CN115234938A - Bypass induced-draft high-Mach-number turbine engine airflow blender and design method - Google Patents
Bypass induced-draft high-Mach-number turbine engine airflow blender and design method Download PDFInfo
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- CN115234938A CN115234938A CN202211146931.3A CN202211146931A CN115234938A CN 115234938 A CN115234938 A CN 115234938A CN 202211146931 A CN202211146931 A CN 202211146931A CN 115234938 A CN115234938 A CN 115234938A
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- lobe
- airflow
- bypass
- blender
- turbine engine
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/06—Arrangement of apertures along the flame tube
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/286—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply having fuel-air premixing devices
Abstract
The invention belongs to the technical field of aero-engines, and discloses a bypass induced draft high-Mach number turbine engine airflow blender and a design method. The airflow blender is a lobe blender, and lobes are in a fin configuration in an inner cavity of the shell; the height of the front edge of each lobe towards the main stream is gradually increased from 0 to the preset lobe content height, and meanwhile, the width of the front edge of the lobe facing the main stream is gradually increased from 0 to the preset lobe width in a bilateral symmetry mode. The airflow blender improves the puncture depth of the culvert airflow, increases the contact area of the culvert airflow and the culvert airflow, and the loss of the flow at the airflow blender is low in a vortex spraying single duct mode; the mixing and the reinforcing are carried out under the bypass bleed air opening multi-duct mode, the loss is small, and the performance requirement of the mixing device of the afterburner of the bypass bleed air high-Mach number turbine engine can be met. The design method of the airflow blender is suitable for the appearance characteristics of a turbine engine, and has clear flow and strong pertinence.
Description
Technical Field
The invention belongs to the technical field of aero-engines, and particularly relates to an airflow mixer for a bypass induced-draft high-Mach-number turbine engine and a design method.
Background
Efficient air mixing is one of important technologies for developing a high-Mach turbine engine, the inlet section of an afterburner of a bypass air-entraining high-Mach turbine engine needs uniform flow field distribution, and the mixing flow between high-temperature fuel gas at a turbine outlet and bypass airflow is very complex, so that the mass, momentum and energy exchange between high-temperature and normal-temperature, high-speed and low-speed, and fresh air and high-temperature fuel gas are involved.
The blender is an important device for blending two air flows with different pressures, temperatures and speeds; the air in the outer culvert of the engine is stably led into the inner culvert so as to ensure that two air flows are mixed to form a more uniform pressure field, temperature field and speed field, promote the fuel to form better atomization and improve the combustion efficiency and the combustion stability.
Common blenders include annular blenders and lobe blenders. The annular blender mixes by means of the shearing action of parallel air inflow of an internal culvert and an external culvert, and a mixing outlet is easy to have a relatively obvious internal culvert high-temperature area and an external culvert low-temperature area, so that the mixing effect is not ideal; the lobe blender forms several pairs of flow direction vortexes with opposite directions through two sides of the lobe, enhances the viscous shearing and blending effect, and has the advantages of high mixing degree, good stability, high efficiency and the like.
However, the lobe of the lobe blender commonly used at present is directed at the outer culvert ring and is admitted air, and the height usually can not be higher, has guaranteed the inlet air temperature of main stabilizer, but also leads to the degree of mixing poor simultaneously, and the combustion efficiency of outer culvert gas is obviously lower. When the bypass bleed-off control device is used for a bypass bleed-off high-Mach turbine engine, the loss is large in a turbojet mode, and the mixing is insufficient in a bypass bleed-off opening mode. To improve this situation, there is a need to develop a bypass bleed high mach number turbine engine air flow blender and design method.
Disclosure of Invention
The invention aims to provide a bypass bleed high-Mach turbine engine airflow blender and another technical problem to be solved is to provide a design method for the bypass bleed high-Mach turbine engine airflow blender.
