CN117365663A - Anti-fly-rotation blade, manufacturing method thereof, aeroengine and aircraft - Google Patents
Anti-fly-rotation blade, manufacturing method thereof, aeroengine and aircraft Download PDFInfo
- Publication number
- CN117365663A CN117365663A CN202210766609.4A CN202210766609A CN117365663A CN 117365663 A CN117365663 A CN 117365663A CN 202210766609 A CN202210766609 A CN 202210766609A CN 117365663 A CN117365663 A CN 117365663A
- Authority
- CN
- China
- Prior art keywords
- blade
- fly
- stacking
- section
- gravity
- 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.)
- Pending
Links
- 238000004519 manufacturing process Methods 0.000 title claims abstract description 12
- 230000005484 gravity Effects 0.000 claims abstract description 46
- 230000007704 transition Effects 0.000 claims abstract description 25
- 238000000034 method Methods 0.000 claims description 7
- 230000002265 prevention Effects 0.000 claims description 4
- 239000002131 composite material Substances 0.000 abstract description 5
- 230000007547 defect Effects 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 240000008042 Zea mays Species 0.000 description 1
- 239000002737 fuel gas Substances 0.000 description 1
- 239000012088 reference solution Substances 0.000 description 1
- 239000000243 solution Substances 0.000 description 1
- 230000008719 thickening Effects 0.000 description 1
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
- 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
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/38—Blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
- F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
- F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade
Abstract
The invention discloses a fly-over preventing blade and a manufacturing method thereof, an aeroengine and an aircraft, wherein the fly-over preventing blade comprises a forward-swept gravity center stacking section, a tail edge stacking section and a transition section, the forward-swept gravity center stacking section is formed by stacking along a forward-swept gravity center stacking line, the tail edge stacking section is formed by stacking along a tail edge stacking line, the transition section is positioned between the forward-swept gravity center stacking section and the tail edge stacking section, and the transition section is formed by smoothly transiting the forward-swept gravity center stacking section and the tail edge stacking section. The anti-fly-rotating blades are compositely overlapped on each modeling section in a forward-swept center of gravity and tail edge composite stacking mode to form the three-dimensional blade, the upper half part of the anti-fly-rotating blades are respectively stacked in two directions, the thickness of local blade shapes can be increased in a tail edge stacking mode, the local strength margin is improved, the anti-fly-rotating function is realized, and the anti-fly-rotating blades can bear larger impact load after the low-pressure turbine shaft is broken.
Description
Technical Field
The invention relates to an anti-fly-through blade, a manufacturing method thereof, an aeroengine and an aircraft.
Background
In the working process of the turbofan engine, after the low-pressure shaft is broken due to some accidental extreme factors such as manufacturing defects, large-sized foreign matters sucked in, and the like, high-pressure fuel gas still continuously flows into the low-pressure turbine to expand and do work. If no effective stopping measures are available, the low-pressure turbine losing the load can rapidly accelerate to a flying state in a very short time, the blades and the discs bear the rapidly increased centrifugal force, the blades can be broken off from the root and thrown out at a high speed far exceeding the design allowable value, and the low-pressure turbine disc can be broken to cause serious non-inclusive accidents, so that the safety of an airplane is endangered.
Disclosure of Invention
The invention aims to overcome the defect that a low-pressure turbine losing load in the prior art can rapidly accelerate to a flying state in a very short time, and a blade can be broken off from the root and thrown out at a high speed, so that the safety of an aircraft is endangered.
The invention solves the technical problems by the following technical scheme:
the present invention provides a fly-away prevention blade comprising:
a forward-swept center-of-gravity stacking section stacked along a forward-swept center-of-gravity stacking line;
a trailing edge stacking section stacked along a trailing edge stacking line;
the transition section is positioned between the forward-swept gravity center stacking section and the tail edge stacking section, and the transition section is formed by smoothly transiting the forward-swept gravity center stacking section and the tail edge stacking section.
