • Y10S416/02Formulas of curves

Description

.Nov. 29,1960 J. R. FOLEY 2,962,260

SWEEP BACK IN BLADING Filed Dec. 13. 1954 INVENTOR JOHN R. FOLEY A TTORNE) States Patent O SWEEP BACK IN BLADING john R Foley, Manchester, Cnn., assignor to United Aircraft Corporatim], East Hartford, Conn., a corporation of Delaware Filed Dec. 1'3, 1954, Ser. No. 474,806

7 Claims. (Cl. 253-78) This invention relates to blading turbo machinery, and particularly to an arrangement by which to obtain sweepback of the leading edge of the vane or blade with respect to the flow of fluid approaching the blade.

One feature of the invention is the realization of the aerodynamic or hydrodynamic advantages of sweepback by leaning the vanes or blades in a tangential direction such that the leading edges of the blades extend at an acute angle to a radius passing through a point in the leading edge. Another feature is the strengthening of the stator assembly by positioning the vanes or blades oblique- 1y to the radius but within the radial plane at right angles to the axis of the row of vanes. More specifically, one feature is the leaning of the blade in a tangential direction such that the leading edge at any one point makes an angle with the radius at the same point which bears a direct and predetermined relation to the angle of approach of the fluid to the blade at that point.

One particular feature is the leaning of stator or rotor blades in an axial flow compressor or turbine for the purpose of obtaining an effective sweepback with respect to the gas approaching the blades.

Other objects and advantages will be apparent from the specification and claims, and from the accompanying drawings which illustrate an embodiment of the invention.

Fig. 1 is an elevation of a part of a row of stator vanes or blades for an axial flow compressor.

Fig. 2 is a sectional view along the line 22 of Fig. 1.

Fig. 3 is a sectional view along the line 33 of Fig. 1.

Fig. 4 is a diagram showing the relative angles in a three dimensional arrangement.

The invention is shown as applied to a row of stator blades for an axial flow compressor. The stator vanes or blades 2, which form the row, extend between an outer shroud 4 and an inner shroud 6 forming a supporting structure and, as shown, these vanes extend at an angle which is oblique to a radial line passing through the axis of the row of vanes and through a point on the leading edge of the vane. That is to say, with reference to the particular vane 2a, the leading edge 8 of this vane forms an acute angle with a radius or radial line 10 passing through the axis of the row of vanes and through a point in the leading edge of the vanes. The inner and outer shrouds define the inner and outer boundaries of the fluid path past the row of blades. The invention resides in obtaining effective sweepback with respect to the air or gas approaching the vane by leaning the vane in a tangential direction, thereby leaving the leading edge of the vanes in a substantially radial plane at right angles to the axis of the row of vanes, as best shown in Fig. 3, the radial plane being represented by the dotted line 12.

The particular advantage of obtaining the sweepback in this way, is that it does not increase the axial length of the compressor or turbine in which the invention is incorporated. It also permits, by providing an effective sweepback, operation of the device at higher critical Mach numbers and with a lower pressure loss through each row of vanes. The arrangement of the vanes at an angle to the radius, as above described, will, when incorporated in a stator assembly, increase the strength of the stator with respect to the aerodynamic forces acting thereon.

To obtain the most eflective sweepback, it has been found that the angle that the leading edge of'the vane makes with a radial line at any point on the vane, for example, the point P, bears a direct relationship to the angle or" the air or gas stream approaching the blade. That is to say, where the effective angle of sweep or the desired angle of sweep between the gas approach angle and the leading edge of the vane is established, the desired angle of lean in a tangential direction can be determined. As shown in Fig. 4, the radial line 10 is represented by the line RR passing through the point P lying in the leading edge of the vane. The line CD in Fig. 4 represents a line tangentially of the row of blades at the point P and lying in the radial plane defined by the leading edges of the vanes and at right angles to the axis of the row of vanes. Thus, the plane T of Fig. 4 is a plane defined by the intersecting lines RR and CD and represents the radial plane at right angles to the axis in which the point P is located.

Fig. 4 also includes the line EF which represents an axial line, as shown in Fig. 3, passing through the point P and parallel to the axis of the row of stator blades. The plane defined by the lines RR and EF is the plane S which represents the axial or longitudinal plane passing through the point P and also the axis of the row of blades.

The third plane W iri Fig. 4 is defined by the intersection of the lines CD and EF and is a tangential plane passing through the point P and at right angles to a radial line from point P to the axis of the row of vanes.

The direction of approach of the air or gas with respect to the row of vanes is represented by the arrow V in the plane W in Figs. 2 and 4 and the angle of the gas approach with respect to the radial plane T is the angle 5. This angle, as will be apparent, is in the tangential plane W and is determinate for all radial points, the structure upstream of the row of vanes being'established. The angle (1), as above referred to, lies in the plane T. The effective sweep angle 0', which is the angle in space between the leading edge 8 andthe gas approach direction V, bears the direct relationship to the other angles which is expressed as follows:

cos a:cos s sin or if the angle 4: is to be determined the equation would be expressed:

cos a cos B sin 4;:

M and APZ= A1 1, (sin 0) S111 0' From the above, it will be apparent that by knowing the critical Mach number and pressure loss in a blade having a radial leading edge, it is possible, by knowing the desired critical Mach number and pressure loss, to'determine the necessary eifective angle of sweep and thereby the necessary angle that the blade must lean in a tangential directionin order to obtain the desired results. It will be understood that the leading edge of the blade is not necessarily located in'a radial plane and that an efiective sweep angle will be obtained by leaning the blade tangentially even. if the leading edge of the blade has, a longitudinal component.

It will be apparent that if the leading edge is a straight line, the effective angle of sweep will vary between the root and tip of each blade since the angle between the leading edge and a radius through a point on the leading edge will be greater where the radius intersects the blade at a point adjacent the root than where the radius intersects the blade at a point adjacent the tip. Accordingly, the leading edge of the blade may be curved so that the angle and, accordingly, the angle provide the optimum benefits at all selected points along the leading edge.

Although the description above given has been directed to a stator assembly for an axial flow compressor, it will be understood that the same benefits are obtainable in a rotating row of blading providing that the centrifugal eflect on leaned rotor blades can be overcome.

For a selected angle of lean qb, the effect is maximum for small angles of the gas approach velocity, that is, where the angle s is small. Accordingly, as the angle of approach of the gas becomes more nearly axial, the effective sweep becomes less and where the gas approach is entirely axial, there is no effective sweep. The invention accordingly, has no pratical application except where the gas approaching the row of blades has a tangential component or in other words, where there is a swirl in the gas approaching the row of blades. This will most frequently occurin multistage compressors and turbines and in other than the first stage. It is well known that a fluid being compressed in a multistage compressor or being expanded through a turbine has a swirl imparted to it by each row of stationary vanes or rotating blades and accordingly after passing the first turning vanes in a compressor the swirl existing in the fluid causes the fluid to approach the plane of the leading edges of a row of blades or vanes at an acute angle which has been commonly referred to as the angle of approach.

It is to be understood that the invention is not limited to the specific embodiment herein illustrated and described, but may be used in other ways without departure from its spirit as defined by the following claims.

I claim:

1. In a turbo machine or like device, a row of blades arranged in a ring and having their leading edges located substantially in a radial plane at right angles to the axis of the ring and in the fluid path through the devices and supporting structure for said blades located at one end at least of each of the blades in the row, each of said leading edges extending at an oblique angle to a radius through said leading edge to produce an effective sweepback with respect to the angle of approach of the fluid in said fluid path to said blade, the relationship of the oblique angle with respect to the angle of approach of the fluid ,9 and the angle of sweepback 0' being approximately cos a=cos 3 sin 5 2. In a turbo machine or like device, an annular row of blades which extend between the inner and outer boundaries of an annular fluid path, and supporting structure for said row of blades located substantially in one of the boundaries of said fluid path, the leading edge of each of said blades at any point therein extending at an oblique angle 5 to a radius through said point to provide an effective sweepback with respect to the angle of approach of the fluid in said path at said point, the relationship of the oblique angle 5 with respect to the angle of approach of the fluid ,9 and the angle of sweepback a' being approximately cos a=oos ,8 sin 3. In an axial flow compressor or turbine, a row of blades arranged in a ring and inner and outer supporting shrouds at opposite ends of the blades of the row, said shrouds defining an annular path for the flow of gas through the device, each of said blades in the row having its leading edge located substantially in a radial plane,

the gas flow in said path having a swirl about the axis of the gas path, and supporting structure for said blades located at one end at least of each of the blades in the row, each of said leading edges extending at an oblique angle & to a radius through said leading edge to produce an eflective sweepback with respect to the angle of approach of the fluid in said fluid path to said blade, the relationship of the oblique angle with respect to the angle of approach of the fluid 13 and the angle of sweepback if being approximately cos 6:608 (3 sin 4. In an axial flow compressor or turbine, a row of blades arranged in a ring and defining an annular path for the flow of gas through the device, each of said blades in the row having its leading edge located substantially in a radial plane at right angles to the axis of the ring, the gas flow in said path having a swirl about the axis of the gas path, and supporting structure for said row of blades located substantially in one of the boundaries of said fluid path, the leading edge of each of said blades at any point therein extending at an oblique angle to a radius through said point to provide an effective sweepback with respect to the angle of approach of the fluid in said path at said point, the relationship of the oblique angle as with respect to the angle of approach of the fluid l8 and the angle of sweepback abeing approximately 5. In a turbo machine or like device, a row of blades arranged in a ring and having the leading edges of the blades located substantially in a radial plane at right angles to the axis of the ring and in the fluid path through the device, the fluid path being an annulus defined by said blades and in which the fluid has a swirl such that the fluid approaches each blade of the ring at an acute angle with respect to the plane of the ring, the angle of approach of the fluid with respect to the plane being determinate at all radial points on each of the blades, the velocity of the fluid past the blades being such that a predetermined angle of sweep back of the leading edges of the blades is desirable, supporting structure for the blades in said row, said structure being located at one end at least of each of the blades in the row, each of said leading edges extending at an oblique angle to a radius through said leading edge such that at any point in the leading edge of the blade the product of the-sine of said oblique angle and the cosine of the angle of approach of the fluid at that point is approximately equal to the cosine of the predetermined angle of sweep back.

6. In a turbo machine or like device, an annular row of blades which extend between the inner and outer boundaries of an annular fluid path and supporting structure for the blades in said row, said structure being located substantially in one of the boundaries of said fluid path, the device being so arranged that the fluid in the path has a swirl, with the fluid approaching each blade of the row at an acute angle with respect to the plane of the row of blades, the angle of approach of the fluid with respect to the plane being determinate at all radial points on each of the blades and the velocity of the fluid past the blades being such that a predetermined angle of sweep back of the leading edges of the blades is desirable, each of said leading edges at any point thereon extending at a selected oblique angle to a radius passing through said leading edge at that point such that the sine of the oblique angle is substantially equal to the cosine of the predetermined angle of sweep back divided by the cosine of the angle of approach of the fluid.

7. In an axial flow compressor or turbine device, a row of blades arranged in a ring and having their leading edges located substantially in a radial plane at right angles to the axis of the ring and in the fluid path through the device, inner and outer supporting shrouds at opposite ends of the blades of the row, said shrouds defining, with said blades, the annular path for the flow of fluid through the device, the fluid in said path having a swirl such that the fluid approaches each blade of the ring at an acute angle with respect to the plane of the ring, the angle of approach of the fluid with respect to the plane being determinate at all radial points on each of the blades, and the velocity of the fluid past the blades being such that a predetermined angle of sweep back of the leading edges of the blades is desirable, each of said leading edges at that point extending at an oblique angle to a radius through said leading edge at the point such that the product of the sine of said oblique angle and the cosine of the angle of approach of the fluid at that point is approximately equal to the cosine of the predetermined angle of sweep back.

References Cited in the file of this patent UNITED STATES PATENTS

US474806A 1954-12-13 1954-12-13 Sweep back in blading Expired - Lifetime US2962260A (en)

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Cited By (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3536414A (en) * 1968-03-06 1970-10-27 Gen Electric Vanes for turning fluid flow in an annular duct
US3883264A (en) * 1971-04-08 1975-05-13 Gadicherla V R Rao Quiet fan with non-radial elements
US3995970A (en) * 1974-09-10 1976-12-07 Mitsubishi Jukogyo Kabushiki Kaisha Axial-flow fan
US4131387A (en) * 1976-02-27 1978-12-26 General Electric Company Curved blade turbomachinery noise reduction
US4168939A (en) * 1977-09-08 1979-09-25 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Acoustically swept rotor
US4826400A (en) * 1986-12-29 1989-05-02 General Electric Company Curvilinear turbine airfoil
US5167489A (en) * 1991-04-15 1992-12-01 General Electric Company Forward swept rotor blade
EP0570106A1 (en) * 1992-05-15 1993-11-18 Gec Alsthom Limited Turbine blade assembly
US5443590A (en) * 1993-06-18 1995-08-22 General Electric Company Rotatable turbine frame
US5588618A (en) * 1994-05-04 1996-12-31 Eurocopter France Counter-torque device with rotor and flow-straightening stator, both of which are ducted, and phase modulation of the blades of the rotor, for helicopter
US5634611A (en) * 1994-05-04 1997-06-03 Eurocopter France Counter-torque device with rotor and flow straightening stator, both of which are ducted, and inclined flow-straightening vanes
US6290465B1 (en) * 1999-07-30 2001-09-18 General Electric Company Rotor blade
US6386830B1 (en) * 2001-03-13 2002-05-14 The United States Of America As Represented By The Secretary Of The Navy Quiet and efficient high-pressure fan assembly
EP1247937A1 (en) * 2001-04-04 2002-10-09 Siemens Aktiengesellschaft Gas turbine blade and gas turbine
US6554564B1 (en) 2001-11-14 2003-04-29 United Technologies Corporation Reduced noise fan exit guide vane configuration for turbofan engines
WO2005081979A2 (en) * 2004-02-23 2005-09-09 Revcor, Inc. Fan assembly and method
US20070160475A1 (en) * 2006-01-12 2007-07-12 Siemens Power Generation, Inc. Tilted turbine vane with impingement cooling
US20070183890A1 (en) * 2006-02-09 2007-08-09 Honeywell International, Inc. Leaned deswirl vanes behind a centrifugal compressor in a gas turbine engine
US20080050220A1 (en) * 2006-08-24 2008-02-28 United Technologies Corporation Leaned high pressure compressor inlet guide vane
US20100028157A1 (en) * 2008-07-30 2010-02-04 General Electric Company Wind turbine blade tip shapes
US20170002670A1 (en) * 2015-07-01 2017-01-05 General Electric Company Bulged nozzle for control of secondary flow and optimal diffuser performance
CN106907185A (en) * 2015-10-15 2017-06-30 通用电气公司 Protrusion nozzle for controlling sidestream and optimal diffuser performance

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US866068A (en) * 1907-02-08 1907-09-17 Gen Electric Nozzle and diaphragm for turbines.
GB190627409A (en) * 1906-12-01 1907-11-30 Rudolf Pawlikowski Improvements in Steam Turbines.
US1475267A (en) * 1922-07-12 1923-11-27 Gen Electric Elastic-fluid turbine
US1475212A (en) * 1922-07-12 1923-11-27 Gen Electric Elastic-fluid turbine
US1641665A (en) * 1925-11-21 1927-09-06 Gen Electric Turbine-nozzle diaphragm
FR823441A (en) * 1936-10-02 1938-01-20 Rateau Sa Apparatus for reducing noise in fans and rotary devices fluid compression
FR892412A (en) * 1942-04-11 1944-04-06 Wagner Hochdruck Dampfturbinen Steering device for steam and gas turbines and the like centrifuges
GB622078A (en) * 1947-03-11 1949-04-26 Arthur Alexander Rubbra Improvements relating to gas turbine engines
US2639886A (en) * 1950-11-17 1953-05-26 Thompson Prod Inc Shrouded wheel

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB190627409A (en) * 1906-12-01 1907-11-30 Rudolf Pawlikowski Improvements in Steam Turbines.
US866068A (en) * 1907-02-08 1907-09-17 Gen Electric Nozzle and diaphragm for turbines.
US1475267A (en) * 1922-07-12 1923-11-27 Gen Electric Elastic-fluid turbine
US1475212A (en) * 1922-07-12 1923-11-27 Gen Electric Elastic-fluid turbine
US1641665A (en) * 1925-11-21 1927-09-06 Gen Electric Turbine-nozzle diaphragm
FR823441A (en) * 1936-10-02 1938-01-20 Rateau Sa Apparatus for reducing noise in fans and rotary devices fluid compression
FR892412A (en) * 1942-04-11 1944-04-06 Wagner Hochdruck Dampfturbinen Steering device for steam and gas turbines and the like centrifuges
GB622078A (en) * 1947-03-11 1949-04-26 Arthur Alexander Rubbra Improvements relating to gas turbine engines
US2639886A (en) * 1950-11-17 1953-05-26 Thompson Prod Inc Shrouded wheel

Cited By (32)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3536414A (en) * 1968-03-06 1970-10-27 Gen Electric Vanes for turning fluid flow in an annular duct
US3883264A (en) * 1971-04-08 1975-05-13 Gadicherla V R Rao Quiet fan with non-radial elements
US3995970A (en) * 1974-09-10 1976-12-07 Mitsubishi Jukogyo Kabushiki Kaisha Axial-flow fan
US4131387A (en) * 1976-02-27 1978-12-26 General Electric Company Curved blade turbomachinery noise reduction
US4168939A (en) * 1977-09-08 1979-09-25 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Acoustically swept rotor
US4826400A (en) * 1986-12-29 1989-05-02 General Electric Company Curvilinear turbine airfoil
US5167489A (en) * 1991-04-15 1992-12-01 General Electric Company Forward swept rotor blade
US5575620A (en) * 1992-05-15 1996-11-19 Gec Alsthom Limited Turbine blade assembly
EP0570106A1 (en) * 1992-05-15 1993-11-18 Gec Alsthom Limited Turbine blade assembly
US5443590A (en) * 1993-06-18 1995-08-22 General Electric Company Rotatable turbine frame
US5588618A (en) * 1994-05-04 1996-12-31 Eurocopter France Counter-torque device with rotor and flow-straightening stator, both of which are ducted, and phase modulation of the blades of the rotor, for helicopter
US5634611A (en) * 1994-05-04 1997-06-03 Eurocopter France Counter-torque device with rotor and flow straightening stator, both of which are ducted, and inclined flow-straightening vanes
US6290465B1 (en) * 1999-07-30 2001-09-18 General Electric Company Rotor blade
SG85715A1 (en) * 1999-07-30 2002-01-15 Gen Electric Rotor blade
US6386830B1 (en) * 2001-03-13 2002-05-14 The United States Of America As Represented By The Secretary Of The Navy Quiet and efficient high-pressure fan assembly
EP1247937A1 (en) * 2001-04-04 2002-10-09 Siemens Aktiengesellschaft Gas turbine blade and gas turbine
CN100366865C (en) * 2001-04-04 2008-02-06 西门子公司 Turbine propeller and turbine engine
US6554564B1 (en) 2001-11-14 2003-04-29 United Technologies Corporation Reduced noise fan exit guide vane configuration for turbofan engines
WO2005081979A3 (en) * 2004-02-23 2008-10-09 Revcor Inc Fan assembly and method
WO2005081979A2 (en) * 2004-02-23 2005-09-09 Revcor, Inc. Fan assembly and method
US20070160475A1 (en) * 2006-01-12 2007-07-12 Siemens Power Generation, Inc. Tilted turbine vane with impingement cooling
US20070183890A1 (en) * 2006-02-09 2007-08-09 Honeywell International, Inc. Leaned deswirl vanes behind a centrifugal compressor in a gas turbine engine
EP1818511A2 (en) * 2006-02-09 2007-08-15 Honeywell International Inc. Leaned deswirl vanes behind a centrifugal compressor in a gas turbine engine
EP1818511A3 (en) * 2006-02-09 2007-12-05 Honeywell International Inc. Leaned deswirl vanes behind a centrifugal compressor in a gas turbine engine
US20080050220A1 (en) * 2006-08-24 2008-02-28 United Technologies Corporation Leaned high pressure compressor inlet guide vane
US7594794B2 (en) * 2006-08-24 2009-09-29 United Technologies Corporation Leaned high pressure compressor inlet guide vane
US20100028157A1 (en) * 2008-07-30 2010-02-04 General Electric Company Wind turbine blade tip shapes
US7854595B2 (en) * 2008-07-30 2010-12-21 General Electric Company Wind turbine blade tip shapes
US20170002670A1 (en) * 2015-07-01 2017-01-05 General Electric Company Bulged nozzle for control of secondary flow and optimal diffuser performance
CN106321156A (en) * 2015-07-01 2017-01-11 通用电气公司 Bulged nozzle for control of secondary flow and optimal diffuser performance
US10323528B2 (en) * 2015-07-01 2019-06-18 General Electric Company Bulged nozzle for control of secondary flow and optimal diffuser performance
CN106907185A (en) * 2015-10-15 2017-06-30 通用电气公司 Protrusion nozzle for controlling sidestream and optimal diffuser performance

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