According to the bypass bleed air high-Mach-number turbine engine airflow blender, the airflow blender is a lobe blender, lobes are distributed in a shell in an axial symmetry mode, and each lobe extends out of the shell through a corresponding bypass bleed air pipe; mixing gas respectively enters the shell from the bypass bleed-air pipes and flows out of the corresponding cavities to form bypass airflow of lobes, and the bypass airflow of the lobes and the bypass airflow of the turbine engine are mixed in the process of flowing along the shell; wherein each lobe is in a fin configuration within the interior cavity of the housing; the height of the front edge of each lobe facing the main flow is gradually increased from 0 to a preset lobe content height h; meanwhile, the width of the front edge of the lobe facing the main flow is increased from 0 in bilateral symmetry and gradually increased to a preset lobe width L; the wall thickness of the lobe a; the outlet plane of the lobe is vertical to the central axis of the shell, and the outer surface of the lobe and the inner wall surface of the shell are in smooth transition through an arc chamfer;
the intersection line of the lobe and the symmetrical plane of the lobe is a lobe ridge line which is a parabola and sweepback along the direction of the connotative airflow; the length of the lobe is the projection length S of the ridge line of the lobe on the inner wall surface of the shell;
the lobe is observed from the central axis of the shell to the direction of the shell, and the outer surface shape of the lobe is a two-dimensional symmetrical airfoil front edge shape.
Further, the ridge line of the lobe is replaced by a hyperbola.
Furthermore, the lobe height h is determined by the inner culvert airflow flow Q1, the outer culvert airflow flow Q2 and the airflow mixing degree X,
。
the invention relates to a design method for a bypass induced-draft high-Mach-number turbine engine airflow blender, which comprises the following steps of:
s1, determining the outer diameter and the inner diameter of a shell of an airflow blender according to the size of an outlet of a turbine engine;
s2, determining the number of lobes corresponding to the bypass air guide pipes one by one according to the number of the bypass air guide pipes of the turbine engine;
s3, calculating the height h of the lobe;
s4, determining a parabola of a lobe crest line of a lobe;
s5, determining the thickness a and the length S of the lobe;
s6, determining a two-dimensional symmetrical wing profile leading edge curve of a lobe;
and S7, determining the arc chamfering shape between the outer surface of the lobe and the inner wall surface of the shell.
The invention discloses an airflow blender for a bypass induced-draft high-Mach turbine engine, which is an airflow blender with reduced mixing enhancement loss. Aiming at the design requirements of a bypass induced-draft high-Mach-number turbine engine blender, the flow characteristics of a low-speed turbojet state and a high-speed bypass induced-draft opening state are comprehensively considered, a conventional lobe structure is optimized, a fin configuration is applied to lobe design based on a bionics theory, and the outer surface shape of a lobe is a two-dimensional symmetrical airfoil leading edge shape. The bypass airflow in the shell flows backwards along the outer surface of the lobe, so that the flow loss can be further reduced, the bypass airflow in the lobe flows backwards along the inner wall surface of the lobe, the penetration depth of the bypass airflow is improved, the contact area of the bypass airflow and the bypass airflow is increased, the flow at the blender of the high-speed turbine is enabled, and the loss is lower in a single-duct turbojet mode; the mixing enhancement is opened under the many ducts mode to the bypass bleed air, and the loss is less.
The bypass bleed high-Mach-number turbine engine airflow mixer only changes the configuration and the shape of lobes, does not generate an additional connecting mechanism, has a simple structure, does not increase the weight of an engine, can meet the performance requirement of a bypass bleed high-Mach-number turbine engine afterburner mixing device, and can also be suitable for a tandem type TBCC engine super-combustor mixing device.
The design method for the bypass induced-draft high-Mach-number turbine engine airflow blender disclosed by the invention is used for adaptively designing the matched airflow blender according to the appearance characteristics of a turbine engine, and has the characteristics of clear flow and strong pertinence.