In the scheme, the structure is adopted, the fly-over preventing blades are compositely overlapped on each modeling section in a forward-swept center of gravity and tail edge composite overlapping mode to form the three-dimensional blade, the upper half parts of the fly-over preventing blades are overlapped in two directions respectively, the thickness of local blade profiles can be increased in a tail edge overlapping mode, the local strength margin is improved, the fly-over preventing function is realized, the fly-over preventing blades can bear larger impact load after the low-pressure turbine shaft is broken, and the purpose of braking is achieved only through the blade profile adjustment technical means without other extra structures. The method for increasing the thickness of the local blade profile is adopted, compared with the whole blade of the blade, the thickness is increased, and the weight of the anti-flying blade is reduced.
Preferably, the transition section is formed by fitting two blade sections selected from the forward swept center of gravity stacking section and two blade sections selected from the trailing edge stacking section.
In the scheme, two blade sections are selected from the forward-swept center-of-gravity stacking section, two blade sections are selected from the trailing edge stacking section, and a transition section is formed by fitting the selected four blade sections, so that smooth transition, smooth curvature and smooth surface are realized among the forward-swept center-of-gravity stacking section, the transition section and the trailing edge stacking section.
Preferably, the forward swept center of gravity stacking section is located below 70% of the blade height of the anti-fly rotor blade, and the trailing edge stacking section is located above 70% of the blade height of the anti-fly rotor blade.
Preferably, the forward swept center of gravity stacking section is arranged below 70% -95% of the blade height of the anti-fly-away rotor blade.
Preferably, the anti-fly-away blade is adapted for a vane of a turbine blade.
The invention provides an aeroengine comprising a fly-away preventing blade as described above.
The present invention provides an aircraft comprising an aeroengine as described above.
The invention provides a manufacturing method of a fly-away preventing blade, which comprises the following steps:
s1, stacking a forward-swept center-of-gravity stacking section along a forward-swept center-of-gravity stacking line, and stacking a trailing edge stacking section along a trailing edge stacking line;
s2, smoothly transiting at the intersection of the sweepforward gravity center stacking section and the tail edge stacking section, and forming a transition section.
In the scheme, the fly-over preventing blades are compositely overlapped on each modeling section in a forward-swept center of gravity and tail edge composite stacking mode to form the three-dimensional blade, the upper half parts of the fly-over preventing blades are respectively stacked in two directions, the thickness of local blade patterns can be increased in a tail edge stacking mode, the local strength margin is improved, the fly-over preventing function is realized, the fly-over preventing blades can bear larger impact load after the low-pressure turbine shaft is broken, and the purpose of braking is achieved only through the blade pattern adjustment technical means without other extra structures. The method for increasing the thickness of the local blade profile is adopted, compared with the whole blade of the blade, the thickness is increased, and the weight of the anti-flying blade is reduced.
Preferably, in step S2, the transition section is formed by selecting two blade sections in the forward swept center of gravity stacking section and selecting two blade section fits in the trailing edge stacking section.
Preferably, the forward-swept center-of-gravity stacking section is stacked along the forward-swept center-of-gravity stacking line below 70% of the blade height of the anti-fly-away blade;
and stacking the trailing edge stacking section along the trailing edge stacking line above 70% of the blade height of the anti-fly vane.
Preferably, the forward sweep center of gravity stacking section is stacked along the forward sweep center of gravity stacking line below 70% -95% of the blade height of the anti-fly-away blade.
The invention has the positive progress effects that:
the anti-fly-rotating blades are compositely overlapped on each modeling section in a forward-swept center of gravity and tail edge composite stacking mode to form the three-dimensional blade, the upper half parts of the anti-fly-rotating blades are respectively stacked in two directions, the thickness of local blade patterns can be increased in a tail edge stacking mode, the local strength margin is improved, the anti-fly-rotating function is realized, the anti-fly-rotating blades can bear larger impact load after the low-pressure turbine shaft is broken, and the purpose of braking is achieved only through the blade pattern adjustment technical means without other extra structures. The method for increasing the thickness of the local blade profile is adopted, compared with the whole blade of the blade, the thickness is increased, and the weight of the anti-flying blade is reduced.
