US20060222488A1 - Nozzle vane with two slopes - Google Patents
Nozzle vane with two slopes Download PDFInfo
- Publication number
- US20060222488A1 US20060222488A1 US10/808,275 US80827504A US2006222488A1 US 20060222488 A1 US20060222488 A1 US 20060222488A1 US 80827504 A US80827504 A US 80827504A US 2006222488 A1 US2006222488 A1 US 2006222488A1
- Authority
- US
- United States
- Prior art keywords
- vane
- intermediate portion
- slope
- tip
- stacking axis
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/141—Shape, i.e. outer, aerodynamic form
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
-
- 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
-
- Y—GENERAL 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
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- the present invention relates to the general field of gas turbine engines, and more particularly to the compressors of such engines.
- a gas turbine engine comprises a combustion part and a turbine part disposed downstream from a compression part.
- An annular passage for passing a gas flow extends axially through these various parts of the engine.
- the gas flow is compressed by the compression part prior to being mixed and burnt with fuel in the combustion part.
- the gases coming from the combustion part then pass through the turbine part so as to provide propulsion thrust and drive the turbines.
- the elements of the compression part are constrained to rotate with the turbines by a drive shaft.
- the compression part of a gas turbine engine may comprise three axial compressors so as to increase compression of the gas flow: a fan; a low pressure compressor; and a high pressure compressor.
- Each compressor is typically constituted by a rotary portion (a rotor) a stationary portion (a stator), and a casing.
- a rotor inner shroud and a stator outer shroud define the radial boundaries of the annular section of the flow of gas passing through the compressor.
- the rotor comprises a plurality of rows of moving blades which extend radially through the flow section from the inner shroud to the vicinity of the outer shroud.
- the stator comprises a plurality of rows of stationary vanes extending from the outer shroud, likewise through the flow section between the outer shroud to the inner shroud.
- Each nozzle-forming row of stationary vanes is disposed between two successive rows of moving blades of the rotor.
- the stationary vanes of the nozzle serve to guide the gas flow coming from the rows of moving blades to take up appropriate speed and direction.
- Each stationary vane is constituted by a plurality of vane sections in alignment along a stacking axis and forming the vane profile.
- the nozzle vanes In order to avoid excessive deformation of the outer shroud and to avoid breaking the vanes of the nozzle, the nozzle vanes generally have a profile with a C-shaped bend (also known as a sail-shape).
- a C-shaped bend also known as a sail-shape.
- Such a shape is characterized by the vane sections situated in the middle of the flow section being tangentially offset relative to lower and upper vane sections that are close to the inner and outer shrouds, thus serving to reduce the buckling strength of the nozzle vanes.
- a vane constituted by a stack of such sections is more flexible and can therefore absorb a fraction of the deformation energy transmitted by the inner shroud.
- sail-shaped slopes penalize the aerodynamic performance of the compressor, particularly in terms of surge margin.
- the tangential offset of the vane has the effect of reducing the angles between the blade and the outer and inner shrouds, and beyond a certain value this is aerodynamically penalizing for the compressor.
- the gas flow passing through the nozzle tends to migrate from the lower and upper sections of the vanes towards the centers thereof. This migration of flow is particularly harmful in terms of surge margin at the base of the vane (bottom sections).
- the present invention thus seeks to mitigate such drawbacks by proposing a novel shape for a nozzle vane of reduced buckling strength, but without penalizing the aerodynamic performance of the nozzle.
- the invention provides a nozzle vane for a rotary disk of a turbomachine, the vane presenting mutually orthogonal longitudinal, tangential, and radial axes, and having pressure side and suction side surfaces extending radially between a base and a tip, and longitudinally between a leading edge and a trailing edge, and a plurality of vane sections having centers of gravity in alignment along a stacking axis, said vane presenting a lower portion, an intermediate portion, and an upper portion, said lower portion extending radially between the base of the vane and a lower limit of the intermediate portion, and said upper portion extending radially between an upper limit of the intermediate portion and the tip of the vane, wherein the stacking axis presents, in the lower and upper portions, a tangential component that is substantially radial, and in the intermediate portion, a tangential component having two slopes.
- Such a stack of vane sections thus makes it possible to conserve angles between the vane and the shrouds that are favorable to the surge margin of the compressor while increasing the tangential offset of the intermediate portion of the vane in order to make the vane more flexible in buckling.
- the lower and upper portions of the vane which present a tangential component that is substantially radial, prevent the gas flow passing through the compressor from migrating excessively towards the intermediate portion, even when the vane deforms.
- the tangential offset of the intermediate portion of the vane also enables the buckling strength of the vane to be decreased.
- the tangential component of the stacking axis in said intermediate portion comprises a first slope in the direction opposite to the direction of rotation of the disk, and a second slope in the direction of rotation of said disk.
- the first slope may present an angle of inclination lying in the range 5° to 45°
- the second slope may present an angle of inclination likewise lying in the range 5° to 45°.
- the tangential component of the stacking axis in the intermediate portion of the vane occupies a radial height lying in the range 35% to 65% of the total radial height between the base and the tip of the vane.
- the tangential component of the stacking axis of the lower portion of the vane occupies a radial height lying in the range 10% to 25% of the total radial height between the base and the tip of the vane.