Drawings
FIG. 1 is a schematic illustration of a typical bypass bleed air high Mach engine J58;
FIG. 2 is a cross-sectional view of the bypass bleed high Mach turbine engine airflow blender of the present invention;
FIG. 3 is a perspective view of a lobe of the present invention used in a bypass bleed high Mach turbine engine airflow blender;
FIG. 4 is a cross-sectional view of a lobe of the present invention used in a bypass bleed high Mach turbine engine airflow blender;
FIG. 5 is a side view of a lobe of the present invention used in a bypass bleed high Mach turbine engine airflow blender.
In the figure, 1, a compressor; 2. a combustion chamber; 3. a turbine; 4. an afterburner; 5. a bypass bleed pipe; 6. a housing; 7. lobes.
Detailed Description
In order to make the technical solutions and advantages of the present application clearer, the technical solutions of the present application will be described in detail with reference to the accompanying drawings, and it should be understood that the specific embodiments described herein are only some of the embodiments of the present application, and are used for explaining the present application and not limiting the present application. It should be noted that, for convenience of description, only the parts related to the present application are shown in the drawings, and other related parts may refer to general designs, and in case of conflict, the embodiments and technical features in the embodiments in the present application may be combined with each other to obtain a new embodiment.
Example 1
The embodiment of the invention will be further described with reference to a typical bypass bleed air high mach number engine J58 as shown in fig. 1, which is also suitable for the blending of incoming streams for a tandem TBCC super combustor.
When the bypass air-entraining engine is started and flies at a low speed, the bypass air-entraining pipe 5 is closed, at the moment, incoming flow is compressed by the compressor 1 and then enters the combustion chamber 2 to be mixed and combusted with fuel oil to obtain high-temperature fuel gas, the high-temperature fuel gas drives the turbine 3 to rotate and enters the afterburner 4, and the residual oxygen and the fuel are mixed and combusted again to provide larger thrust at the cost of high oil consumption and low efficiency. When the supersonic aircraft needs to fly at a speed of more than 2 to 3 times of sound velocity, the bypass bleed air engine utilizes 6 bypass bleed air pipes 5 to lead most of airflow out of the fourth stage of the air compressor 1, directly crosses the middle combustion chamber 2 and the turbine 3, is mixed with fuel oil in the afterburner 4 for combustion, and works in a stamping mode to reduce thrust loss.
As shown in fig. 2 to 5, the bypass bleed air high mach number turbine engine air flow blender of the present embodiment is a lobe blender, lobes 7 are axisymmetrically distributed in the housing 6, and each lobe 7 extends out of the housing 6 through a corresponding bypass bleed air pipe 5; the mixing gas respectively enters the shell 6 from the bypass bleed air pipes 5 and flows out from the corresponding cavities to form bypass airflow of the lobe 7, and the bypass airflow of the lobe 7 and the bypass airflow of the turbine engine are mixed in the process of flowing along the shell 6; characterized in that each lobe 7 is in the configuration of a fin in the inner cavity of the housing 6; the height of the front edge of each lobe 7 facing the main stream is gradually increased from 0 to a preset lobe content height h, meanwhile, the width of the front edge of each lobe 7 facing the main stream is symmetrical from 0 to the left and right and is gradually increased to a preset lobe width L, and the wall thickness a of each lobe 7; the outlet plane of the lobe 7 is vertical to the central axis of the shell 6, and the outer surface of the lobe 7 and the inner wall surface of the shell 6 are in smooth transition through an arc chamfer;
the intersection line of the lobe 7 and the symmetrical plane of the lobe 7 is a lobe crest line which is a parabola and sweepback along the direction of the connotative airflow; the length of the lobe 7 is the projection length S of the ridge line of the lobe on the inner wall surface of the shell 6;
when the lobe 7 is viewed from the central axis of the casing 6 toward the casing 6, the outer surface of the lobe 7 has a two-dimensional symmetrical airfoil leading edge shape.
Further, the ridge line of the lobe is replaced by a hyperbola.