Drawings
FIG. 1 is a schematic view of a rotor blade according to a preferred embodiment of the present invention.
FIG. 2 is a schematic view of the maximum thickness of a cross-sectional profile of a rotor blade according to a preferred embodiment of the present invention.
FIG. 3 is a graph comparing the maximum thickness distribution of the radial blade profile of each section of the anti-fly rotor blade according to the preferred embodiment of the present invention with the maximum thickness distribution of the radial blade profile of each section of the blade according to the reference scheme.
FIG. 4 is a comparison of the cross-sectional profile of a rotor blade of a preferred embodiment of the present invention at 90% of the blade height with the cross-sectional profile of a reference version at 90% of the blade height.
Fig. 5 is a graph comparing isentropic mach number distribution of a cross-sectional profile surface at 90% of the blade height of a rotor blade according to a preferred embodiment of the present invention with isentropic mach number distribution of a cross-sectional profile surface at 90% of the blade height of a blade according to a reference scheme.
Fig. 6 is a graph comparing the outlet airflow angle of the anti-fly rotor blade according to the preferred embodiment of the invention with the outlet airflow angle of the blade in the reference version.
Fig. 7 is a graph showing the comparison of the distribution of the energy loss coefficient of the anti-fly rotor blade according to the preferred embodiment of the present invention and the distribution of the energy loss coefficient of the blade according to the reference scheme.
FIG. 8 is a flow chart of a method for manufacturing a rotor blade according to a preferred embodiment of the invention.
Reference numerals illustrate:
front sweep center of gravity stacking section 1
Forward swept center of gravity stacking line 11
Trailing edge stacking section 2
Trailing edge stacking line 21
Transition section 3
Detailed Description
The invention is further illustrated by means of the following examples, which are not, however, intended to limit the scope of the invention.
The invention provides a fly-away preventing vane, as shown in figure 1. The anti-fly-away blade comprises a forward-swept gravity center stacking section 1, a tail edge stacking section 2 and a transition section 3, wherein the forward-swept gravity center stacking section 1 is formed by stacking along a forward-swept gravity center stacking line 11, the tail edge stacking section 2 is formed by stacking along a tail edge stacking line 21 close to a casing part, the transition section 3 is positioned between the forward-swept gravity center stacking section 1 and the tail edge stacking section 2, and the transition section 3 is formed by smoothly transiting the forward-swept gravity center stacking section 1 and the tail edge stacking section 2.
In this embodiment, the anti-fly-rotating blades are compositely stacked on each modeling section in a mode of "forward-swept center of gravity+trailing edge" to form a three-dimensional blade, and compared with the reference blades in the reference scheme in a single stacking mode, such as center of gravity stacking, the anti-fly-rotating blades are stacked respectively in two directions, the upper half part can increase the thickness of a local blade profile in a trailing edge stacking mode, the local strength margin is improved, the anti-fly-rotating function is realized, the anti-fly-rotating blades can bear larger impact load after the low-pressure turbine shaft is broken, and the purpose of braking is achieved only by the technical means of blade profile adjustment without other additional structures. The thickness of the local blade profile is increased, so that the weight of the anti-flying blade is reduced compared with the thickness thickening mode in the height direction of the whole blade. The diameter of the circle containing the center inside the sectional blade profile in fig. 2 is defined as the maximum thickness of the sectional blade profile of the anti-fly vane.
In order to prove the technical effect of the invention, three-dimensional viscous flow field analysis is carried out, and a comparison test is carried out on the anti-fly-rotor blade in the anti-fly-rotor scheme and the reference blades of the reference scheme so as to show that the design of the anti-fly-rotor blade meets the specified flow field conditions.
Under the anti-fly rotation scheme and the reference scheme, the values of the heights of the two blades in the radial direction are different, and the comparison of the maximum thickness distribution of the section blade profiles of the two blades is shown in figure 3, so that the anti-fly rotation blades adopt local blade height and large thickness distribution, wherein the maximum thickness of the anti-fly rotation blades is 2-4 times of the maximum thickness of the reference blades in the area of more than 50% of the blade height in the blades. In FIG. 3, the abscissa represents the maximum Thickness (Max Thickness), and the ordinate represents the radial height range (Span) ranging from 0 to 1, i.e., the percentage of the radial height of the blade.