- the tangential component of the stacking axis of the upper portion of the vane advantageously occupies a radial height lying in the range 10% to 25% of the total radial height between the base and the tip of the vane.
- FIG. 1 is a fragmentary view in longitudinal section of a compression part of a gas turbine engine
- FIG. 2 is a fragmentary perspective view of a compression part including nozzle vanes of the invention.
- FIG. 3 is a diagram showing a tangential stacking axis for the sections of a vane of the invention.
- FIG. 1 shows a portion of a compression part 10 in a gas turbine engine.
- This compression part presents an annular gas flow section 12 which extends axially through the compression part and radially between an outer shroud 14 of a stator and an inner shroud 16 of a rotor disk.
- the inner shroud is set into rotation about a longitudinal axis 17 of the engine in the direction indicated by arrow 18 , while the outer shroud remains stationary.
- the travel direction of the gas flow passing through the compression part is represented by arrow F.
- the rotor disk comprises a plurality of rows of moving blades 20 which extend radially through the flow section 12 from the inner shroud 16 to the vicinity of the outer shroud 14 .
- Each moving blade 20 presents a root 22 of dovetail shape which is engaged in a recess 24 provided to receive it in the inner shroud 16 .
- the stator comprises a plurality of rows of nozzle vanes 26 that are secured to the outer shroud 16 and that likewise extend across the flow section between the outer shroud and the inner shroud 14 .
- Each nozzle vane 26 is made integrally with the outer shroud 16 .
- the nozzle vanes 26 could likewise have respective roots for engaging in recesses in the outer shroud.
- each nozzle vane 26 is associated with three orthogonal axes: a longitudinal axis X; a tangential axis Y; and a radial axis Z.
- the longitudinal axis X extends in the flow direction F of the gas passing through the compression part.
- the tangential axis Y extends in the direction of rotation 18 of the inner shroud 16
- the radial axis Z extends radially from the inner shroud 16 towards the outer shroud.
- Each nozzle vane 26 has a pressure side surface 28 and a suction side surface 30 extending firstly radially between a base 32 and a tip 34 , and secondly axially between a leading edge 36 and a trailing edge 38 .
- the nozzle vane 26 is also constituted by a plurality of vane sections (not shown in the figures) whose centers of gravity are stacked along a stacking axis 40 from the base 32 to the tip 34 of the vane.
- the vane is subdivided along the radial axis Z into three portions: a lower portion 42 a ; an intermediate portion 42 b ; and an upper portion 42 c .
- the lower portion 42 a extends radially from the base of the vane up to a lower limit 44 of the intermediate portion 42 b
- the upper portion 42 c extends radially from an upper limit 46 of the intermediate portion to the tip of the vane.
- FIG. 3 shows a tangential stacking axis of a FIG. 2 nozzle vane. More precisely, the axis 48 shown in this figure corresponds to a tangential component of the stacking axis 40 shown in FIG. 2 , as projected onto the plane defined by the tangential axis Y and the radial axis Z of the vane.
- the stacking axis 40 of the sections of the nozzle vane 26 present, in the lower and upper portions 42 a and 42 c of the vane, a tangential component 48 that is substantially radial, and in the intermediate portion 42 b , a tangential component 48 that presents two slopes.
- the tangential component 48 of the stacking axis 40 in the lower and upper portions 42 a and 42 c of the vane extends parallel to the radial axis Z of the vane.
- the tangential component 48 of the stacking axis of the intermediate portion 42 b of the vane presents first slope 50 a at its lower limit 44 and second slope 50 b at its upper limit 46 .
- the first slope 50 a preferably extends in the opposite direction to the direction of rotation 18 of the inner shroud 16 , while the second slope 50 b takes place in the direction of rotation 18 .
- the first slope presents an angle of inclination ⁇ with respect to axis lying in the range 5° to 45°
- the second slope presents an angle of inclination ⁇ likewise lying in the range 5° to 45°
- the first slope 50 a extends radially between the lower limit 44 of the intermediate portion 42 b of the vane to a bend point 52 in the tangential component of the stacking axis
- the second slope 50 b extends radially from said bend point to the upper limit 46 of the intermediate portion.
- the tangential component 48 of the stacking axis in the intermediate portion 42 b of the vane extends over a radial height lying in the range 35% to 65% of the total radial height between the base and the tip of said vane.
- the tangential component 48 of the stacking axis in the lower portion 42 a of the vane extends over a radial height lying in the range 10% to 25% of the total radial height between the base and the tip of said vane.
- the tangential component 48 of the stacking axis in the upper portion 42 c of the vane advantageously extends over a radial height lying in the range 10% to 25% of the total radial height between the base and the tip of said vane.
- the nozzle vane of the invention presents better ability to withstand the mechanical stresses to which it is subjected while conserving accessible aerodynamic performance.
- the reduction in the buckling strength of the vane is illustrated in particular by FIG. 3 where the dashed line shows the deformation that the vane can accept without breaking.
- the nozzle vanes absorb a large portion of the deformation energy transmitted from the inner shroud, thus making it possible to limit damage of the outer shroud.
Abstract
Description
- The present invention relates to the general field of gas turbine engines, and more particularly to the compressors of such engines.