Furthermore, the height h of the lobe 7 is determined by the inner bypass airflow flow Q1, the outer bypass airflow flow Q2 and the airflow mixing degree X,
。
the design method for the bypass induced-draft high-Mach turbine engine airflow blender comprises the following steps of:
s1, determining the outer diameter and the inner diameter of a shell 6 of the airflow blender according to the size of an outlet of a turbine engine;
s2, determining the number of lobes 7 corresponding to the bypass bleed air pipes 5 one to one according to the number of the bypass bleed air pipes 5 of the turbine engine;
s3, calculating the height h of the lobe 7;
s4, determining a parabola of a lobe crest line of the lobe 7;
s5, determining the thickness a and the length S of the lobe 7;
s6, determining a two-dimensional symmetrical wing profile leading edge curve of the lobe 7;
and S7, determining the arc chamfering shape between the outer surface of the lobe 7 and the inner wall surface of the shell 6.
The above embodiments are merely illustrative of the technical ideas of the present invention, and the technical ideas of the present invention are not limited thereto, so that any technical changes made in the technical solutions according to the technical ideas presented by the present invention fall within the protective scope of the present invention.
Claims (4)
1. The airflow blender is used for the bypass bleed high Mach number turbine engine and is a lobe blender, lobes (7) are distributed in a shell (6) in an axisymmetric mode, and each lobe (7) extends out of the shell (6) through a corresponding bypass bleed air pipe (5); the mixing gas respectively enters the shell (6) from the bypass bleed air pipe (5) and flows out from the corresponding cavity to form bypass airflow of the lobe (7), and the bypass airflow of the lobe (7) and the bypass airflow of the turbine engine are mixed in the flowing process along the shell (6); characterized in that each lobe (7) is in the form of a fin in the interior of the housing (6); the height of the front edge of each lobe (7) facing the main flow is gradually increased from 0 to a preset lobe content height h; meanwhile, the width of the front edge of the lobe (7) facing the main flow is increased from 0 in bilateral symmetry and gradually increased to a preset lobe width L; the wall thickness a of the lobe (7); the outlet plane of the lobe (7) is vertical to the central axis of the shell (6), and the outer surface of the lobe (7) and the inner wall surface of the shell (6) are in smooth transition through a circular arc chamfer;
the intersection line of the lobe (7) and the symmetrical plane of the lobe (7) is a lobe crest line which is a parabola and sweepbacks along the direction of the connotative airflow; the length of the lobe (7) is the projection length S of the ridge line of the lobe on the inner wall surface of the shell (6);
the lobe (7) is observed from the central axis of the shell (6) to the direction of the shell (6), and the outer surface shape of the lobe (7) is a two-dimensional symmetrical airfoil shape leading edge shape.
2. The high mach turbine engine flow blender for bypass induced air according to claim 1, wherein the lobe ridges are replaced with hyperbolic curves.
3. The bypass induced draft high mach number turbine engine airflow blender of claim 1, wherein said lobe (7) height h is determined by a bypass airflow rate Q1, a bypass airflow rate Q2 and an airflow blending degree X,
。
4. a method of designing a bypass bleed high mach number turbine engine air flow blender in accordance with any of claims 1 to 3, comprising the steps of:
s1, determining the outer diameter and the inner diameter of a shell (6) of an airflow blender according to the size of an outlet of a turbine engine;
s2, determining the number of lobes (7) corresponding to the bypass bleed air pipes (5) one by one according to the number of the bypass bleed air pipes (5) of the turbine engine;
s3, calculating the height h of the lobe (7);
s4, determining the parabola of the lobe crest line of the lobe (7);
s5, determining the thickness a and the length S of the lobe (7);
s6, determining a two-dimensional symmetrical wing profile leading edge curve of the lobe (7);
s7, determining the arc chamfering shape between the outer surface of the lobe (7) and the inner wall surface of the shell (6).
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| Application Number | Priority Date | Filing Date | Title |
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| CN202211146931.3A CN115234938B (en) | 2022-09-21 | 2022-09-21 | Bypass induced-draft high-Mach-number turbine engine airflow blender and design method |
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| CN202211146931.3A CN115234938B (en) | 2022-09-21 | 2022-09-21 | Bypass induced-draft high-Mach-number turbine engine airflow blender and design method |
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| CN115234938B CN115234938B (en) | 2022-11-25 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116379002A (en) * | 2023-06-05 | 2023-07-04 | 中国空气动力研究与发展中心空天技术研究所 | Design method of equal-rotation-speed reversing diffuser structure and diffuser structure |
Citations (18)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH08145361A (en) * | 1994-11-16 | 1996-06-07 | Ishikawajima Harima Heavy Ind Co Ltd | Fuel injection valve for gas turbine |
| US5755092A (en) * | 1995-10-18 | 1998-05-26 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A) | Exhaust nozzle mixer for a turbofan engine |
| JP2003065156A (en) * | 2001-08-29 | 2003-03-05 | Ishikawajima Harima Heavy Ind Co Ltd | Lobe mixer |
| US20030145578A1 (en) * | 2002-02-06 | 2003-08-07 | Ishikawajima-Harima Heavy Industries Co., Ltd. | Lobe mixer for jet flow |
| JP2003314368A (en) * | 2002-04-25 | 2003-11-06 | Ishikawajima Harima Heavy Ind Co Ltd | Lobe mixer for gas turbine |
| JP2005147064A (en) * | 2003-11-18 | 2005-06-09 | Ishikawajima Harima Heavy Ind Co Ltd | Lobe mixer |
| CN1942665A (en) * | 2004-02-20 | 2007-04-04 | 杰克·H·安德森 | Jet nozzle mixer |
| CN101031713A (en) * | 2004-09-29 | 2007-09-05 | Snecma固体燃料推进器公司 | Mixer for separate-flow nozzle |
| CN102588145A (en) * | 2012-03-09 | 2012-07-18 | 董国光 | Air-breathing rotary detonation wave injection and multimode ram detonation coupled circulating propulsion system |
| CN104405530A (en) * | 2014-10-16 | 2015-03-11 | 南京航空航天大学 | Exhaust noise reduction system of separated exhaust type turbofan engine |
| CN205064048U (en) * | 2015-10-13 | 2016-03-02 | 中国航天空气动力技术研究院 | Tail gas mixing device |
| CN106246408A (en) * | 2016-08-03 | 2016-12-21 | 中国科学院工程热物理研究所 | A kind of ultra-compact high blending efficiency racemization lobe mixing exhaust system structure |
| CN109253470A (en) * | 2018-09-11 | 2019-01-22 | 中国人民解放军国防科技大学 | Supersonic mixing reinforcing device |
| CN110748436A (en) * | 2019-10-16 | 2020-02-04 | 南京航空航天大学 | Double-layer cooling mixing pipe injection type infrared suppressor |
| CN112902225A (en) * | 2021-02-04 | 2021-06-04 | 厦门大学 | Multistage afterburning chamber with outer ring rotary detonation supercharged combustion chamber |
| CN113028448A (en) * | 2021-03-15 | 2021-06-25 | 中国航发沈阳发动机研究所 | Non-uniform lobe mixer for turbo-fan engine afterburner |
| CN114183272A (en) * | 2021-10-25 | 2022-03-15 | 北京航空航天大学 | Injection mixer and control method thereof |
| CN114934856A (en) * | 2022-06-02 | 2022-08-23 | 江苏科技大学 | Marine ejector waste heat recovery system |
-
2022
- 2022-09-21 CN CN202211146931.3A patent/CN115234938B/en active Active