As shown in fig. 4, a comparison of the sectional profile of the anti-fly-rotor blade at 90% of the blade height in the anti-fly-rotor solution and the sectional profile of the anti-fly-rotor blade at 90% of the blade height in the reference solution is shown.
As shown in fig. 5, a comparison graph of isentropic mach number distribution of the section blade profile surface at 90% of the blade height of the anti-fly-rotor blade in the anti-fly-rotor scheme and isentropic mach number distribution of the section blade profile surface at 90% of the blade height of the blade in the reference scheme is shown, so as to illustrate the influence on the blade profile performance after the blade is thickened. In fig. 5, the abscissa represents the percentage of the axial chord length (Xa), and the ordinate represents the isentropic mach number (Mais).
As shown in fig. 6, a comparison of the outlet air flow angle of the anti-fly vane in the anti-fly scheme and the outlet air flow angle of the vane in the reference scheme, in which the abscissa represents the outlet air flow angle (Velocity Flow Angle) and the ordinate represents the Normalized radial height range (Span Normalized), is shown with different values in the radial height. In fig. 6, the average value of the air flow angle sections of the outlets of the anti-flying blades is different by 0.14 degrees, the stage matching condition of the seven-stage low-pressure turbine is not affected, and the design of the anti-flying blades is compared with a reference scheme, so that the design of the anti-flying blades meets the specified flow field condition, and the anti-flying function can be realized.
As shown in fig. 7, a comparison graph of the energy loss coefficient distribution of the anti-fly rotor blade in the anti-fly rotation scheme and the energy loss coefficient distribution of the blade in the reference scheme, in which the abscissa represents the energy loss coefficient (Energy Loss Coefficient) and the ordinate represents the Normalized radial height range (Span Normalized), is shown with different values in the radial height.
In this embodiment, the transition section 3 is fitted by two blade sections selected from the swept-forward center of gravity stacking section 1 and two blade sections selected from the trailing edge stacking section 2. In other words, two blade sections are selected from the forward-swept center-of-gravity stacking section 1, two blade sections are selected from the trailing edge stacking section 2, and the transition section 3 is formed by fitting the selected four blade sections, so that the forward-swept center-of-gravity stacking section 1, the transition section 3 and the trailing edge stacking section 2 are smoothly transited, have smooth curvature and smooth surfaces.
In this embodiment, the forward swept center of gravity stacking segment 1 is located below 70% of the blade height of the anti-fly rotor blade and the trailing edge stacking segment 2 is located above 70% of the blade height of the anti-fly rotor blade. Preferably, the forward swept center of gravity stacking section 1 is disposed below 70% -95% of the blade height of the anti-fly rotor blade.
In the present embodiment, the anti-fly-rotation blade is applied to the stator blade of the turbine blade. Specifically, the anti-fly-away blade is suitable for one row of stationary blades in 10 th row to 14 th row in a seven-stage low pressure turbine.
The embodiment of the invention also provides an aeroengine, which comprises the anti-flying vane according to any embodiment.
The embodiment of the invention also provides an aircraft, which comprises the aeroengine.
The embodiment of the invention also provides a manufacturing method of the anti-fly-rotor blade, as shown in fig. 8, comprising the following steps:
s1, stacking a forward-swept center-of-gravity stacking section 1 along a forward-swept center-of-gravity stacking line 11, and stacking a trailing edge stacking section 2 along a trailing edge stacking line 21;
s2, smoothly transiting at the intersection of the forward-swept gravity center stacking section 1 and the tail edge stacking section 2, and forming a transition section 3.