- A gas turbine engine comprises a combustion part and a turbine part disposed downstream from a compression part. An annular passage for passing a gas flow extends axially through these various parts of the engine. The gas flow is compressed by the compression part prior to being mixed and burnt with fuel in the combustion part. The gases coming from the combustion part then pass through the turbine part so as to provide propulsion thrust and drive the turbines. The elements of the compression part are constrained to rotate with the turbines by a drive shaft.
- The compression part of a gas turbine engine may comprise three axial compressors so as to increase compression of the gas flow: a fan; a low pressure compressor; and a high pressure compressor. Each compressor is typically constituted by a rotary portion (a rotor) a stationary portion (a stator), and a casing. A rotor inner shroud and a stator outer shroud define the radial boundaries of the annular section of the flow of gas passing through the compressor. The rotor comprises a plurality of rows of moving blades which extend radially through the flow section from the inner shroud to the vicinity of the outer shroud. The stator comprises a plurality of rows of stationary vanes extending from the outer shroud, likewise through the flow section between the outer shroud to the inner shroud. Each nozzle-forming row of stationary vanes is disposed between two successive rows of moving blades of the rotor. The stationary vanes of the nozzle serve to guide the gas flow coming from the rows of moving blades to take up appropriate speed and direction. Each stationary vane is constituted by a plurality of vane sections in alignment along a stacking axis and forming the vane profile.
- In normal operation of the engine, the rotation of the shaft driving the compression part gives rise to an unbalance phenomenon. The unbalance leads to cyclical loading and vibration that the rotor communicates to the stator of the engine with significant risk of the engine being damaged. In the compressors, this unbalance phenomenon leads to orbital movement of the inner shroud due to its rotation. By the inner shroud making contact with the stationary vanes of the nozzle, this orbital movement is transmitted in the form of radial displacement which has the consequence of deforming the outer shroud to which the vanes are fixed. Furthermore, the fixed nozzle vanes subjected to such radial displacement bend and run the risk of breaking (buckling phenomenon).
- In order to avoid excessive deformation of the outer shroud and to avoid breaking the vanes of the nozzle, the nozzle vanes generally have a profile with a C-shaped bend (also known as a sail-shape). Such a shape is characterized by the vane sections situated in the middle of the flow section being tangentially offset relative to lower and upper vane sections that are close to the inner and outer shrouds, thus serving to reduce the buckling strength of the nozzle vanes. A vane constituted by a stack of such sections is more flexible and can therefore absorb a fraction of the deformation energy transmitted by the inner shroud.
- Nevertheless, sail-shaped slopes penalize the aerodynamic performance of the compressor, particularly in terms of surge margin. The tangential offset of the vane has the effect of reducing the angles between the blade and the outer and inner shrouds, and beyond a certain value this is aerodynamically penalizing for the compressor. The gas flow passing through the nozzle tends to migrate from the lower and upper sections of the vanes towards the centers thereof. This migration of flow is particularly harmful in terms of surge margin at the base of the vane (bottom sections).
- The present invention thus seeks to mitigate such drawbacks by proposing a novel shape for a nozzle vane of reduced buckling strength, but without penalizing the aerodynamic performance of the nozzle.
- To this end, the invention provides a nozzle vane for a rotary disk of a turbomachine, the vane presenting mutually orthogonal longitudinal, tangential, and radial axes, and having pressure side and suction side surfaces extending radially between a base and a tip, and longitudinally between a leading edge and a trailing edge, and a plurality of vane sections having centers of gravity in alignment along a stacking axis, said vane presenting a lower portion, an intermediate portion, and an upper portion, said lower portion extending radially between the base of the vane and a lower limit of the intermediate portion, and said upper portion extending radially between an upper limit of the intermediate portion and the tip of the vane, wherein the stacking axis presents, in the lower and upper portions, a tangential component that is substantially radial, and in the intermediate portion, a tangential component having two slopes.
- Such a stack of vane sections thus makes it possible to conserve angles between the vane and the shrouds that are favorable to the surge margin of the compressor while increasing the tangential offset of the intermediate portion of the vane in order to make the vane more flexible in buckling. The lower and upper portions of the vane, which present a tangential component that is substantially radial, prevent the gas flow passing through the compressor from migrating excessively towards the intermediate portion, even when the vane deforms. The tangential offset of the intermediate portion of the vane also enables the buckling strength of the vane to be decreased.
- Preferably, the tangential component of the stacking axis in said intermediate portion comprises a first slope in the direction opposite to the direction of rotation of the disk, and a second slope in the direction of rotation of said disk. The first slope may present an angle of inclination lying in the range 5° to 45°, and the second slope may present an angle of inclination likewise lying in the range 5° to 45°.
- Advantageously, the tangential component of the stacking axis in the intermediate portion of the vane occupies a radial height lying in the range 35% to 65% of the total radial height between the base and the tip of the vane.