Patent Citations (18)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH08145361A (en) * | 1994-11-16 | 1996-06-07 | Ishikawajima Harima Heavy Ind Co Ltd | Fuel injection valve for gas turbine |
| US5755092A (en) * | 1995-10-18 | 1998-05-26 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A) | Exhaust nozzle mixer for a turbofan engine |
| JP2003065156A (en) * | 2001-08-29 | 2003-03-05 | Ishikawajima Harima Heavy Ind Co Ltd | Lobe mixer |
| US20030145578A1 (en) * | 2002-02-06 | 2003-08-07 | Ishikawajima-Harima Heavy Industries Co., Ltd. | Lobe mixer for jet flow |
| JP2003314368A (en) * | 2002-04-25 | 2003-11-06 | Ishikawajima Harima Heavy Ind Co Ltd | Lobe mixer for gas turbine |
| JP2005147064A (en) * | 2003-11-18 | 2005-06-09 | Ishikawajima Harima Heavy Ind Co Ltd | Lobe mixer |
| CN1942665A (en) * | 2004-02-20 | 2007-04-04 | 杰克·H·安德森 | Jet nozzle mixer |
| CN101031713A (en) * | 2004-09-29 | 2007-09-05 | Snecma固体燃料推进器公司 | Mixer for separate-flow nozzle |
| CN102588145A (en) * | 2012-03-09 | 2012-07-18 | 董国光 | Air-breathing rotary detonation wave injection and multimode ram detonation coupled circulating propulsion system |
| CN104405530A (en) * | 2014-10-16 | 2015-03-11 | 南京航空航天大学 | Exhaust noise reduction system of separated exhaust type turbofan engine |
| CN205064048U (en) * | 2015-10-13 | 2016-03-02 | 中国航天空气动力技术研究院 | Tail gas mixing device |
| CN106246408A (en) * | 2016-08-03 | 2016-12-21 | 中国科学院工程热物理研究所 | A kind of ultra-compact high blending efficiency racemization lobe mixing exhaust system structure |
| CN109253470A (en) * | 2018-09-11 | 2019-01-22 | 中国人民解放军国防科技大学 | Supersonic mixing reinforcing device |
| CN110748436A (en) * | 2019-10-16 | 2020-02-04 | 南京航空航天大学 | Double-layer cooling mixing pipe injection type infrared suppressor |
| CN112902225A (en) * | 2021-02-04 | 2021-06-04 | 厦门大学 | Multistage afterburning chamber with outer ring rotary detonation supercharged combustion chamber |
| CN113028448A (en) * | 2021-03-15 | 2021-06-25 | 中国航发沈阳发动机研究所 | Non-uniform lobe mixer for turbo-fan engine afterburner |
| CN114183272A (en) * | 2021-10-25 | 2022-03-15 | 北京航空航天大学 | Injection mixer and control method thereof |
| CN114934856A (en) * | 2022-06-02 | 2022-08-23 | 江苏科技大学 | Marine ejector waste heat recovery system |
Non-Patent Citations (6)
| Title |
|---|
| ZHI-QIANG SHENG: "Jet mixing of lobed nozzles with spoilers located at lobe peaks", 《APPLIED THERMAL ENGINEERING》 * |
| 潘丞雄等: "斜切对抑制引射式波瓣喷管内部流动分离的效果研究", 《航空学报》 * |
| 谢翌等: "波瓣帽罩结构和脊线对混合排气系统性能影响", 《推进技术》 * |
| 邵万仁等: "波瓣形排气引射混合器的试验研究", 《航空动力学报》 * |
| 雷晗: "第三章 混合扩压器气动性能数值分析", 《中国优秀硕士论文电子期刊网》 * |
| 韩东等: "新型斜波瓣超声速混合器设计研究及其数值计算", 《航空动力学报》 * |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116379002A (en) * | 2023-06-05 | 2023-07-04 | 中国空气动力研究与发展中心空天技术研究所 | Design method of equal-rotation-speed reversing diffuser structure and diffuser structure |
| CN116379002B (en) * | 2023-06-05 | 2023-08-11 | 中国空气动力研究与发展中心空天技术研究所 | Design method of equal-rotation-speed reversing diffuser structure and diffuser structure |
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| CN115234938B (en) | 2022-11-25 |
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