The anti-fly-rotating blades are compositely overlapped on each modeling section in a forward-swept center of gravity and tail edge composite stacking mode to form the three-dimensional blade, the anti-fly-rotating blades are respectively stacked in two directions, the upper half part can increase the thickness of local blade patterns in a tail edge stacking mode, the local strength margin is improved, the anti-fly-rotating function is realized, the anti-fly-rotating blades can bear larger impact load after the low-pressure turbine shaft is broken, and the purpose of braking is achieved only through the blade pattern adjustment technical means without other extra structures. The method for increasing the thickness of the local blade profile is adopted, compared with the whole blade of the blade, the thickness is increased, and the weight of the anti-flying blade is reduced.
In this embodiment, in step S2, the transition section 3 is formed by selecting two blade sections in the forward swept center of gravity stacking section 1 and selecting two blade section fits in the trailing edge stacking section 2.
In the embodiment, below 70% of the blade height of the anti-fly-away blade, a forward-swept center-of-gravity stacking section 1 is stacked along a forward-swept center-of-gravity stacking line 11; the trailing edge stacking section 2 is stacked along the trailing edge stacking line 21 at a blade height of 70% or more of the anti-fly vane. Preferably, the forward swept center of gravity stacking section 1 is stacked along the forward swept center of gravity stacking line 11 below 70% -95% of the blade height of the anti-fly rotor blade.
While specific embodiments of the invention have been described above, it will be appreciated by those skilled in the art that this is by way of example only, and the scope of the invention is defined by the appended claims. Various changes and modifications to these embodiments may be made by those skilled in the art without departing from the principles and spirit of the invention, but such changes and modifications fall within the scope of the invention.
Claims (11)
1. A fly-away prevention blade, characterized in that the fly-away prevention blade comprises:
a forward-swept center-of-gravity stacking section stacked along a forward-swept center-of-gravity stacking line;
a trailing edge stacking section stacked along a trailing edge stacking line;
the transition section is positioned between the forward-swept gravity center stacking section and the tail edge stacking section, and the transition section is formed by smoothly transiting the forward-swept gravity center stacking section and the tail edge stacking section.
2. The anti-fly-away blade of claim 1, wherein the transition section is fitted by two blade sections selected from the forward swept center of gravity stacking section and two blade sections selected from the trailing edge stacking section.
3. The anti-fly-away blade of claim 2, wherein the forward swept center of gravity stacking section is below 70% of the blade height of the anti-fly-away blade and the trailing edge stacking section is above 70% of the blade height of the anti-fly-away blade.
4. A rotor blade according to claim 3, wherein the forward swept center of gravity stacking section is disposed below 70% -95% of the blade height of the rotor blade.
5. The anti-fly blade according to claim 1, wherein the anti-fly blade is adapted for use with a vane of a turbine blade.
6. An aircraft engine, characterized in that it comprises a anti-fly vane according to any one of claims 1 to 5.
7. An aircraft comprising the aeroengine of claim 6.
8. A method of manufacturing a fly-away vane, the method comprising the steps of:
s1, stacking a forward-swept center-of-gravity stacking section along a forward-swept center-of-gravity stacking line, and stacking a trailing edge stacking section along a trailing edge stacking line;
s2, smoothly transiting at the intersection of the sweepforward gravity center stacking section and the tail edge stacking section, and forming a transition section.
9. A method of manufacturing a fly-away prevention blade according to claim 8, wherein in step S2 the transition section is formed by selecting two blade sections in the forward swept center of gravity stacking section and selecting two blade section fits in the trailing edge stacking section.
10. The method of manufacturing a fly-away turning vane as set forth in claim 8, wherein the forward swept center of gravity stacking section is stacked along the forward swept center of gravity stacking line below 70% of the vane height of the fly-away turning vane;
and stacking the trailing edge stacking section along the trailing edge stacking line above 70% of the blade height of the anti-fly vane.
11. A method of manufacturing a fly-away rotor blade according to claim 8 or 10, wherein the forward swept center of gravity stacking section is stacked along the forward swept center of gravity stacking line below 70% -95% of the blade height of the fly-away rotor blade.