- Also advantageously, the tangential component of the stacking axis of the lower portion of the vane occupies a radial height lying in the
range 10% to 25% of the total radial height between the base and the tip of the vane. Similarly, the tangential component of the stacking axis of the upper portion of the vane advantageously occupies a radial height lying in therange 10% to 25% of the total radial height between the base and the tip of the vane. - Other characteristics and advantages of the present invention appear from the following description given with reference to the accompanying drawings which show an embodiment having no limiting character. In the figures:
-
FIG. 1 is a fragmentary view in longitudinal section of a compression part of a gas turbine engine; -
FIG. 2 is a fragmentary perspective view of a compression part including nozzle vanes of the invention; and -
FIG. 3 is a diagram showing a tangential stacking axis for the sections of a vane of the invention. -
FIG. 1 shows a portion of acompression part 10 in a gas turbine engine. This compression part presents an annulargas flow section 12 which extends axially through the compression part and radially between anouter shroud 14 of a stator and aninner shroud 16 of a rotor disk. The inner shroud is set into rotation about alongitudinal axis 17 of the engine in the direction indicated byarrow 18, while the outer shroud remains stationary. The travel direction of the gas flow passing through the compression part is represented by arrow F. - The rotor disk comprises a plurality of rows of moving
blades 20 which extend radially through theflow section 12 from theinner shroud 16 to the vicinity of theouter shroud 14. Each movingblade 20 presents aroot 22 of dovetail shape which is engaged in arecess 24 provided to receive it in theinner shroud 16. The stator comprises a plurality of rows ofnozzle vanes 26 that are secured to theouter shroud 16 and that likewise extend across the flow section between the outer shroud and theinner shroud 14. Eachnozzle vane 26 is made integrally with theouter shroud 16. Alternatively, thenozzle vanes 26 could likewise have respective roots for engaging in recesses in the outer shroud. - As shown in
FIG. 2 , eachnozzle vane 26 is associated with three orthogonal axes: a longitudinal axis X; a tangential axis Y; and a radial axis Z. The longitudinal axis X extends in the flow direction F of the gas passing through the compression part. The tangential axis Y extends in the direction ofrotation 18 of theinner shroud 16, and the radial axis Z extends radially from theinner shroud 16 towards the outer shroud. - Each
nozzle vane 26 has apressure side surface 28 and asuction side surface 30 extending firstly radially between abase 32 and atip 34, and secondly axially between a leadingedge 36 and atrailing edge 38. Thenozzle vane 26 is also constituted by a plurality of vane sections (not shown in the figures) whose centers of gravity are stacked along astacking axis 40 from thebase 32 to thetip 34 of the vane. - In
FIG. 3 , the vane is subdivided along the radial axis Z into three portions: alower portion 42 a; anintermediate portion 42 b; and anupper portion 42 c. Thelower portion 42 a extends radially from the base of the vane up to alower limit 44 of theintermediate portion 42 b, and theupper portion 42 c extends radially from anupper limit 46 of the intermediate portion to the tip of the vane. -
FIG. 3 shows a tangential stacking axis of aFIG. 2 nozzle vane. More precisely, theaxis 48 shown in this figure corresponds to a tangential component of thestacking axis 40 shown inFIG. 2 , as projected onto the plane defined by the tangential axis Y and the radial axis Z of the vane. - In the invention, the
stacking axis 40 of the sections of thenozzle vane 26 present, in the lower andupper portions tangential component 48 that is substantially radial, and in theintermediate portion 42 b, atangential component 48 that presents two slopes. As shown inFIG. 3 , thetangential component 48 of thestacking axis 40 in the lower andupper portions tangential component 48 of the stacking axis of theintermediate portion 42 b of the vane presentsfirst slope 50 a at itslower limit 44 andsecond slope 50 b at itsupper limit 46. Thefirst slope 50 a preferably extends in the opposite direction to the direction ofrotation 18 of theinner shroud 16, while thesecond slope 50 b takes place in the direction ofrotation 18. -
- The
first slope 50 a extends radially between thelower limit 44 of theintermediate portion 42 b of the vane to abend point 52 in the tangential component of the stacking axis, and thesecond slope 50 b extends radially from said bend point to theupper limit 46 of the intermediate portion. Advantageously, thetangential component 48 of the stacking axis in theintermediate portion 42 b of the vane extends over a radial height lying in the range 35% to 65% of the total radial height between the base and the tip of said vane. - According to another advantageous characteristic of the invention, the
tangential component 48 of the stacking axis in thelower portion 42 a of the vane extends over a radial height lying in therange 10% to 25% of the total radial height between the base and the tip of said vane. Similarly, thetangential component 48 of the stacking axis in theupper portion 42 c of the vane advantageously extends over a radial height lying in therange 10% to 25% of the total radial height between the base and the tip of said vane. - The nozzle vane of the invention presents better ability to withstand the mechanical stresses to which it is subjected while conserving accessible aerodynamic performance. The reduction in the buckling strength of the vane is illustrated in particular by
FIG. 3 where the dashed line shows the deformation that the vane can accept without breaking. Thus, in the event of unbalance, the nozzle vanes absorb a large portion of the deformation energy transmitted from the inner shroud, thus making it possible to limit damage of the outer shroud. By avoiding transfer of excessive mechanical force to the stationary portions of the engine, it is possible to reduce the dimensions of said stationary portions. InFIG. 3 , it can also be seen that in the lower and upper portions of the vane, the deformation that the vane can accept decreases the angle between the vane and the outer and inner shrouds to a small extent only. This limitation of the angle between the vane and the shrouds prevents the flow of gas passing through the nozzle from migrating excessively towards the intermediate portion of the vane, thereby conserving acceptable aerodynamic performance.
Claims (20)
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
FR0303754A FR2853022B1 (en) | 2003-03-27 | 2003-03-27 | DOUBLE CURVED RECTIFIER DRAW |
FR0303754 | 2003-03-27 |
Publications (2)
Publication Number | Publication Date |
---|---|
US20060222488A1 true US20060222488A1 (en) | 2006-10-05 |
US7121792B1 US7121792B1 (en) | 2006-10-17 |
Family
ID=32799760
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US10/808,275 Active 2024-05-10 US7121792B1 (en) | 2003-03-27 | 2004-03-25 | Nozzle vane with two slopes |
Country Status (5)
Country | Link |
---|---|
US (1) | US7121792B1 (en) |
EP (1) | EP1462608B1 (en) |
DE (1) | DE602004001531T2 (en) |
FR (1) | FR2853022B1 (en) |
RU (1) | RU2341660C2 (en) |
Cited By (31)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JP2008157246A (en) * | 2006-12-22 | 2008-07-10 | General Electric Co <Ge> | Gas turbine engine including inclined stator vane and method for assembling the same |
US20110268575A1 (en) * | 2008-12-19 | 2011-11-03 | Volvo Aero Corporation | Spoke for a stator component, stator component and method for manufacturing a stator component |
JP2011236909A (en) * | 2010-05-11 | 2011-11-24 | General Electric Co <Ge> | Turbomachine nozzle |
US8613592B2 (en) | 2010-04-10 | 2013-12-24 | Mtu Aero Engines Gmbh | Guide blade of a turbomachine |
JP2014508895A (en) * | 2011-03-25 | 2014-04-10 | ゼネラル・エレクトリック・カンパニイ | High camber stator vane |
CN103857880A (en) * | 2011-10-13 | 2014-06-11 | 斯奈克玛 | Turbomachine centre blade comprising a curved portion |
EP3108108A4 (en) * | 2014-02-19 | 2017-03-01 | United Technologies Corporation | Gas turbine engine airfoil |
US9752439B2 (en) | 2014-02-19 | 2017-09-05 | United Technologies Corporation | Gas turbine engine airfoil |
US9777580B2 (en) | 2014-02-19 | 2017-10-03 | United Technologies Corporation | Gas turbine engine airfoil |
US10036257B2 (en) | 2014-02-19 | 2018-07-31 | United Technologies Corporation | Gas turbine engine airfoil |
CN108869396A (en) * | 2017-05-16 | 2018-11-23 | 劳斯莱斯有限公司 | Compressor aerofoil components |
US10184483B2 (en) | 2014-02-19 | 2019-01-22 | United Technologies Corporation | Gas turbine engine airfoil |
US10208765B2 (en) | 2015-01-28 | 2019-02-19 | MTU Aero Engines AG | Gas turbine axial compressor |
US10309414B2 (en) | 2014-02-19 | 2019-06-04 | United Technologies Corporation | Gas turbine engine airfoil |
US10352331B2 (en) | 2014-02-19 | 2019-07-16 | United Technologies Corporation | Gas turbine engine airfoil |
US10358925B2 (en) | 2014-02-19 | 2019-07-23 | United Technologies Corporation | Gas turbine engine airfoil |
US10370974B2 (en) | 2014-02-19 | 2019-08-06 | United Technologies Corporation | Gas turbine engine airfoil |
US10385866B2 (en) | 2014-02-19 | 2019-08-20 | United Technologies Corporation | Gas turbine engine airfoil |
US10393139B2 (en) | 2014-02-19 | 2019-08-27 | United Technologies Corporation | Gas turbine engine airfoil |
US10422226B2 (en) | 2014-02-19 | 2019-09-24 | United Technologies Corporation | Gas turbine engine airfoil |
US10465702B2 (en) | 2014-02-19 | 2019-11-05 | United Technologies Corporation | Gas turbine engine airfoil |
US10495106B2 (en) | 2014-02-19 | 2019-12-03 | United Technologies Corporation | Gas turbine engine airfoil |
US10502229B2 (en) | 2014-02-19 | 2019-12-10 | United Technologies Corporation | Gas turbine engine airfoil |
US10519971B2 (en) | 2014-02-19 | 2019-12-31 | United Technologies Corporation | Gas turbine engine airfoil |
US10550852B2 (en) | 2014-02-19 | 2020-02-04 | United Technologies Corporation | Gas turbine engine airfoil |
US10557477B2 (en) | 2014-02-19 | 2020-02-11 | United Technologies Corporation | Gas turbine engine airfoil |
US10570915B2 (en) | 2014-02-19 | 2020-02-25 | United Technologies Corporation | Gas turbine engine airfoil |
US10570916B2 (en) | 2014-02-19 | 2020-02-25 | United Technologies Corporation | Gas turbine engine airfoil |
US10584715B2 (en) | 2014-02-19 | 2020-03-10 | United Technologies Corporation | Gas turbine engine airfoil |
US10590775B2 (en) | 2014-02-19 | 2020-03-17 | United Technologies Corporation | Gas turbine engine airfoil |
US10605259B2 (en) | 2014-02-19 | 2020-03-31 | United Technologies Corporation | Gas turbine engine airfoil |
Families Citing this family (18)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US7967571B2 (en) * | 2006-11-30 | 2011-06-28 | General Electric Company | Advanced booster rotor blade |
US8292574B2 (en) | 2006-11-30 | 2012-10-23 | General Electric Company | Advanced booster system |
US8087884B2 (en) * | 2006-11-30 | 2012-01-03 | General Electric Company | Advanced booster stator vane |
US7806653B2 (en) * | 2006-12-22 | 2010-10-05 | General Electric Company | Gas turbine engines including multi-curve stator vanes and methods of assembling the same |
US7758306B2 (en) * | 2006-12-22 | 2010-07-20 | General Electric Company | Turbine assembly for a gas turbine engine and method of manufacturing the same |
GB0704426D0 (en) * | 2007-03-08 | 2007-04-18 | Rolls Royce Plc | Aerofoil members for a turbomachine |
US8075259B2 (en) * | 2009-02-13 | 2011-12-13 | United Technologies Corporation | Turbine vane airfoil with turning flow and axial/circumferential trailing edge configuration |
FR2972380A1 (en) * | 2011-03-11 | 2012-09-14 | Alstom Technology Ltd | METHOD FOR MANUFACTURING STEAM TURBINE DIAPHRAGM |
ITTO20110728A1 (en) * | 2011-08-04 | 2013-02-05 | Avio Spa | STATIC PALLETED SEGMENT OF A GAS TURBINE FOR AERONAUTICAL MOTORS |
ITTO20111009A1 (en) * | 2011-11-03 | 2013-05-04 | Avio Spa | AERODYNAMIC PROFILE OF A TURBINE |
FR2989999B1 (en) * | 2012-04-26 | 2016-01-01 | Sdmo Ind | COOLING DEVICE COMPRISING AN AXIAL FAN WITH CENTRAL FLOW RECTIFICATION AND CORRESPONDING ELECTROGEN GROUP. |
US9004850B2 (en) | 2012-04-27 | 2015-04-14 | Pratt & Whitney Canada Corp. | Twisted variable inlet guide vane |
US20140072433A1 (en) * | 2012-09-10 | 2014-03-13 | General Electric Company | Method of clocking a turbine by reshaping the turbine's downstream airfoils |
US9435221B2 (en) | 2013-08-09 | 2016-09-06 | General Electric Company | Turbomachine airfoil positioning |
FR3009589B1 (en) * | 2013-08-12 | 2015-09-04 | Snecma | TURBOMACHINE RECTIFIER BOLT |
EP3055507B1 (en) | 2013-10-08 | 2020-01-01 | United Technologies Corporation | Rotor blade with compound lean contour and corresponding gas turbine engine |
FR3040071B1 (en) | 2015-08-11 | 2020-03-27 | Safran Aircraft Engines | TURBOMACHINE ROTOR DAWN |
CN106122107B (en) * | 2016-09-05 | 2019-08-16 | 上海电气燃气轮机有限公司 | Complex bend stator blade for multi stage axial flow compressor |
Citations (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3745629A (en) * | 1972-04-12 | 1973-07-17 | Secr Defence | Method of determining optimal shapes for stator blades |
US6195983B1 (en) * | 1999-02-12 | 2001-03-06 | General Electric Company | Leaned and swept fan outlet guide vanes |
US6312219B1 (en) * | 1999-11-05 | 2001-11-06 | General Electric Company | Narrow waist vane |
US6491493B1 (en) * | 1998-06-12 | 2002-12-10 | Ebara Corporation | Turbine nozzle vane |
US6508630B2 (en) * | 2001-03-30 | 2003-01-21 | General Electric Company | Twisted stator vane |
US6554569B2 (en) * | 2001-08-17 | 2003-04-29 | General Electric Company | Compressor outlet guide vane and diffuser assembly |
Family Cites Families (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5088892A (en) * | 1990-02-07 | 1992-02-18 | United Technologies Corporation | Bowed airfoil for the compression section of a rotary machine |
US6331100B1 (en) * | 1999-12-06 | 2001-12-18 | General Electric Company | Doubled bowed compressor airfoil |
-
2003
- 2003-03-27 FR FR0303754A patent/FR2853022B1/en not_active Expired - Fee Related
-
2004
- 2004-03-11 EP EP04290660A patent/EP1462608B1/en not_active Expired - Lifetime
- 2004-03-11 DE DE602004001531T patent/DE602004001531T2/en not_active Expired - Lifetime
- 2004-03-24 RU RU2004108621/06A patent/RU2341660C2/en active
- 2004-03-25 US US10/808,275 patent/US7121792B1/en active Active
Patent Citations (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3745629A (en) * | 1972-04-12 | 1973-07-17 | Secr Defence | Method of determining optimal shapes for stator blades |
US6491493B1 (en) * | 1998-06-12 | 2002-12-10 | Ebara Corporation | Turbine nozzle vane |
US6195983B1 (en) * | 1999-02-12 | 2001-03-06 | General Electric Company | Leaned and swept fan outlet guide vanes |
US6312219B1 (en) * | 1999-11-05 | 2001-11-06 | General Electric Company | Narrow waist vane |
US6508630B2 (en) * | 2001-03-30 | 2003-01-21 | General Electric Company | Twisted stator vane |
US6554569B2 (en) * | 2001-08-17 | 2003-04-29 | General Electric Company | Compressor outlet guide vane and diffuser assembly |
Cited By (45)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JP2008157246A (en) * | 2006-12-22 | 2008-07-10 | General Electric Co <Ge> | Gas turbine engine including inclined stator vane and method for assembling the same |
US20110268575A1 (en) * | 2008-12-19 | 2011-11-03 | Volvo Aero Corporation | Spoke for a stator component, stator component and method for manufacturing a stator component |
US8613592B2 (en) | 2010-04-10 | 2013-12-24 | Mtu Aero Engines Gmbh | Guide blade of a turbomachine |
JP2011236909A (en) * | 2010-05-11 | 2011-11-24 | General Electric Co <Ge> | Turbomachine nozzle |
JP2014508895A (en) * | 2011-03-25 | 2014-04-10 | ゼネラル・エレクトリック・カンパニイ | High camber stator vane |
US10309419B2 (en) * | 2011-10-13 | 2019-06-04 | Safran Aircarft Engines | Turbomachine centre blade comprising a curved portion |
CN103857880A (en) * | 2011-10-13 | 2014-06-11 | 斯奈克玛 | Turbomachine centre blade comprising a curved portion |
US20140248144A1 (en) * | 2011-10-13 | 2014-09-04 | Snecma | Turbomachine centre blade comprising a curved portion |
JP2014528552A (en) * | 2011-10-13 | 2014-10-27 | スネクマ | Turbomachine center blade including curved part |
US10393139B2 (en) | 2014-02-19 | 2019-08-27 | United Technologies Corporation | Gas turbine engine airfoil |
US10502229B2 (en) | 2014-02-19 | 2019-12-10 | United Technologies Corporation | Gas turbine engine airfoil |
US9988908B2 (en) | 2014-02-19 | 2018-06-05 | United Technologies Corporation | Gas turbine engine airfoil |
US10036257B2 (en) | 2014-02-19 | 2018-07-31 | United Technologies Corporation | Gas turbine engine airfoil |
US11867195B2 (en) | 2014-02-19 | 2024-01-09 | Rtx Corporation | Gas turbine engine airfoil |
US10184483B2 (en) | 2014-02-19 | 2019-01-22 | United Technologies Corporation | Gas turbine engine airfoil |
US11767856B2 (en) | 2014-02-19 | 2023-09-26 | Rtx Corporation | Gas turbine engine airfoil |
US9752439B2 (en) | 2014-02-19 | 2017-09-05 | United Technologies Corporation | Gas turbine engine airfoil |
US10309414B2 (en) | 2014-02-19 | 2019-06-04 | United Technologies Corporation | Gas turbine engine airfoil |
US10352331B2 (en) | 2014-02-19 | 2019-07-16 | United Technologies Corporation | Gas turbine engine airfoil |
US10358925B2 (en) | 2014-02-19 | 2019-07-23 | United Technologies Corporation | Gas turbine engine airfoil |
US10370974B2 (en) | 2014-02-19 | 2019-08-06 | United Technologies Corporation | Gas turbine engine airfoil |
US10385866B2 (en) | 2014-02-19 | 2019-08-20 | United Technologies Corporation | Gas turbine engine airfoil |
EP3108108A4 (en) * | 2014-02-19 | 2017-03-01 | United Technologies Corporation | Gas turbine engine airfoil |
US10422226B2 (en) | 2014-02-19 | 2019-09-24 | United Technologies Corporation | Gas turbine engine airfoil |
US10465702B2 (en) | 2014-02-19 | 2019-11-05 | United Technologies Corporation | Gas turbine engine airfoil |
US10495106B2 (en) | 2014-02-19 | 2019-12-03 | United Technologies Corporation | Gas turbine engine airfoil |
US9777580B2 (en) | 2014-02-19 | 2017-10-03 | United Technologies Corporation | Gas turbine engine airfoil |
US10519971B2 (en) | 2014-02-19 | 2019-12-31 | United Technologies Corporation | Gas turbine engine airfoil |
US10550852B2 (en) | 2014-02-19 | 2020-02-04 | United Technologies Corporation | Gas turbine engine airfoil |
US10557477B2 (en) | 2014-02-19 | 2020-02-11 | United Technologies Corporation | Gas turbine engine airfoil |
US10570915B2 (en) | 2014-02-19 | 2020-02-25 | United Technologies Corporation | Gas turbine engine airfoil |
US10570916B2 (en) | 2014-02-19 | 2020-02-25 | United Technologies Corporation | Gas turbine engine airfoil |
US10584715B2 (en) | 2014-02-19 | 2020-03-10 | United Technologies Corporation | Gas turbine engine airfoil |
US10590775B2 (en) | 2014-02-19 | 2020-03-17 | United Technologies Corporation | Gas turbine engine airfoil |
US10605259B2 (en) | 2014-02-19 | 2020-03-31 | United Technologies Corporation | Gas turbine engine airfoil |
US10890195B2 (en) | 2014-02-19 | 2021-01-12 | Raytheon Technologies Corporation | Gas turbine engine airfoil |
US10914315B2 (en) | 2014-02-19 | 2021-02-09 | Raytheon Technologies Corporation | Gas turbine engine airfoil |
US11041507B2 (en) | 2014-02-19 | 2021-06-22 | Raytheon Technologies Corporation | Gas turbine engine airfoil |
US11193497B2 (en) | 2014-02-19 | 2021-12-07 | Raytheon Technologies Corporation | Gas turbine engine airfoil |
US11193496B2 (en) | 2014-02-19 | 2021-12-07 | Raytheon Technologies Corporation | Gas turbine engine airfoil |
US11209013B2 (en) | 2014-02-19 | 2021-12-28 | Raytheon Technologies Corporation | Gas turbine engine airfoil |
US11391294B2 (en) | 2014-02-19 | 2022-07-19 | Raytheon Technologies Corporation | Gas turbine engine airfoil |
US11408436B2 (en) | 2014-02-19 | 2022-08-09 | Raytheon Technologies Corporation | Gas turbine engine airfoil |
US10208765B2 (en) | 2015-01-28 | 2019-02-19 | MTU Aero Engines AG | Gas turbine axial compressor |
CN108869396A (en) * | 2017-05-16 | 2018-11-23 | 劳斯莱斯有限公司 | Compressor aerofoil components |
Also Published As
Publication number | Publication date |
---|---|
RU2341660C2 (en) | 2008-12-20 |
DE602004001531D1 (en) | 2006-08-31 |
FR2853022A1 (en) | 2004-10-01 |
FR2853022B1 (en) | 2006-07-28 |
DE602004001531T2 (en) | 2007-07-19 |
EP1462608B1 (en) | 2006-07-19 |
RU2004108621A (en) | 2005-10-20 |
EP1462608A1 (en) | 2004-09-29 |
US7121792B1 (en) | 2006-10-17 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US7121792B1 (en) | Nozzle vane with two slopes | |
US10018050B2 (en) | Turbomachine rotor blade | |
EP3378780B1 (en) | Boundary layer ingestion engine with integrally bladed fan disk | |
EP2689108B1 (en) | Compressor airfoil with tip dihedral | |
US8147207B2 (en) | Compressor blade having a ratio of leading edge sweep to leading edge dihedral in a range of 1:1 to 3:1 along the radially outer portion | |
JP6047141B2 (en) | High camber stator vane | |
US8105037B2 (en) | Endwall with leading-edge hump | |
EP1505302B1 (en) | Compressor airfoil | |
CA2333843C (en) | Fluted compressor flowpath | |
US9546555B2 (en) | Tapered part-span shroud | |
US20080118362A1 (en) | Transonic compressor rotors with non-monotonic meanline angle distributions | |
EP2378075A1 (en) | Rotor blade and corresponding gas turbine engine | |
EP2354462A2 (en) | Compressor | |
CN112983885A (en) | Shroud for a splitter and rotor airfoil of a fan of a gas turbine engine | |
US20210372288A1 (en) | Compressor stator with leading edge fillet | |
EP2738351A1 (en) | Rotor blade with tear-drop shaped part-span shroud | |
RU2727823C2 (en) | Turbomachine rotor blade, disc with blades, rotor and turbomachine | |
CN113260770B (en) | Turbine blade with maximum buckling law with high flutter margin | |
KR102376903B1 (en) | Blade, compressor and gas turbine having the same | |
CN113272520B (en) | Turbine blade having maximum thickness law with high flutter margin | |
US11220910B2 (en) | Compressor stator | |
RU2792505C2 (en) | Gas turbine engine blade made according to the rule of deflection of the blade profile with a large flutter margin |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
AS | Assignment |
Owner name: SNECMA MOTEURS, FRANCE Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:FESSOU, PHILIPPE;BLOZOVSKI, MARIO;QUINIOU, HERVE;REEL/FRAME:015137/0265 Effective date: 20040318 |
|
STCF | Information on status: patent grant |
Free format text: PATENTED CASE |
|
AS | Assignment |
Owner name: SNECMA, FRANCE Free format text: CHANGE OF NAME;ASSIGNOR:SNECMA MOTEURS;REEL/FRAME:020609/0569 Effective date: 20050512 Owner name: SNECMA,FRANCE Free format text: CHANGE OF NAME;ASSIGNOR:SNECMA MOTEURS;REEL/FRAME:020609/0569 Effective date: 20050512 |
|
FEPP | Fee payment procedure |
Free format text: PAYOR NUMBER ASSIGNED (ORIGINAL EVENT CODE: ASPN); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY |
|
FPAY | Fee payment |
Year of fee payment: 4 |
|
FPAY | Fee payment |
Year of fee payment: 8 |
|
MAFP | Maintenance fee payment |
Free format text: PAYMENT OF MAINTENANCE FEE, 12TH YEAR, LARGE ENTITY (ORIGINAL EVENT CODE: M1553) Year of fee payment: 12 |
|
AS | Assignment |
Owner name: SAFRAN AIRCRAFT ENGINES, FRANCE Free format text: CHANGE OF NAME;ASSIGNOR:SNECMA;REEL/FRAME:046479/0807 Effective date: 20160803 |
|
AS | Assignment |
Owner name: SAFRAN AIRCRAFT ENGINES, FRANCE Free format text: CORRECTIVE ASSIGNMENT TO CORRECT THE COVER SHEET TO REMOVE APPLICATION NOS. 10250419, 10786507, 10786409, 12416418, 12531115, 12996294, 12094637 12416422 PREVIOUSLY RECORDED ON REEL 046479 FRAME 0807. ASSIGNOR(S) HEREBY CONFIRMS THE CHANGE OF NAME;ASSIGNOR:SNECMA;REEL/FRAME:046939/0336 Effective date: 20160803 |