Priority Applications (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210766609.4A CN117365663A (en) | 2022-06-30 | 2022-06-30 | Anti-fly-rotation blade, manufacturing method thereof, aeroengine and aircraft |
| PCT/CN2023/103641 WO2024002212A1 (en) | 2022-06-30 | 2023-06-29 | Anti-spin blade and manufacturing method therefor, aviation engine, and aircraft |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210766609.4A CN117365663A (en) | 2022-06-30 | 2022-06-30 | Anti-fly-rotation blade, manufacturing method thereof, aeroengine and aircraft |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN117365663A true CN117365663A (en) | 2024-01-09 |
Family
ID=89383158
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202210766609.4A Pending CN117365663A (en) | 2022-06-30 | 2022-06-30 | Anti-fly-rotation blade, manufacturing method thereof, aeroengine and aircraft |
Country Status (2)
| Country | Link |
|---|---|
| CN (1) | CN117365663A (en) |
| WO (1) | WO2024002212A1 (en) |
Family Cites Families (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH08312303A (en) * | 1995-05-18 | 1996-11-26 | Mitsubishi Heavy Ind Ltd | Curved stacking method for axial compressor |
| CN105090098A (en) * | 2014-05-09 | 2015-11-25 | 贵州航空发动机研究所 | Transonic fan rotor blade |
| CN106570213B (en) * | 2016-10-11 | 2019-07-16 | 北京航空航天大学 | The design method and blade of variable inlet guide vane, compressor |
| FR3070448B1 (en) * | 2017-08-28 | 2019-09-06 | Safran Aircraft Engines | TURBOMACHINE BLOWER RECTIFIER DRAWER, TURBOMACHINE ASSEMBLY COMPRISING SUCH A BLADE AND TURBOMACHINE EQUIPPED WITH SAID DAUTH OR DUDIT TOGETHER |
| CN109505790B (en) * | 2018-12-28 | 2020-10-23 | 哈尔滨工业大学 | High-load high-through-flow-capacity axial flow fan |
| WO2020161943A1 (en) * | 2019-02-07 | 2020-08-13 | 株式会社Ihi | Method for designing blade for axial flow type fan, compressor and turbine, and blade obtained by means of said design |
-
2022
- 2022-06-30 CN CN202210766609.4A patent/CN117365663A/en active Pending
-
2023
- 2023-06-29 WO PCT/CN2023/103641 patent/WO2024002212A1/en unknown
Also Published As
| Publication number | Publication date |
|---|---|
| WO2024002212A1 (en) | 2024-01-04 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US11118459B2 (en) | Turbofan arrangement with blade channel variations | |
| US10697471B2 (en) | Gas turbine engine vanes | |
| EP1930598B1 (en) | Advanced booster rotor blade | |
| EP1930600B1 (en) | Advanced booster stator vane | |
| EP1930599B1 (en) | Advanced booster system | |
| US10577956B2 (en) | Gas turbine engine vanes | |
| EP3378780B1 (en) | Boundary layer ingestion engine with integrally bladed fan disk | |
| EP3489461B1 (en) | Gas turbine engine | |
| CN105736460B (en) | Axial compressor rotor incorporating non-axisymmetric hub flowpath and splitter blades | |
| CN115875085B (en) | Gas turbine engine with inlet pre-swirl feature | |
| WO2018219611A1 (en) | Compressor stator vane for axial compressors having a corrugated tip contour | |
| US11473434B2 (en) | Gas turbine engine airfoil | |
| CN117365663A (en) | Anti-fly-rotation blade, manufacturing method thereof, aeroengine and aircraft | |
| US11286779B2 (en) | Characteristic distribution for rotor blade of booster rotor | |
| US11371354B2 (en) | Characteristic distribution for rotor blade of booster rotor | |
| US20230392502A1 (en) | Fan blade with zero tip dihedral at the head | |
| US20210381385A1 (en) | Characteristic distribution for rotor blade of booster rotor | |
| CN113389599B (en) | Turbine engine with high acceleration and low blade turning airfoils | |
| GB2545909A (en) | Fan disk and gas turbine engine |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |