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Небесная энциклопедия

Космические корабли и станции, автоматические КА и методы их проектирования, бортовые комплексы управления, системы и средства жизнеобеспечения, особенности технологии производства ракетно-космических систем

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Мониторинг СМИ

Мониторинг СМИ и социальных сетей. Сканирование интернета, новостных сайтов, специализированных контентных площадок на базе мессенджеров. Гибкие настройки фильтров и первоначальных источников.

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Поддерживает ввод нескольких поисковых фраз (по одной на строку). При поиске обеспечивает поддержку морфологии русского и английского языка
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Применить Всего найдено 76. Отображено 76.
23-06-2023 дата публикации

Oil supply system and combustion system integrated detection method and device

Номер: CN116296417A
Автор: SUN XIAOFENG, QIN LEI, CHENG RONGHUI, ZHANG GUANGYU, LIU BAO, WANG XIAOYU
Принадлежит:

The invention relates to an integrated detection method and device for an oil supply system and a combustion system, and the method comprises the steps: enabling fuel oil to pass through a main fuel oil pump and an equivalent valve in the oil supply system and then enter a combustion chamber of the combustion system; acquiring first fuel pressure after the fuel passes through the main fuel pump, first mass flow of the fuel flowing through the equivalent valve, pressure drop generated in the pipeline after the fuel passes through the equivalent valve, and first pipeline information of the pipeline; determining the pressure in the combustion chamber based on the first fuel pressure, the first mass flow, the pressure drop and the first pipeline information; and when the combustion chamber pressure is larger than a first preset threshold value, pressure abnormity of the combustion system is determined. The oil supply system and the combustion system are integrally considered, so that the coupling ...

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Номер записи: 1
07-07-2023 дата публикации

Afterburner aerodynamic characteristic simulation device for aero-engine complete turbine test

Номер: CN116399605A
Автор: LIU WEICHEN, CHENG RONGHUI, ZHANG XIAOYU, BAO ZHANYANG, NIU CHUNYAN, ZHOU CHUNYANG, GUO HONGTAO, LIU BAO, JIANG YU, XU XINGPING, WANG TAO
Принадлежит:

The invention belongs to the technical field of non-variable-capacity engine design and structural component testing, and particularly relates to an afterburner aerodynamic characteristic simulation device for aero-engine whole turbine testing. Only an outer casing, a confluence ring and an inner cone are reserved, and the afterburner structure is designed in a simplified mode; when an aero-engine complete turbine test is carried out, an actual afterburner can be replaced, it can be guaranteed that enough space is provided for installation and operation of turbine test equipment, wire leading is convenient, in addition, the confluence ring is modified and designed, the confluence ring is designed to be of a hollow structure with the rear edge open, and the confluence ring is convenient to install and operate. The annular plate with the flow control hole is detachably connected to the rear edge, so that outer culvert airflow can be introduced into the confluence ring and flows out from the ...

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Номер записи: 2
04-04-2023 дата публикации

Variable-aperture impingement cooling structure applied to suction surface of turbine guide vane

Номер: CN115898556A
Автор: LI HAIWANG, ZHU XIAOHUA, LI XIAOLIN, LIU SONG, YOU RUQUAN, XIE GANG, CHENG RONGHUI, GUO WEN
Принадлежит:

The invention discloses a variable aperture impingement cooling structure applied to a turbine guide vane suction surface, which comprises a plurality of impingement sheets, any two adjacent impingement sheets are connected through a separation rib, the top of each impingement sheet is provided with a plurality of rows of impingement holes, a turbulent flow part is mounted between any two adjacent rows of impingement holes, one end of the turbulent flow part abuts against the separation rib, and the other end of the turbulent flow part abuts against the separation rib. A plurality of impact cavities are formed in the impact thin plate, an impact target surface is arranged on the side, perpendicular to the separation ribs, of the impact thin plate, and a cold air outlet is formed in one end of the impact target surface and located in the impact cavities. The method has the advantages of being simple, convenient to machine and obvious in effect, and can be applied to various turbine blade ...

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Номер записи: 3
09-05-2023 дата публикации

Damper-rotor system response analysis method

Номер: CN116090135A
Автор: CHENG RONGHUI, SHI BIN, PANG YANLONG, ZENG YAO, ZHANG SHAOPING, LEI XINLIANG
Принадлежит:

The invention belongs to the technical field of aero-engines and gas turbines, and particularly relates to a damper-rotor system response analysis method. According to the method, the oil film force considering the oil film inertia force is introduced into the damper-rotor system, the unbalance response analysis method with the damper is established, and the analysis precision is higher.

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Номер записи: 4
29-08-2023 дата публикации

Compressor outlet total pressure correction method under core engine environment based on wake loss

Номер: CN116658451A
Автор: CHENG RONGHUI, ZHANG JINCHENG, ZHANG SHAOPING, ZHANG XUEBAO, YU HUAWEI, ZHANG JUN
Принадлежит:

The invention relates to the technical field of aero-engines, and discloses a method for correcting the total pressure of an outlet of a gas compressor in a core engine environment based on wake loss. The ratio of the average value of the second outlet total pressure at the same radial height to the average value of the first outlet total pressure at the corresponding radial height is taken as a first outlet total pressure correction coefficient by carrying out a core engine test; using the average value of the first outlet total pressure recovery coefficients under all radial heights obtained by all core engine tests as an actual outlet total pressure correction coefficient, and correcting the actually measured outlet total pressure of the gas compressor to obtain a corrected outlet total pressure value of the gas compressor; the accurate compressor pressure ratio characteristic considering wake loss under the core engine environment can be obtained, and the overall matching precision ...

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Номер записи: 5
15-08-2023 дата публикации

Cross-generation development small bypass ratio turbofan engine main engine and stress application matching method

Номер: CN116595680A
Автор: CHENG RONGHUI, CAO MAOGUO, HAOBISI GALATU, CHEN ZHONGGUANG, XU XINGPING, ZHANG ZHISHU, YUAN JILAI, JIANG YU, XUE HAIBO, CHEN DI, JIANG FANSHENG, CHEN ZEHUA, RUAN WENBO, ZHANG XUEDONG, ZHU ZHENKUN
Принадлежит:

The invention belongs to the field of aero-engine design, and relates to a cross-generation development small-bypass-ratio turbofan engine main engine and afterburner matching method, which comprises the following steps: preliminarily determining a main engine and afterburner design criterion, and then carrying out main engine and afterburner matching work preliminary scheme design; and processing hardware until the engine performance or the main engine and thrust augmentation matching design criterion requirements are met, then carrying out a full-size afterburner test, collecting test data, correcting the working abnormity of the afterburner, putting all parts of the full-size afterburner test into the whole engine, carrying out the whole-process test of the whole engine, and completing the whole-process test of the whole engine. The actual performance of all parts under the overall condition is obtained, after the work matching problem of a main engine and an afterburner is optimized ...

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Номер записи: 6
29-08-2023 дата публикации

Aero-engine invalid air leakage analysis method and suppression device

Номер: CN116663157A
Автор: CHENG RONGHUI, ZOU MI, ZHANG SHAOPING, HU YANLI, XU LIANQIANG, MA JIANDONG, SU CHANGMING, LIANG JINHUA
Принадлежит:

The invention relates to the technical field of aero-engines, and discloses an aero-engine invalid air leakage analysis method and a suppression device. Whether invalid air leakage exists or not is judged according to the deviation between an engine air system test result and a theoretical design result; then change sensitivity parameter simulation analysis is carried out on a labyrinth gap and a sealing gap in combination with an engine structure, the position where air leakage exists is determined, rapid recognition of the invalid air leakage position is achieved, a basis is provided for actual performance evaluation and improved design of an engine, and design robustness of an aero-engine air system is ensured; according to the analysis result, the cavity separating sleeve between the third cavity and the fourth cavity is improved, the original cavity separating sleeve is replaced with the cavity separating plate, the invalid air leakage restraining device is formed, the problem that ...

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Номер записи: 7
23-06-2023 дата публикации

Quick warm-up control method and device for ensuring reliable work of aero-engine

Номер: CN116291898A
Автор: CHENG RONGHUI, ZHANG XUEDONG, WANG JINGKE, CAO MAOGUO, JIA LINYUAN, BAI SHUAIYU, ZHAI YINGHAN, ZHU ZHENKUN, BING LIANXI, ZHANG ZHICHENG, LI CHENGLONG, BAI HANSONG, YUAN JILAI, JIANG FANSHENG, ZHANG ZHISHU, ZHANG XICHANG, CHEN ZHONGGUANG, WU YASHUAI
Принадлежит:

The invention belongs to the technical field of engine control, and particularly relates to a rapid warm-up control method and device for ensuring reliable work of an aero-engine. The method comprises the following steps: S1, acquiring a use mode of an engine, wherein the use mode comprises a combat mode or a training mode; s2, the starting state of the engine is obtained, wherein the starting state comprises a cold starting state or a hot starting state; and S3, four engine warming strategies are formed according to the combination of the use mode and the starting state of the engine, engine warming is carried out according to the different slow rotating speed staying time and the set rotating speed staying time at the set rotating speed higher than the slow rotating speed according to the engine warming strategies, and under the engine warming strategy with the combat mode and the cold-state starting state combined, engine warming is carried out according to the set rotating speed staying ...

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Номер записи: 8
13-06-2023 дата публикации

Tool for directly decomposing split sections of outer culvert casing in complete machine state of turbofan engine

Номер: CN116255211A
Автор: CHENG RONGHUI, ZHAO JIANGWEI, LI BO, CHEN XIAOFEI, WANG DANDAN, MA BO, WU ZHENGHONG, MA JIAN, LIN LEI
Принадлежит:

The invention relates to the technical field of aero-engines, and discloses a tool for directly decomposing a split section of an outer culvert casing in the complete machine state of a turbofan engine, the tool comprises a base, and a supporting seat and a supporting column are mounted on the base; the two supporting columns are arranged on the two sides of the whole ring section correspondingly. The two supporting columns are each provided with a force bearing pin capable of fixing an auxiliary installation section of the whole-ring-section reinforcing ring. A C-shaped frame is arranged between the two supporting columns; a bolt hole is formed in the C-shaped frame, and a force transmission pin capable of limiting the reinforcing ring is inserted into the bolt hole. The reinforcing ring is further limited and fixed by inserting the force transmission pin into the bolt hole, so that the whole ring section of the engine can be axially and horizontally restrained while the outer ring casing ...

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Номер записи: 9
04-08-2023 дата публикации

Thermal shock and thermal shock combined test device for engine hot end component

Номер: CN116539317A
Автор: TIAN WEI, FU YU, CHENG RONGHUI, HE JIN, ZHANG SHAOPING, ZHANG HANG, WANG QING, ZHONG YAN, LIU YANFEI
Принадлежит:

The invention provides a thermal shock and thermal shock combined test device for an engine hot end component, and belongs to the technical field of engine hot end component tests. The cooling liquid tank is mounted on the test table top, and an opening is formed in the top of the cooling liquid tank; the quartz lamp is mounted above the cooling liquid through a support frame; the test piece supporting mechanism is mounted between the quartz lamp and the bottom of the cooling liquid box through a lifting mechanism; the lifting mechanism is installed on the test table, the output end of the lifting mechanism is connected with the test piece bearing mechanism, and the lifting mechanism drives the test piece bearing mechanism to descend to enter the cooling liquid box or ascend to leave the cooling liquid box; the air cooling mechanism is mounted on the table top of the test table, and an air outlet of the air cooling mechanism faces the test piece on the test piece bearing mechanism; and ...

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Номер записи: 10
04-08-2023 дата публикации

Rotor connection structure deformation analysis method based on bolt assembly torque change

Номер: CN116542072A
Автор: HONG JIE, CHENG RONGHUI, MA YONGBO, MA YANHONG, CHEN XUEQI
Принадлежит:

The invention belongs to the technical field of aeroengine rotor dynamic characteristics and complete machine vibration, and particularly relates to a bolt assembly torque change-based rotor connection structure deformation analysis method. The method comprises the following steps of: extracting structure parameters of a rotor connection structure in an aeroengine; establishing a mechanical model of the flange-bolt connection structure; establishing a mathematical relationship between the bolt pre-tightening force in the bolt and the mechanical property; establishing a mathematical relationship between the assembly torque and the mechanical property; according to bolt loosening torque values, measured in engineering practice, of the connecting structure at different circumferential positions, a bolt assembly torque circumferential distribution diagram is drawn; and performing statistics and analysis on each numerical value in the assembly torque database, and establishing a statistical ...

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Номер записи: 11
04-04-2023 дата публикации

Hydraulic loading method and device suitable for aero-engine ground test

Номер: CN115901277A
Автор: LIU ZUOHONG, CHENG RONGHUI, LIN SHAN, GAO WEIMIN, XU GUANGLEI, CAI CHENGYANG, HE ZHIQIANG, GAO WEI, SUN ZHAOHUI, KANG CHEN
Принадлежит:

The invention belongs to the field of aero-engine design, and relates to a hydraulic loading method and device suitable for an aero-engine ground test, and the method comprises the steps: firstly, adding a large-flow proportional valve, a small-flow proportional valve, a throttling nozzle and a corresponding pipeline on the basis of rack loading equipment and a pipeline, and calibrating the hydraulic loading flow which can pass through the throttling nozzle; pipeline design is completed; then, the test debugging type is judged, short-term debugging can be completed through cooperation of the large-flow proportional valve and the small-flow proportional valve, and long-term debugging is completed by controlling the small-flow proportional valve and the throttling nozzles at the permutation and combination positions to be opened, finally, the small-flow proportional valve is controlled to conduct stepless regulation flow loading, and the throttling nozzles are controlled to conduct fixed ...

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Номер записи: 12
01-08-2023 дата публикации

Engine main combustion chamber smoke number analysis method based on axial partition

Номер: CN116522046A
Автор: CHENG RONGHUI, ZHAO YATING, SU JIAN, LI PENGYU, ZHOU SUNYU, WANG PEI, QIU JUNYUAN, XU BING, LIU XINSHENG, LI WU, ZHANG JIANGUO, LI QIANXIANG, TANG RUIYAN
Принадлежит:

The invention provides an engine main combustion chamber smoke number analysis method based on axial partition, and belongs to the field of aero-engines. The method includes the steps that working parameters of engine take-off working conditions serve as main combustion chamber smoke number calculation parameters; carrying out distribution calculation on air flow entering a flame tube of a main combustion chamber of the engine, wherein the air flow comprises a flame tube head, a main combustion area air flow and a flame tube total air quantity; according to the flame tube head air flow, the main combustion area air flow, the flame tube total air flow and the fuel oil flow entering the main combustion chamber, the flame tube head air-gas ratio, the main combustion area air-gas ratio and the main combustion chamber total air-gas ratio are calculated; and according to the inlet pressure, temperature and air flow of the main combustion chamber, based on an axial partitioning method, the gas-oil ...

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Номер записи: 13
23-06-2023 дата публикации

Novel integrated high-strength bearing cavity oil supply and heat insulation structure

Номер: CN116291884A
Автор: LENG ZIHAO, CHENG RONGHUI, CAO MAOGUO, SU ZHUANG, GUO SONG, ZHANG JIEYI, ZHENG KAI, LI GUOQUAN
Принадлежит:

The invention belongs to the technical field of aero-engines, and particularly relates to a novel integrated high-strength bearing cavity oil supply and heat insulation structure, through an integrated rear cover assembly structure, an oil supply pipe assembly and a skin heat insulation assembly structure are integrated, the assembly work of a rotor and a stator can be completed under the visual condition, and meanwhile through the design of a jackscrew hole structure, the assembly efficiency is improved. The disassembly work of the rear cover assembly can be very conveniently completed, structures such as a sealing device are prevented from being damaged, and the assembly and disassembly reliability of the bearing cavity oil supply and heat insulation structure is greatly improved; through the inner skin structure with the reinforcing ribs, the skin heat insulation assembly structure is optimized, under the condition of high pressure difference, the outer skin stress and the skin heat ...

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Номер записи: 14
18-04-2023 дата публикации

Method for designing gap between threaded main mounting joint mounting system and thrust pin

Номер: CN115982894A
Автор: ZENG YAO, CHENG RONGHUI, ZHANG SHAOPING, PANG YANLONG, SUN DENGKE, SU HANSHENG, ZHANG QIONG, LIU TIANWEN, ZHOU DAQING
Принадлежит:

The invention discloses a method for designing a gap between a mounting system with a threaded main mounting joint and a thrust pin, and belongs to the field of aero-engines. The method comprises the following steps: determining a highest temperature state and a first maximum horizontal overload load as a minimum gap design load by carrying out engine full working state load analysis, and obtaining a minimum gap value between a main mounting joint and a thrust pin by adopting a formula or a finite element analysis method; carrying out stress analysis on threaded connection of the main mounting section under a blade flying load condition, reversely deducing a thread tooth bearing limit by adopting an allowable stress value of a material in order to prevent a fracture load of the thread tooth, and determining that the superposed thread tooth bearing limit in a highest temperature state is a maximum gap design load; and a finite element analysis method is adopted to obtain the maximum gap ...

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Номер записи: 15
02-06-2023 дата публикации

Design method for supersonic speed and transonic speed blade profile of gas compressor rotor

Номер: CN116205008A
Автор: CHENG RONGHUI, LI ZHENG, YU HUAWEI, WANG QUANQI, ZHANG JUN, ZHANG SHAOPING
Принадлежит:

The invention provides a design method for a supersonic speed and transonic speed blade profile of a gas compressor rotor, and belongs to the technical field of gas compressor rotor blade design, and the design method comprises the steps that a gas compressor blade is sequentially divided into a first section, a second section and a third section according to the flowing direction of gas flow along a blade profile, the first section is a part in front of a gas flow blade channel, and the second section is a part in front of a gas flow blade channel; the second section is a section from airflow entering a blade channel to a shock wave generating section, and the third section is a shock wave rear flowing section; the first section is sequentially divided into a first increasing section and a second increasing section according to the flowing direction of airflow along the blade profile, and the ratio of the relative airflow bending angle to the relative chord length of the first increasing ...

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Номер записи: 16
30-06-2023 дата публикации

Excitation device of large-rigidity blade

Номер: CN116358879A
Автор: CHENG RONGHUI, GENG WEIMIN, WANG HUAITAO, FAN XINGCHAO, WANG MING
Принадлежит:

The invention provides an excitation device for a large-rigidity blade, and the excitation device is characterized in that a sealing cabin cover is arranged at the top of a test cabin body, and the test cabin body and the sealing cabin cover form a first accommodation chamber; the driving mandrel is arranged in the first containing cavity, one end of the driving mandrel penetrates through the sealing cabin cover and extends into the center position of the first containing cavity, and the other end of the driving mandrel is connected with the rotary driving device; the cover plate and the containing assembly are both arranged in the first containing cavity, the cover plate is arranged at the top of the containing assembly, the containing assembly and the cover plate form a cylindrical second containing cavity, the switching assembly is arranged in the second containing cavity, and the switching assembly is fixed to the driving mandrel; the rotating blade and the exciter are both arranged ...

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Номер записи: 17
01-08-2023 дата публикации

Optimization method for resisting chordwise bending vibration of stator blade of gas compressor with margin plate

Номер: CN116522545A
Автор: CHENG RONGHUI, GAO WEISI, ZHANG SHAOPING, PANG YANLONG, SUN HAIHE, ZHOU YUNLAI, ZHANG XUEBAO, ZHANG YIBIN, YIN LING, YANG FEIBING, LI SHENGXIANG
Принадлежит:

The invention provides a chordwise bending vibration resistance optimization method for a stator blade of a gas compressor with a margin plate. The chordwise bending vibration resistance optimization method comprises the following steps: acquiring a weak position in the middle of a blade body and a weak position at the root of the blade body of the stator blade; defining a dynamic intensity reserve ratio and an upper limit value and a lower limit value of a dynamic intensity reserve ratio threshold value; calculating the dynamic strength reserve ratio of the stator blade; comparing the dynamic strength reserve ratio with a dynamic strength reserve ratio threshold value, and optimizing the structural parameters of the stator blade according to the comparison result; and performing high-cycle fatigue test or simulation analysis on the optimized stator blade, and evaluating the chordwise bending vibration resistance of the optimized stator blade. According to the optimization method designed ...

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Номер записи: 18
23-06-2023 дата публикации

Method and system for judging anti-icing working state of aero-engine

Номер: CN116295592A
Автор: CHENG RONGHUI, LI LINGHAN, WU XIN, ZHU CHUANLONG, QI DONGXING
Принадлежит:

The invention provides a method and a system for judging an anti-icing working state of an aero-engine, and belongs to the technical field of aero-engines, the method comprises the following steps: constructing a relational expression between an anti-icing bleed air inlet temperature and an air compressor outlet temperature, and obtaining the anti-icing bleed air inlet temperature according to a measured value of the air compressor outlet temperature and the relational expression; a temperature sensor is arranged at an outlet of the anti-icing control device, a first-order inertia link transfer function of the temperature of an anti-icing air-entraining inlet and the temperature of the outlet of the anti-icing control device is constructed, and a temperature calculation value at the outlet of the anti-icing air-entraining device is obtained through calculation according to the temperature of the anti-icing air-entraining inlet and the first-order inertia transfer function; and the anti-icing ...

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Номер записи: 19
23-06-2023 дата публикации

Turbine rear thrust increasing component of aviation turbine engine

Номер: CN116291889A
Автор: CHENG RONGHUI, GUO HONGTAO, BAO ZHANYANG, XU XINGPING, LIU BAO, ZHANG XIAOYU, SONG WEIFENG
Принадлежит:

The invention belongs to the technical field of non-variable-capacity engine component design, and particularly relates to a turbine rear thrust increasing component of an aviation turbine engine, which is characterized in that an assembly composed of a heat shield, an oil inlet pipe, an oil nozzle and a flow adjusting sheet is designed to replace an existing oil injection rod, external culvert air is introduced to cool the component, the component is prevented from being damaged, and the service life of the component is prolonged. The flow of air introduced into the outer culvert can be controlled, the amount of air mixed with fuel oil is changed, the position of a flame front is adjusted conveniently and rapidly, combustion pulsation generated in a thrust increasing component can be effectively reduced, and the problems of insufficient combustion and flameout are solved.

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Номер записи: 20
15-08-2023 дата публикации

Accurate determination method for low-conduction capacity requirement of whole engine

Номер: CN116595790A
Автор: CHENG RONGHUI, HAOBISI GALATU, CHEN ZEHUA, YANG LONGLONG, CHEN ZHONGGUANG, ZHANG ZHISHU, CAO MAOGUO, HUANG YUJUAN, YU MING, RUAN WENBO, ZHANG XUEDONG, ZHU ZHENKUN, JIANG FANSHENG, XUE HAIBO, WU YASHUAI
Принадлежит:

The invention belongs to the field of aero-engine design, and relates to a method for accurately determining the low conduction capacity requirement of a complete engine, which comprises the following steps of: firstly, respectively carrying out part test, core engine and complete engine test on the engine by adopting the same set of hardware and the same set of test scheme to obtain the actual performance of each part under the overall condition; then, the high-pressure turbine expansion ratio and the external and internal culvert outlet pressure ratio are determined, the subsequent design direction is determined by judging the high-pressure turbine expansion ratio and the expansion ratio design value and the external and internal culvert outlet pressure ratio and the pressure ratio design value respectively, and then the influence of low-pressure turbine guider area adjustment on adjustable parameters is scored, so that the design direction is determined. According to the method, the ...

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Номер записи: 21
11-04-2023 дата публикации

Method for calculating disc center equivalent stress of aero-engine turbine disc

Номер: CN115952628A
Автор: CHENG RONGHUI, CHEN YANYAN, CHEN YUNFEI, PANG YANLONG, YAN YEHAO, ZENG YAO, SUN HAIHE, ZHANG SHAOPING
Принадлежит:

The invention provides a method for calculating equivalent stress of the center of a turbine disc of an aero-engine, belongs to the technical field of turbine discs of aero-engines, and particularly relates to a method for calculating equivalent stress of the center of the turbine disc of the aero-engine by analyzing transition state stress of the turbine disc under a given load spectrum to obtain stress history and temperature history of the center of the turbine disc, unifying the temperature of the stress history and then calculating equivalent stress of the center of the turbine disc of the aero-engine. The method comprises the following steps: calculating equivalent stress of a consideration symbol at each time point in a stress process, then carrying out rain flow counting to obtain accurate peak value time and valley value time, correspondingly subtracting unidirectional stress at the peak value time and the valley value time, and then obtaining an equivalent stress range according ...

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Номер записи: 22
06-06-2023 дата публикации

Single-crystal turbine blade shroud of aero-engine

Номер: CN116220833A
Автор: CHENG RONGHUI, CONG PEIHONG, JIN LI, GUO YONG, GUO JIE
Принадлежит:

The invention belongs to the technical field of non-variable-capacity engine blade design, and particularly relates to an aero-engine single-crystal turbine blade shroud, a wear-resistant layer is welded to the working face of the shroud, a base body exists between the wear-resistant layer and the working face of the shroud and between the wear-resistant layer and a non-working face transfer concave angle, and the length of the base body is larger than that of a welding fusion area. The blade crown working faces of the adjacent aero-engine single-crystal turbine blades are in contact through the wear-resisting layer-base body and the wear-resisting layer-wear-resisting layer, the mutual wear performance is guaranteed, meanwhile, a welding fusion area is far away from a transfer concave angle between the blade crown working face and a non-working face, the welding fusion area can be prevented from entering the transfer concave angle, and the welding quality of the blade crown working faces ...

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Номер записи: 23
30-05-2023 дата публикации

Anti-collision and anti-abrasion design method for blade shoulder of adjustable blade

Номер: CN116187110A
Автор: CHENG RONGHUI, FAN XIAOHONG, ZHANG QIANG, JIANG XUANKAI, ZHANG XUEBAO, MA WEITING, ZHANG SHAOPING, FU YU
Принадлежит:

The invention provides an anti-collision and anti-abrasion design method for a blade shoulder of an adjustable blade, and belongs to the technical field of design of adjustable blades of aero-engines, the anti-collision and anti-abrasion design method comprises the steps that an initial three-dimensional model of a cartridge receiver runner surface, an inner ring runner surface and a blade body is constructed, and an upper blade shoulder and a lower blade shoulder of the blade body intersect with the cartridge receiver runner surface and the inner ring runner surface respectively; the initial three-dimensional model is controlled to rotate around the rotation center line, and a plurality of basic three-dimensional models of the blade body at different rotation angles are obtained; the multiple basic three-dimensional models rotate around the rotating center line to the same position, the intersection of the multiple basic three-dimensional models is taken as a basic three-dimensional model ...

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Номер записи: 24
01-08-2023 дата публикации

Turbine blade engine state flow prediction method and system

Номер: CN116522826A
Автор: CHENG RONGHUI, ZHU XIAOHUA, HE JIN, GUO WEN, GU DONG, LU HONGKANG, ZHOU MENGZHI, SUN CHENG
Принадлежит:

The invention relates to the technical field of aero-engine turbine blade design, and discloses a turbine blade engine state flow prediction method and system.The turbine blade is divided into a plurality of area positions according to blade airflow flowing characteristics, and the relation between flow tests and engine state flow is established through all the areas; according to the method, a one-dimensional internal flow calculation model is obtained, so that the first actual flow of each partition is calculated, then the one-dimensional internal flow calculation model is corrected according to the first actual flow and the theoretical flow, the corrected one-dimensional internal flow prediction model of the engine state flow of each area is obtained, and the prediction model can improve the prediction precision of the turbine blade test state flow; the function and effect of the flow test are broadened, data support is provided for high-precision estimation of the turbine blade engine ...

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Номер записи: 25
07-04-2023 дата публикации

Aerial turbofan engine shaft component cyclic fatigue life load spectrum compilation method

Номер: CN115935523A
Автор: ZHENG MAOJUN, CHENG RONGHUI, ZHANG YONG, YANG CHANGQI, CONG PEIHONG, CAO MAOGUO, WANG HAIXU, GUO YONG
Принадлежит:

The invention provides an aviation turbofan engine shaft component cycle fatigue life load spectrum compilation method comprising the following steps: 1, collecting an engine typical mission profile and mixing frequency data thereof, and combining the height, the Mach number and the engine working state to form an engine typical working point; 2, corresponding overall performance parameters are collected according to typical working points of the engine and converted into torque characteristic values, and a graphic format of a typical task section is established with the torque characteristic values as ordinates and task section serial numbers as abscissas; 3, performing statistics on each typical mission profile according to the torque characteristic value by adopting a rain flow counting method to obtain all peak/valley value points of a plurality of cycle pairs and corresponding typical working points of the engine; and 4, selecting or clustering peak/valley value points from the typical ...

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Номер записи: 26
21-04-2023 дата публикации

Engine turbine rear thrust increasing component

Номер: CN115992776A
Автор: CHENG RONGHUI, LIU BAO, SUN XIAOFENG, CAO MAOGUO, XU XINGPING, YOU QINGJIANG, MA HONGYU, ZHU JIAN, LIU WEICHEN, JIANG YU, BAO ZHANYANG, CHEN DI, ZHANG ZHIXUE, LI JIANGNING, ZHOU CHUNYANG, GUO HONGTAO, ZHANG XIAOYU, MA YU, PAN XINZHENG, YAN JINSHENG
Принадлежит:

The invention belongs to the technical field of design of non-variable-capacity engine components, and particularly relates to an engine turbine rear thrust increasing component which comprises an annular outer wall, an inner wall and an outer wall. The confluence ring is arranged in the annular outer wall, and an outer culvert is formed between the confluence ring and the annular outer wall; the inner cone is arranged in the confluence ring, and an inner connotation is formed between the inner cone and the confluence ring; the annular heat insulation screen is arranged in the annular outer wall, is located at the outlet part of the annular outer wall and is corrugated, and a plurality of rows of air film holes are formed in each windward side of the corrugation; the outer side air entraining cylinders are arranged between the annular outer wall and the annular heat insulation screen, and the rear ends of the outer side air entraining cylinders are connected to crest positions of the corrugations ...

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Номер записи: 27
28-04-2023 дата публикации

Aero-engine air inlet distortion test method and device

Номер: CN116026604A
Автор: LIU ZUOHONG, CHENG RONGHUI, HE ZHIQIANG, CAI CHENGYANG, GAO LEI, KANG CHEN, YANG LONGLONG, GAO WEI, JIANG FANSHENG, BAI WEI
Принадлежит:

The invention belongs to the field of aero-engine tests, and relates to an aero-engine air inlet distortion test method and device, and the method comprises the steps: firstly assembling a plugboard type pressure distortion generator at an engine inlet, installing a data collection device for an engine air inlet measurement cross section, and carrying out the data collection of the engine air inlet cross section after driving; static debugging and slow-running inspection are carried out on the engine, the inserting plate type pressure distortion generator and the rack surge judging system, it is guaranteed that all devices on the engine are in a normal working state, after preparation work is completed, non-inserting plate performance recording of all states of the whole engine is carried out firstly, all parameters in an air inlet channel of the engine are recorded, and then the whole engine is subjected to static debugging and slow-running inspection; the data is stored in a data acquisition ...

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Номер записи: 28
27-06-2023 дата публикации

Aero-engine lubricating oil ventilation system with high sealing reliability

Номер: CN116335827A
Автор: LENG ZIHAO, CHENG RONGHUI, CAO MAOGUO, SU ZHUANG, HOU HONGJIAN, HU XINGLONG, GU ZHIYING, SONG GUANLIN, LI GUOQUAN
Принадлежит:

The invention belongs to the field of aero-engines, and particularly relates to a high-sealing-reliability aero-engine lubricating oil ventilation system which improves the oil return effect of a bearing cavity and avoids lubricating oil leakage by increasing the thickness of front and middle cavity oil return stage rotors in an oil return pump set and increasing the oil return capacity of front and middle cavities. Controllable ventilation electromagnetic valves are arranged on ventilation pipelines of a front bearing cavity and a middle bearing cavity, when an engine is at a low rotating speed, ventilation of the front cavity and the middle cavity is closed, and under the suction effect of an oil scavenging pump set, the cavity pressure of the front cavity and the middle cavity is reduced, the sealing pressure difference is increased, the sealing reliability is improved, and lubricating oil leakage of the front cavity and the middle cavity is avoided.

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Номер записи: 29
26-05-2023 дата публикации

Turbine guider and exhaust area measuring method thereof

Номер: CN116164695A
Автор: FU XINGHAO, CHENG RONGHUI, HUANG YUJUAN, MA ZHILE, LI JI, YANG KE, QU YUNFENG
Принадлежит:

The invention provides a turbine guider and an exhaust area measuring method thereof, and belongs to the technical field of aero-engines. When the measuring method is used for measuring the width of an exhaust window, the measuring planes of a left blade and a right blade are respectively perpendicular to respective stacking shafts, and the radial heights of mutually corresponding scanning curves of the left blade and the right blade are consistent; the window width is obtained based on the shortest distance between the blade profiles of the same radial height and is better matched with the actual width direction of the exhaust window. According to the method, during width measurement, the selectable radial height is not limited by the measurement method, a width measurement plane can be selected at will according to actual needs, the method has better applicability to bent and twisted blades and blades with complex exhaust window width distribution rules along the radial direction, and ...

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Номер записи: 30
14-07-2023 дата публикации

Aerial turbofan engine turbine rear thrust increasing device

Номер: CN116428074A
Автор: CHENG RONGHUI, ZHOU CHUNYANG, CHEN DI, BAO ZHANYANG, XU XINGPING, LIU BAO, JIANG YU
Принадлежит:

The invention belongs to the technical field of non-variable-capacity engine component design, and particularly relates to an aero-turbofan engine turbine rear thrust increasing device. According to the design, a radial stabilizer and a flame transfer groove are adopted for ignition, and combustion is organized; the turbine rear thrust increasing device mainly comprises an outer casing, a confluence ring, an inner cone, a radial stabilizer, an oil injection rod, an outer flame transfer groove section, a flame transfer groove section tail plate, an inner flame transfer groove section, a flame transfer groove section baffle plate and the like, and the ignition, flame transfer and organization combustion capacity of the turbine rear thrust increasing device can be effectively improved.

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Номер записи: 31
30-05-2023 дата публикации

Implementation method of blade modeling variable thickness distribution curve

Номер: CN116186945A
Автор: CHENG RONGHUI, WANG QUANQI, YU HUAWEI, ZHANG SHAOPING, ZHANG XUEBAO, LI ZHENG, ZHANG JUN, FU YU
Принадлежит:

The invention provides an implementation method of a blade modeling variable thickness distribution curve, and belongs to the technical field of fan or gas compressor design of gas turbine engines, and the implementation method specifically comprises the following steps: dividing a blade into a front edge section, a tail edge section and a main body section between the front edge section and the tail edge section; the thickness distribution of the front edge section is controlled by a circular curve to form a thickness distribution curve of the front edge section; the thickness distribution of the trailing edge section is controlled by a circular curve to form a thickness distribution curve of the trailing edge section; the position of the maximum thickness of the blade serves as a demarcation point, the main body section is divided into a front section and a rear section in the chord length direction of the blade, the thickness distribution of the front section is controlled through an ...

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Номер записи: 32
11-08-2023 дата публикации

Method for rapidly determining infrared stealth performance of aero-engine

Номер: CN116579135A
Автор: CHEN HANZE, CHENG RONGHUI, SHANG SHOUTANG, CAO MAOGUO, DENG HONGWEI, LU HAOHAO, WANG DIANLEI
Принадлежит:

The invention relates to a method for rapidly determining the infrared stealth performance of an aero-engine, and the method comprises the steps: carrying out the simulation calculation on an infrared stealth material on each backward visual part on the basis of setting that the infrared stealth material has a plurality of different damage areas; the infrared stealth performance of the aero-engine under the combination of different damage areas of the infrared stealth material on each backward visual part is obtained, so that a regression equation associated between the infrared stealth performance of the aero-engine and the damage area of the infrared stealth material on each backward visual part is trained, and a regression coefficient is obtained; selecting the aero-engine, measuring the damage area of each backward visual component, obtaining the infrared stealth performance of the aero-engine through simulation calculation, checking the accuracy of a regression equation, and quickly ...

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Номер записи: 33
04-08-2023 дата публикации

Method and system for determining full envelope vibration stress of engine turbine blade

Номер: CN116542182A
Автор: CHENG RONGHUI, WU JIAN, YANG YUANLONG, LI KUN, ZHANG SHAOPING
Принадлежит:

The invention relates to the technical field of aero-engine turbine blade design, discloses an engine turbine blade full envelope vibration stress determination method and system, and aims to accurately analyze and predict the engine full envelope turbine blade vibration stress based on the real test data with the limited engine working envelope range under the existing test conditions. The maximum vibration stress of the full envelope of the turbine blade is obtained, accurate analysis of the maximum vibration stress of the full envelope of the engine turbine blade is achieved, and the working reliability of the engine turbine blade can be remarkably improved.

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Номер записи: 34
16-05-2023 дата публикации

Duck-web-shaped turbine blade air film hole and turbine blade

Номер: CN116122916A
Автор: CHENG RONGHUI, LIANG JINHUA, CUI TINGTING, GUO WEN, SU YUNLIANG, ZOU MI, MA JIANDONG, XU LIANQIANG, ZHAO WEIWEI
Принадлежит:

The invention provides a duck-web-shaped turbine blade air film hole and a turbine blade, the air film hole comprises a cylindrical hole section, according to the cold air flowing direction, one end of the cylindrical hole section is an upstream cold air inlet end, and the other end of the cylindrical hole section is a downstream cold air outlet end; the expansion section is communicated with the downstream cold air outlet end of the cylindrical hole section; the expansion section is of a groove-shaped structure with an opening in the top face, a groove body of the groove-shaped structure comprises a protrusion extending outwards in the cold air flowing direction, a first branch and a second branch, the first branch and the second branch are located on the two sides of the protrusion respectively, and the first branch and the second branch are of symmetrical structures located on the two sides of the protrusion. The first branch, the protrusion and the second branch are in arc transition ...

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Номер записи: 35
09-05-2023 дата публикации

Passive control method and device for heat source mounting surface, inner wall surface and outer wall surface of combustion chamber

Номер: CN116090107A
Автор: SUN XIAOFENG, QIN LEI, CHENG RONGHUI, ZHANG GUANGYU, LIU BAO, WANG XIAOYU
Принадлежит:

The invention provides a passive control method and device for a heat source mounting surface and inner and outer wall surfaces of a combustion chamber, and belongs to the field of aviation. A scattering sound field function formed by an equivalent monopole source describing heat source installation surface impedance in the diffusion cavity, formed by an equivalent monopole source describing combustion chamber outer wall surface acoustic liner impedance in the acoustic liner back cavity and formed by an equivalent monopole source describing combustion chamber inner wall surface acoustic liner impedance in the acoustic liner back cavity is constructed; obtaining a characteristic equation through Laplace transformation based on the pressure disturbance function and each scattering sound field function; according to the characteristic solution of the characteristic equation, the inhibition effect of the pore parameter combination on combustion instability is judged so as to determine a pore ...

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Номер записи: 36
29-08-2023 дата публикации

Method and device for determining time constant of temperature sensor of turbofan engine

Номер: CN116659870A
Автор: ZHANG XUEDONG, CHENG RONGHUI, XIA YU, TANG ZHENGFU, JIA LINYUAN, BAI SHUAIYU, YUAN JILAI, JIANG FANSHENG, ZHU ZHENKUN, BING LIANXI, SUN HAOBO, ZHANG ZHISHU, ZHANG XICHANG, CHEN ZHONGGUANG, GALATU, HOBES, RUAN WENBO, ZHANG ZHICHENG
Принадлежит:

The invention belongs to the technical field of engine control, and particularly relates to a turbofan engine temperature sensor time constant determination method and device. The method comprises the following steps: S1, obtaining a reference time constant given based on a wind tunnel test, wherein the reference time constant is a time constant obtained by a low-pressure turbine outlet temperature sensor in a reference environment of a wind tunnel calibration laboratory; s2, the conversion rotating speed of an inlet of the high-pressure compressor is obtained; and S3, based on a given constant K and a given sensor grade index m, determining a calibrated time constant through a calibration formula. According to the method, the time constant of the low-pressure turbine outlet temperature sensor under any rotating speed state in the full envelope transient process of the turbofan engine can be obtained, and the problem that the time constant of the low-pressure turbine outlet temperature ...

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Номер записи: 37
18-07-2023 дата публикации

Bearing cavity oil return and ventilation structure with low lubricating oil consumption

Номер: CN116447013A
Автор: LENG ZIHAO, CHENG RONGHUI, CAO MAOGUO, SU ZHUANG, SONG GUANLIN, HU XINGLONG, HAN JINZAI, GUO SONG, ZHANG JIEYI, YU LI, CHEN XUEQI, MAO HONGTU
Принадлежит:

The invention belongs to the field of aero-engines, and particularly relates to a bearing cavity oil return and ventilation structure with low lubricating oil consumption, by designing a bearing cavity oil return oil-gas separation structure, a high-speed rotating gas cavity and a lubricating oil return cavity are isolated, oil and gas in an oil collection cavity are effectively separated, interference of high-speed flowing gas on flowing of lubricating oil in the oil return cavity is reduced, and the lubricating oil in the lubricating oil return cavity is effectively separated. By designing a ventilation oil-gas separation structure of the bearing cavity, oil gas near the ventilation hole is effectively separated, the amount of the lubricating oil entering the ventilation cavity is reduced, the oil return efficiency is improved, and the lubricating oil consumption is reduced.

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Номер записи: 38
28-07-2023 дата публикации

Main fuel control method and device in acceleration process of aero-engine

Номер: CN116498448A
Автор: ZHANG XUEDONG, CHENG RONGHUI, TANG ZHENGFU, XIA YU, JIA LINYUAN, BAI SHUAIYU, SHI LEI, ZHAI YINGHAN, YUAN JILAI, JIANG FANSHENG, ZHU ZHENKUN, BING LIANXI, CHEN ZHONGGUANG, ZHANG ZHISHU, ZHANG XICHANG, HAOBISI GALATU, RUAN WENBO, ZHANG ZHICHENG
Принадлежит:

The invention belongs to the technical field of engine control, and particularly relates to a main fuel control method and device in the acceleration process of an aero-engine. The method comprises the steps that S1, a main fuel oil supply flow conversion value is calculated based on the conversion relation between the main fuel oil supply flow conversion value and the high-pressure compressor inlet conversion rotating speed; s2, the fan inlet total temperature of the engine at the current moment is obtained, and a first correction value is calculated based on a set first correction function; s3, the fan inlet total pressure of the engine at the current moment is obtained, and a second correction value is calculated based on a set second correction function in combination with the high-pressure compressor inlet conversion rotating speed; and S4, the main fuel oil supply flow conversion value is corrected according to the first correction value and the second correction value, and the main ...

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Номер записи: 39
28-07-2023 дата публикации

Turbine blade temperature correction method and device based on temperature test data

Номер: CN116502469A
Автор: CHENG RONGHUI, KOFUYU, GUO WEN, KANG YAJIE, ZHU XIAOHUA, WANG CHAO, SHI JIAWEI, ZHOU YINGLU, WANG SIYI
Принадлежит:

The invention relates to the technical field of aero-engine turbine blade design, and discloses a turbine blade temperature correction method and device based on temperature test data, and the method comprises the steps: carrying out the partitioning of a turbine blade, and obtaining a corrected simulation node temperature value in each partition based on the temperature measurement data of a temperature measurement point; the corrected simulation node temperature value is closer to an actual temperature field in a use environment, the turbine blade temperature simulation precision and reliability are improved, it is ensured that blade strength evaluation can be carried out more accurately, important support is provided for turbine blade design and even development of a whole engine, and the method has good engineering practical value; direct high-fidelity application of limited test data in turbine blade temperature simulation is achieved, the economic cost is low, the utilization efficiency ...

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Номер записи: 40
21-07-2023 дата публикации

Sawtooth crown blade and blade crown tightness design method thereof

Номер: CN116467804A
Автор: CHENG RONGHUI, GUO RUI, CAO MAOGUO, MA SHIYAN, WEI WENTAO, YANG YANGHUA, YU RONGBIN
Принадлежит:

The invention provides a sawtooth crown blade shroud tightness design method which comprises the following steps: constructing a relation between blade shroud torsional mechanics and a geometric position according to a blade shroud stress condition of a turbine blade in an assembly state; introducing a tightness change rate, and constructing a relational expression between the shroud tightness function and the shroud meshing angle, the tightness change rate and the shroud pre-torsion angle; drawing a plurality of curves with equal pre-torsion angles and equal meshing angles by taking the tightness change rate as an abscissa and the shroud tightness function as an ordinate to obtain a shroud tightness extrusion force function curve; determining a shroud tightness function, a tightness change rate and a shroud meshing angle range, drawing a maximum straight line and a minimum straight line of the shroud tightness function on a shroud tightness function curve, and drawing a maximum curve and ...

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Номер записи: 41
04-04-2023 дата публикации

Air film hole structure of turbine blade

Номер: CN115898554A
Автор: CHENG RONGHUI, LIANG JINHUA, CUI TINGTING, GUO WEN, SU YUNLIANG, ZOU MI, MA JIANDONG, XU LIANQIANG, ZHAO WEIWEI
Принадлежит:

The invention discloses an air film hole structure of a turbine blade, belongs to the technical field of turbine blades, and solves the technical problems that in the prior art, the air film coverage rate of an air film hole in a blade is low, and the cooling effect of the blade is reduced. The air film hole structure comprises a cylindrical section, a protruding section, a contraction section, an expansion section and an outlet face which are sequentially arranged in an integrated mode in the airflow direction so as to be suitable for external cooling of the turbine blade, and the cold air covering effect of the turbine blade is improved.

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Номер записи: 42
04-08-2023 дата публикации

Compressor blade vibration reduction optimization design evaluation method based on energy method

Номер: CN116541970A
Автор: CHENG RONGHUI, GAO WEISI, ZHANG SHAOPING, PANG YANLONG, SUN HAIHE, ZHOU YUNLAI, ZHANG XUEBAO, LI ZIWEI, YIN LING, XU YONGQIANG, YANG FEIBING, LI SHENGXIANG
Принадлежит:

The invention provides a compressor blade vibration reduction optimization design evaluation method based on an energy method. The compressor blade vibration reduction optimization design evaluation method comprises the steps that a dangerous area of a blade before optimization is obtained; calculating the total kinetic energy ratio of the total kinetic energy of the blades before optimization to the total kinetic energy of the blades after optimization; based on modal analysis, calculating a first strain energy ratio of the dangerous area of the blade before optimization and a second strain energy ratio of the dangerous area of the blade after optimization; according to the total kinetic energy ratio, the first strain energy ratio and the second strain energy ratio, the dangerous area strain energy ratio of the blade before optimization to the blade after optimization is calculated; based on the dangerous area strain energy ratio, obtaining a vibration stress ratio of dangerous areas of ...

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Номер записи: 43
01-08-2023 дата публикации

Turbine blade air supply system pressure robustness analysis method

Номер: CN116522693A
Автор: CHENG RONGHUI, HU YANLI, ZHANG SHAOPING, GUO WEN, XU LIANQIANG, MA JIANDONG, ZOU MI, SU CHANGMING
Принадлежит:

The invention provides a turbine blade air supply system pressure robustness analysis method, and relates to the technical field of aero-engines. Comprising the following steps: determining the maximum opening gap delta of the rotor rear baffle sealing structure; defining the ratio of the prewhirl nozzle outlet pressure PQA of the turbine blade gas supply system to the total outlet pressure P3 of a gas compressor as a gas supply pressure characteristic parameter PQA/P3; the engine load serves as the boundary condition, and the minimum value X of PQA/P3 needed by the turbine blade in the maximum take-off state is obtained; according to the maximum gap delta, the PQA/P3 change range Y before and after the sealing structure fails is determined; the current PQA/P3 is compared with X and Y, and if PQA/P3 is smaller than X or (PQA/P3) n-(PQA/P3) 1 is larger than Y, it is indicated that non-design air leakage exists in the rotor sealing structure. Whether the rotor sealing structure works stably ...

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Номер записи: 44
23-06-2023 дата публикации

Turbine rear casing structure typical bending part stress test piece

Номер: CN116296302A
Автор: XIE LIJUAN, CHENG RONGHUI, CONG PEIHONG, BAI HANSONG, LIU MEI, SONG YANG, LYU TIANBO
Принадлежит:

The invention belongs to the technical field of structural design of typical structure simulators of turbine rear casings, and particularly relates to a stress test piece for typical bending parts of turbine rear casing structures, which comprises a plate-shaped test piece, two ends of the plate-shaped test piece are test assessment areas, and the width of the two test assessment areas is the wall thickness L1 of the typical bending parts of the turbine rear casing structures; a plurality of runway-shaped holes arranged in the width direction of the plate-shaped test piece are formed in the plate-shaped test piece in a range L2 away from the two test assessment areas; the size of L2 and the distribution position of each runway-shaped hole are limited, so that the stress on one side is the maximum and the stress on the other side tends to be zero when the two test assessment areas are loaded.

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Номер записи: 45
15-08-2023 дата публикации

High-temperature area threaded connection method and assembly with anti-bonding, anti-rotation and retaining functions

Номер: CN116592031A
Автор: CHENG RONGHUI, LIU CHAO, YU ZUCHAO, ZHU CHENG, SHAO JIANBO, YING ZHUOJUN, ZHAO DAN, XU LIANGLIANG, WU ZHENGHONG, LIN LEI
Принадлежит:

The invention relates to the technical field of aero-engines, and discloses a high-temperature area threaded connection method and assembly with anti-bonding, anti-rotation and retaining functions, a conical uniform-load hole is formed in the tail of a D-head bolt in the axial direction, and the conical uniform-load hole extends to the middle of the bolt from the tail of the D-head bolt, so that the variable stiffness design of a bolt rod is met, and the service life of the bolt rod is prolonged; the thread teeth can be ensured to be stressed more uniformly during working; especially, the problems that when the internal thread and the external thread are matched, the stress conditions of the thread teeth are sequentially decreased in the direction from the head to the tail of the bolt, the first several thread teeth bear about 70% of loads, the extrusion loads of the thread teeth are large, consequently, the thread teeth are deformed and clamped, friction between the thread teeth is too ...

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Номер записи: 46
04-08-2023 дата публикации

Compressor bolt connection part temperature field calculation method based on thermal resistance

Номер: CN116541969A
Автор: CHENG RONGHUI, LI XIAOLIN, ZHANG SHAOPING, REN FANG, HE JIN, GUO WEN
Принадлежит:

The invention provides a gas compressor bolt connection part temperature field calculation method based on thermal resistance. The gas compressor bolt connection part temperature field calculation method comprises the steps of obtaining a first high-temperature side reference temperature and a first low-temperature side reference temperature in a three-dimensional temperature field; constructing a two-dimensional simulation model of the bolt connection part; acquiring the application thermal resistance of the bolt connection part; averagely distributing the applied thermal resistance and calculating a two-dimensional temperature field initial value of the bolt connection part; calculating a second high-temperature side reference temperature and a second low-temperature side reference temperature in the two-dimensional temperature field; correcting the applied thermal resistance; and adjusting an application thermal resistance distribution proportion, and calculating a two-dimensional temperature ...

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Номер записи: 47
25-04-2023 дата публикации

Device and method for assembling turbine disc of aero-engine and front sealing disc of turbine disc

Номер: CN116000604A
Автор: ZHAO DANDAN, ZHANG JINMING, CHENG RONGHUI, LIU JIANXIAN, LI QINGHUI, GUO TIEYING
Принадлежит:

The invention provides an aero-engine turbine disc and front sealing disc assembling device and method thereof.The aero-engine turbine disc and front sealing disc assembling device comprises a pull disc, a front sealing disc and a rear sealing disc, the center of the pull disc is provided with a mounting hole, and the edge of the pull disc is provided with countersunk pull holes and large receding holes which are distributed in the circumferential direction; each countersunk pull hole is correspondingly communicated with one large abdicating hole; the hook end of each drag hook is correspondingly hooked into one cooling hole of the front sealing disc, and the root of each drag hook correspondingly penetrates through one large receding hole and is clamped into the corresponding countersunk pull hole; the hydraulic cylinder is screwed in the mounting hole; the pressure plate is pressed at the end part of the front shaft neck of the sealing plate, is positioned with the end part of the front ...

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Номер записи: 48
04-04-2023 дата публикации

Fan-shaped on-duty stabilizer and oil supply method thereof

Номер: CN115899762A
Автор: CHENG RONGHUI, JIANG YU, CAO MAOGUO, XU XINGPING, LIU BAO, CHEN DI, BAO ZHANYANG, YOU QINGJIANG, ZHANG XIAOYU, MA YU, KANG BOFENG
Принадлежит:

The invention belongs to the field of aero-engine design, and particularly relates to a fan-shaped on-duty stabilizer and an oil supply method thereof. The radial stabilizers are installed on the inner wall of the afterburner casing in the circumferential direction, the roots of the radial stabilizers are installed on the inner wall face of the afterburner casing, and the tips of the radial stabilizers point to the center line of the afterburner casing; wherein each radial stabilizer is provided with a cavity, each cavity is provided with a main spray rod and an auxiliary spray rod, and the main spray rods and the auxiliary spray rods spray oil through nozzles on the side walls of the radial stabilizers; the main spraying rod is connected with a main pipe for supplying oil; the auxiliary spraying rod comprises a first group of auxiliary spraying rods and a second auxiliary spraying rod, the first group of auxiliary spraying rods are connected with a first auxiliary pipe for supplying oil ...

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Номер записи: 49
29-08-2023 дата публикации

Method and device for screening blades of blisk of gas compressor with controllable frequency dispersibility

Номер: CN116663200A
Автор: CHENG RONGHUI, YANG YUANLONG, PANG YANLONG, WU JIAN, LI KUN, ZHANG SHAOPING, SUN HAIHE, TIAN HONGYU, HE JIN
Принадлежит:

The invention relates to the technical field of aero-engines, and discloses a method and device for screening blades of a blisk of a gas compressor with controllable frequency dispersity. The ratio of the second-order bending frequency to the first-order bending frequency of the blades corresponding to a lower difference model of the blades of the blisk of the gas compressor is used as the minimum value of blade profile parameters; the maximum value of the blade profile parameter is obtained according to the minimum value of the blade profile parameter and the frequency dispersion control deviation analysis of the blade of the integral blade disc of the gas compressor; and then testing to obtain the actual first-order bending frequency and the actual second-order bending frequency of each blade to be tested, and selecting the blade of which the actual blade profile parameter of the corresponding blade to be tested is within the range of the minimum blade profile parameter value and the ...

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Номер записи: 50
06-06-2023 дата публикации

Turbine disc shaft connecting structure for improving robustness of aero-engine

Номер: CN116220822A
Автор: CHENG RONGHUI, GENG XU, WANG DONG, LYU CHUNGUANG, LIU YU, LIU WENDONG, JIA PENGZHI
Принадлежит:

The invention belongs to the field of aero-engine design, and relates to a turbine disc and shaft connecting structure for improving the robustness of an aero-engine, the turbine disc and shaft connecting structure comprises a turbine disc and a turbine shaft, an oblique cone wall extends out of the turbine disc, vent holes are formed in the oblique cone wall, and the oblique cone wall comprises a straight part and a vertical part; the straight part is arranged in the axial direction of the aero-engine, and a lengthened section is arranged on the straight part. The oblique cone wall forms a bent structure through the vertical part and the straight part, and the oblique cone wall of the turbine disc is lengthened, so that the problem that the strength of the cone wall is weakened due to vent holes is solved under the condition that the vent holes of the cone wall of the turbine disc ensure the same ventilation area; by arranging the lengthening section, the mass center distance between the ...

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Номер записи: 51
04-08-2023 дата публикации

Method for recording steady-state performance of compression part of aero-engine

Номер: CN116539311A
Автор: CHENG RONGHUI, WU SENLIN, ZHAO ZHENG, LIANG JUN, WAN QIANJUN, FENG FEI
Принадлежит:

The invention provides a steady-state performance recording method for an aero-engine compression part. The steady-state performance recording method comprises the steps that firstly, measured parameters are configured; step 2, calculating the average value of the total outlet temperature of the compression part collected in each second; step 3, performing primary linear fitting on the outlet total temperature in the first continuous acquisition time every second to obtain the slope of a fitting straight line; 4, the slope value in the second continuous acquisition time period is judged every second, and whether the test state is stable or not is determined according to the judgment result; step 5, when the test state is stable, sending an acquisition prompt through acquisition software; and 6, recording the equal rotating speed characteristics of the compression part and evaluating the performance of the compression part. According to the invention, by evaluating the measurement range ...

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Номер записи: 52
06-06-2023 дата публикации

Method for constructing maneuvering overload spectrum of aero-engine

Номер: CN116227148A
Автор: ZHENG MAOJUN, ZHANG YONG, CHENG RONGHUI, CONG PEIHONG, CAO MAOGUO, BAI HANSONG
Принадлежит:

The invention provides an aero-engine maneuvering overload spectrum construction method, which is used for constructing a maneuvering load spectrum of a new trans-generation research model, and takes demand analysis as a starting point, based on analysis of a design usage of the new trans-generation research model, various existing limitation data conditions are collected, analyzed, absorbed and integrated, and the maneuvering overload spectrum of the aero-engine is constructed. According to the method, the problem that the maneuvering overload spectrum of the aero-engine cannot be directly obtained through out-field use data statistics due to the fact that the maneuvering overload spectrum of the aero-engine has no direct design input and the difference between active service model usage is large is solved, and a basis is provided for life design and test of an important force bearing structure of the engine; due to the fact that the method is closely attached to a cross-generation model ...

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Номер записи: 53
01-08-2023 дата публикации

Axial symmetry type drag reduction cover and molded line parametric modeling method thereof

Номер: CN116522491A
Автор: CHENG RONGHUI, FU XINGHAO, MA ZHILE, HUANG YUJUAN
Принадлежит:

The method comprises the steps that the front edge/the tail edge of the molded line of the resistance reduction cover is configured to be in an arc shape, and the connecting line of the circle centers of the arc shapes of the front edge/the tail edge is a symmetry axis; determining an end point of a symmetric axis according to the axial limiting length of the drag reduction cover, symmetrically offsetting the symmetric axis towards two sides by a preset distance to construct a maximum width line of the drag reduction cover, and determining a profile control point of the drag reduction cover by combining the maximum width axial relative position; giving the radius of a front/trailing edge small circle, and determining the center position of the front/trailing edge small circle; the front/tail edge small circle of the drag reduction cover and the symmetry axis respectively have two intersection points, a front/tail edge molded line tangent point position coefficient is constructed, the arc-shaped ...

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Номер записи: 54
28-07-2023 дата публикации

Turbine blade margin plate heat exchange boundary construction method

Номер: CN116502369A
Автор: CHENG RONGHUI, KOFUYU, ZHANG SHAOPING, ZHU XIAOHUA, WANG SIYI, KANG YAJIE, WANG CHAO, SHI JIAWEI, ZHOU YINGLU, LI KUN
Принадлежит:

The invention provides a turbine blade margin plate heat exchange boundary construction method, and relates to the field of aero-engines. Comprising the following steps: S1, acquiring geometrical characteristic parameters of a single-blade margin plate surface, and constructing an outer ring side line and an inner ring side line; s2, obtaining a front edge endpoint coordinate and a rear edge endpoint coordinate of an inner ring side line, and dividing the inner ring side line into a blade basin curve and a blade back curve; obtaining the coordinates of a front edge line projection point and a rear edge line projection point through the projection of the front edge end point and the rear edge end point on the front edge line and the rear edge line; s3, physical parameters for constructing a single-blade marginal plate surface heat exchange boundary are obtained; s4, constructing a runner edge plate surface between two adjacent blades, and dividing the runner edge plate surface into orthogonal ...

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Номер записи: 55
04-08-2023 дата публикации

Inter-stage matching characterization method of multi-stage axial flow gas compressor

Номер: CN116542077A
Автор: CHENG RONGHUI, ZHANG JUN, ZHANG SHAOPING, ZHANG JINCHENG, ZHANG XUEBAO, YU HUAWEI
Принадлежит:

The invention provides an inter-stage matching characterization method for a multistage axial-flow gas compressor. The inter-stage matching characterization method comprises the steps that a full-rotation-speed characteristic curve of each stage of the multistage axial-flow gas compressor is obtained; calculating a stage adiabatic efficiency characteristic curve and a stage pressure coefficient curve according to the full rotation speed characteristic curve of each stage and the inlet flow field parameters; extracting a first flow coefficient corresponding to the highest adiabatic efficiency point and a first pressure coefficient corresponding to the first flow coefficient; based on the common working line and the full-rotating-speed characteristic curve of each stage, extracting a working point of each stage at each rotating speed; extracting a second flow coefficient and a second pressure coefficient according to each stage of inlet flow field parameter of the working point; and calculating ...

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Номер записи: 56
08-08-2023 дата публикации

Aircraft engine compressor rotor spigot tightness design method

Номер: CN116562106A
Автор: CHENG RONGHUI, ZHANG SHENZHI, PANG YANLONG, FAN XINGCHAO, ZHANG SHAOPING, GAO WEISI, DU WENJUN
Принадлежит:

The invention provides an aero-engine compressor rotor spigot tightness design method, and relates to the technical field of aero-engines. Comprising the steps that the engine working state with the minimum spigot extrusion stress is obtained; under the state, the extrusion stress on the gas compressor rotor spigot contact surface is obtained; through iterative calculation, the corresponding seam allowance interference magnitude when the extrusion stress is larger than 0 is obtained, and the seam allowance interference magnitude is the minimum assembly tightness; the engine working state with the maximum spigot extrusion stress is obtained; under the state, obtaining extrusion stress on a gas compressor rotor spigot contact surface and shear stress of a spigot root section, calculating to obtain yield strength reserve and shear yield strength reserve, and obtaining corresponding spigot interference magnitudes when the yield strength reserve is greater than or equal to 1 and the shear yield ...

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Номер записи: 57
04-07-2023 дата публикации

Thermal protection structure for fuel nozzle of aero-engine

Номер: CN116379474A
Автор: ZHAO XIAN, SU JIAN, CHENG RONGHUI, LI PENGYU, XU BING, WANG QING, LI QIANXIANG, QIU JUNYUAN, ZHANG JIANGUO
Принадлежит:

The invention provides a thermal protection structure for a fuel nozzle of an aero-engine, which belongs to the technical field of aero-engines and specifically comprises a heat shield and a part wrapping the nozzle and positioned in a casing, and the heat shield is provided with an opening corresponding to the fuel outlet end face of the nozzle; the welding boss is arranged on the outer wall of the end, close to the mounting flange, of the rod part, and the end, close to the mounting flange, of the heat shield is fixedly connected with the welding boss; the limiting bosses are arranged on the periphery of the end, close to the head, of the rod part, and the multiple limiting bosses are distributed in the circumferential direction of the rod part; a plurality of weakening holes are formed in the side wall of the end, close to the mounting flange, of the heat shield, sealing rings are arranged on the sides, away from the welding bosses, of the weakening holes of the rod portion of the nozzle ...

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Номер записи: 58
04-07-2023 дата публикации

Method for inspecting and controlling production quality of Ti60 titanium alloy blisk

Номер: CN116380892A
Автор: CHENG RONGHUI, TIAN WEI, ZHANG SHAOPING, FU YU, GUO HUIMING, ZHONG YAN, HE JIN, BAI YUNRUI, ZHANG XUEBAO, ZHANG HANG
Принадлежит:

The invention relates to a method for inspecting and controlling the production quality of a Ti60 titanium alloy blisk, which belongs to the technical field of manufacturing of key materials and parts of aero-engines, and comprises the following steps of: performing die forging process treatment, solid solution heat treatment and aging heat treatment on a plurality of batches of blanks, and forming a Ti60 titanium alloy blisk forge piece; one forge piece is extracted from each batch of blade disc forge pieces and is dissected in the radial direction and the axial direction of the forge piece; microscopic structure samples with different sizes are cut at least from the blade, the rim, the radial plate and the hub; each microscopic structure sample is pretreated to prepare a to-be-detected surface; and optical microscopes with different multiples are adopted to carry out microscopic structure testing on each inspection surface, whether the parameter characteristics of the beta crystal grains ...

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Номер записи: 59
14-04-2023 дата публикации

Afterburner for aero-engine

Номер: CN115962484A
Автор: CHENG RONGHUI, CHEN DI, ZHOU CHUNYANG, XU XINGPING, LIU BAO, BAO ZHANYANG, YOU QINGJIANG, JIANG YU, LIU WEICHEN
Принадлежит:

The invention belongs to the technical field of afterburner design for general engines, and particularly relates to an afterburner for an aero-engine. Mainly comprises an outer casing, a confluence ring, an inner cone, an annular stabilizer outer wall, an annular stabilizer inner wall, an annular stabilizer front wall, an annular stabilizer rear wall, an annular stabilizer inner ring plate, an annular stabilizer inner ring pipe, an annular stabilizer oil injection rod, a radial stabilizer, a radial stabilizer inner ring pipe and a radial stabilizer oil injection rod. And the ignition, cross flame and stable combustion capability in the combustion chamber can be ensured.

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Номер записи: 60
04-08-2023 дата публикации

Compressor rotor temperature stress deformation analysis method and device based on overall parameters

Номер: CN116542114A
Автор: CHENG RONGHUI, YANG YUANLONG, PANG YANLONG, WU JIAN, TIAN HONGYU, ZHANG SHAOPING, LI KUN, HE JIN
Принадлежит:

The invention relates to the technical field of aero-engine turbine blade design, and discloses a gas compressor rotor temperature stress deformation analysis method and device based on overall parameters. A function model between the temperature of the gas compressor and the overall parameter, a function model between the equivalent stress of the gas compressor and the overall parameter and a function model between the total displacement of the gas compressor and the overall parameter are established by using typical state points in the flight envelope of the engine, and are used for calculating a temperature value, an equivalent stress value and a deformation value of a to-be-measured state in a segmented manner. The fitting model not only can completely cover the working state points of the engine, but also avoids the problem of data extrapolation, ensures the calculation precision of each function model, can obtain thousands of times of calculation efficiency improvement, and achieves ...

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Номер записи: 61
30-06-2023 дата публикации

Composite damage analysis method under interaction influence of wear damage and contact fatigue damage

Номер: CN116362058A
Автор: MA YANHONG, CHENG RONGHUI, HAN XUDONG, HONG JIE, CHEN XUEQI
Принадлежит:

The invention relates to the technical field of mechanical engineering, and particularly provides a composite damage analysis method under the interaction influence of wear damage and contact fatigue damage, which comprises the following steps: firstly, establishing a mechanical model, and then solving the wear damage rate when the wear damage independently occurs and the fatigue damage rate when the contact fatigue damage independently occurs; and then based on the interactive influence relationship between the contact fatigue damage and the wear damage, correcting the wear damage rate and the contact fatigue damage rate, then establishing an interface composite damage tolerance criterion, and finally obtaining the interface composite damage failure life. According to the method, the problem that an interface life evaluation method based on linear superposition of two kinds of damage is not accurate enough is solved, the interaction influence effect of abrasion damage and contact fatigue ...

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Номер записи: 62
02-05-2023 дата публикации

Turbine rear casing structure part fatigue test loading determination and test method

Номер: CN116046408A
Автор: XIE LIJUAN, BAI HANSONG, CONG PEIHONG, CAO MAOGUO, CHENG RONGHUI, LYU TIANBO, SONG YANG
Принадлежит:

The invention relates to a turbine rear casing structure part fatigue test loading determination and test method, and the method comprises the steps: carrying out the fatigue simulation analysis of a turbine rear casing structure in a design state, and obtaining the stress distribution of the turbine rear casing structure in the design state; determining a fatigue key examination part of the turbine rear casing structure according to the stress distribution of the turbine rear casing structure in the design state; in the part test state, fatigue simulation analysis is conducted on the turbine rear casing structure, and stress distribution of the turbine rear casing structure in the part test state is obtained; adjusting the magnitude of the loading load of each test load loading point on the turbine rear casing structure in the part test state, so that the stress distribution of the fatigue key examination part of the turbine rear casing structure in the part test state and the stress distribution ...

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Номер записи: 63
04-08-2023 дата публикации

Pre-whirl air supply system with labyrinth clearance control function

Номер: CN116537895A
Автор: CHENG RONGHUI, XU LIANQIANG, HE JIN, HU YANLI, KANG YAJIE, MA JIANDONG, ZOU MI, SU CHANGMING
Принадлежит:

The invention relates to the technical field of aero-engines, and discloses a pre-whirl air supply system with labyrinth clearance control, a first honeycomb ring is arranged to be of a floating ring type structure, so that the first honeycomb ring deforms freely and independently, and the problem that the first honeycomb ring is affected by deformation of turbine stator blades and an inner supporting ring is solved; the first impact holes are formed in the inner supporting ring and used for impacting and cooling the first honeycomb ring, the temperature of the first honeycomb ring can be reduced, the thermal expansion amount of the first honeycomb ring is reduced, and the gap of the sealing comb teeth is controlled. In addition, the temperature response speed of the first honeycomb ring can be increased, the stability of the thermal expansion amount in the transition state process is kept, and it is ensured that the gap of the sealing comb tooth structure at the rim can be kept stable ...

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Номер записи: 64
14-07-2023 дата публикации

Temperature protection method and system for engine based on fuel supply control

Номер: CN116428066A
Автор: CHENG RONGHUI, ZHANG XUEDONG, JIANG FANSHENG, ZHANG ZHISHU, ZHANG SHENGLIANG, CHEN ZHONGGUANG, BING LIANXI, ZHU ZHENKUN, YUAN JILAI, XIA YU, BAI SHUAIYU, SHI LEI, RUAN WENBO, CHEN ZEHUA, BIAN JIALIANG, ZHANG ZHICHENG, HAOBISI GALATU, LYU ANQI
Принадлежит:

The invention relates to a temperature protection method and system for an engine based on fuel supply control, and the method comprises the steps: setting the yellow warning temperature T4'of the total temperature T4 of an outlet of a main combustion chamber; setting red limiting temperatures T4 ''and T4'' gt of the main combustion chamber outlet total temperature T4; t4 '; when the total temperature T4 of an outlet of the main combustion chamber exceeds the yellow warning temperature T4'and the duration time exceeds t ', oil cutting is conducted according to the formula W = W' * e-kt, fuel oil supply of the main combustion chamber is reduced, and W is fuel oil supply of the main combustion chamber; w'is fuel oil supply of the main combustion chamber at the beginning of oil cutting; k is an oil cutting depth coefficient of the main combustion chamber; t is the duration of oil cutting of the main combustion chamber and is stopped when the total temperature T4 of an outlet of the main combustion ...

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Номер записи: 65
28-07-2023 дата публикации

Thermal management system model of aviation gas turbine engine

Номер: CN116500913A
Автор: CHENG RONGHUI, LI LINGHAN, WU XIN, JIA LINYUAN, LENG ZIHAO
Принадлежит:

The invention provides an aero-gas turbine engine thermal management system model, and belongs to the technical field of aero-engine thermal management, and the system model comprises an engine controller sub-model which is used for calculating a current engine geometric adjustable variable and a fuel oil flow according to a current atmospheric environment condition and an engine control instruction, the hot oil return flow is calculated according to the limiting temperature corresponding to the limiting temperature point determined according to the temperature bearing characteristic of the engine and the safety requirement and the actual feedback temperature; the engine performance calculation sub-model is used for calculating engine performance parameters required by the engine mechanical system sub-model and/or the engine fuel system sub-model; the engine mechanical system sub-model is used for calculating engine bearing heat generation and air heat exchange outside a bearing cavity; ...

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Номер записи: 66
30-05-2023 дата публикации

Combustion chamber of gas turbine

Номер: CN116182195A
Автор: CHENG RONGHUI, LIU BAO, SUN XIAOFENG, QIN LEI, ZHANG GUANGYU, LI LEI, WANG XIAOYU, CAO MAOGUO, XU XINGPING, MA HONGYU, CONG PEIHONG, ZHANG ZHIXUE
Принадлежит:

The invention belongs to the technical field of gas turbine combustion chamber design, and particularly relates to a gas turbine combustion chamber which comprises an annular combustion chamber outer wall. The combustion chamber inner wall is annular and is arranged in the outer wall; the combustion chamber heads are circumferentially distributed at the inlet part between the combustion chamber outer wall and the combustion chamber inner wall, and the peripheries of the combustion chamber heads are provided with annular switching sections; each switching section is connected between the combustion chamber outer wall and the combustion chamber inner wall, a plurality of cooling holes are formed in each switching section, and the cooling holes close to the combustion chamber outer wall have relatively large density and/or radius.

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Номер записи: 67
18-08-2023 дата публикации

Method for constructing flow calculation simulation model of engine turbine guider

Номер: CN116611174A
Автор: CHENG RONGHUI, ZOU MI, FU YU, GUO WEN, HU YANLI, XU LIANQIANG, MA JIANDONG, SU CHANGMING, LIANG JINHUA
Принадлежит:

The invention belongs to the field of engines, relates to a turbine guider design technology, and provides a method for constructing a flow calculation simulation model of an engine turbine guider, which comprises the following steps of: modeling each part of the guider into a loss unit; calculating a flow test value output by each loss unit according to an engine design state complete machine test result; an air entraining cavity, an exhaust cavity and all loss units are adopted to build a flow simulation calculation network; acquiring boundary conditions of each loss unit; constructing a guider flow calculation simulation model by using the flow simulation calculation network and the boundary conditions; and performing iterative correction on the flow simulation calculation model to obtain a corrected flow simulation calculation model. According to the construction method, the flow of the turbine guider can be rapidly and efficiently calculated, great significance is achieved for evaluation ...

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Номер записи: 68
28-04-2023 дата публикации

Afterburner of aero-engine

Номер: CN116025926A
Автор: CHENG RONGHUI, LIU BAO, LIU SHENG, SUN XIAOFENG, XU XINGPING, CAO MAOGUO, LI LEI, MA HONGYU, CONG PEIHONG, ZHANG ZHIXUE, YOU QINGJIANG, BAO ZHANYANG, CHEN DI, JIANG YU, LIU WEICHEN, ZHOU CHUNYANG, GUO HONGTAO, ZHANG XIAOYU, PAN XINZHENG, YAN JINSHENG
Принадлежит:

The invention belongs to the technical field of general combustion engines and aero-engines, and particularly relates to an afterburner of an aero-engine. The confluence ring is arranged in the outer casing, an outer culvert is formed between the confluence ring and the outer casing, and a plurality of cooling air inlets distributed in the circumferential direction are formed in the side wall of the tail edge of the confluence ring; the inner cone is arranged in the confluence ring, and an inner connotation is formed between the inner cone and the confluence ring; the radial stabilizer outer walls are connected into the confluence ring and point to the inner cone, the cross sections of the radial stabilizer outer walls are V-shaped, openings of the radial stabilizer outer walls face backwards, and tooth grooves distributed in a staggered mode are formed in the tail edges of the two side walls; the radial stabilizer inner walls are connected into the confluence ring and correspondingly connected ...

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Номер записи: 69
29-08-2023 дата публикации

Checking method for performance simulation model of multi-stage axial flow compressor

Номер: CN116663202A
Автор: CHENG RONGHUI, ZHANG JUN, ZHANG SHAOPING, ZHOU ZHENG, TIAN XIAOHONG
Принадлежит:

The invention belongs to the technical field of multi-stage axial flow gas compressor performance simulation, and provides a multi-stage axial flow gas compressor performance simulation model checking method which comprises the steps that S1, a plurality of key models in a gas compressor performance simulation model are extracted; s2, performing a performance test on the gas compressor to obtain a performance test result; and S3, based on a performance test result, the compressor performance influence importance sequence and simulation calculation of the compressor performance simulation model, checking the compressor performance simulation model for multiple times, sequentially obtaining checking coefficients of all key models, and obtaining a final compressor performance simulation model. According to the final compressor performance simulation model checked by the checking method, the calculation precision of a compressor performance simulation program can be greatly improved, and the ...

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Номер записи: 70
11-04-2023 дата публикации

Rotor shaft squirrel cage force measurement structure calibration device and method

Номер: CN115950581A
Автор: CHENG RONGHUI, YANG FEIBING, WEI SONGHAN, GAO WEIQIANG, TENG GUANGRONG, LI SHENGXIANG
Принадлежит:

The invention relates to the technical field of aero-engine and gas turbine tests, and discloses a rotor shaft squirrel cage force measurement structure calibration device and method.While an axial force application assembly is used for applying multiple sets of axial force to a rotor shaft, a radial force application assembly is used for applying dynamic radial force and/or static radial force to the rotor shaft; acquiring strain values on the elastic ring force measuring piece in the application process of each group of axial force and radial force; and carrying out curve fitting by taking the applied axial force as a dependent variable, the axial force strain value obtained corresponding to the elastic ring force measuring piece and the dynamic and/or static radial force as an independent variable, and determining a strain-load calibration curve of the elastic ring force measuring piece. The axial load and the radial load are simultaneously applied to the elastic force measuring piece ...

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Номер записи: 71
06-06-2023 дата публикации

High-reliability engine pressing joint sealing structure

Номер: CN116221533A
Автор: CHENG RONGHUI, LU WENRUI, JIA ZHIYUAN, LI JIAN, YANG RANSHENG, LIANG JING, SUN JIAWEI, YANG YONGHUA, DING SONG, DENG RUTAN
Принадлежит:

The invention belongs to the field of design of parts of aero-engines, and relates to a high-reliability engine pressing connector sealing structure which comprises a pressing connecting nozzle, an oil inlet connecting nozzle, an oil outlet connecting nozzle and a spherical sealing plug. The spherical sealing plug comprises a plug cap and a nut; the plug cap and the nut are coaxially arranged, the plug cap is connected into the nut in a sliding mode, and the nut is connected into the pressing connecting nozzle in a threaded mode. The plug cap comprises a ball head inserted in the pressing connection nozzle, the diameter of the ball head is larger than that of an inner cavity of the nut, the pressing connection nozzle is provided with a conical ring surface on the inner wall of the inner side of the nut, and the ball head can abut against the inner wall of the conical ring surface; during assembly, tightening force applied to the nut by a torque wrench is converted into pressing force to ...

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Номер записи: 72
25-08-2023 дата публикации

Design method and system of rotor and stator rim auxiliary sealing hole

Номер: CN116644523A
Автор: CHENG RONGHUI, MA JIANDONG, ZHANG SHAOPING, GUO WEN, HU YANLI, XU LIANQIANG, ZOU MI, SU CHANGMING, LIANG JINHUA
Принадлежит:

The invention relates to the technical field of aero-engine turbine blade design, and discloses a design method and system for a rotor and stator rim auxiliary sealing hole. A rim auxiliary sealing hole structure communicating with a sealing cavity is additionally arranged at the position of a lower edge plate of a stator guide vane; according to data such as main runner gas temperature, sealing airflow temperature, sealing airflow flow and rotor rotating speed under the design working condition of the aero-engine; according to the method, the characteristic temperature in the sealing cavity with the auxiliary sealing hole is rapidly and efficiently obtained in engineering evaluation by using the temperature dimensionless sealing efficiency related to the rim efficiency, the flow evaluation of the auxiliary sealing hole structure can be conveniently and efficiently performed in the engineering design stage, and the structural parameter design of the auxiliary sealing hole can be guided.

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Номер записи: 73
11-04-2023 дата публикации

Method for calculating hole edge stress of turbine blade film hole

Номер: CN115952622A
Автор: CHENG RONGHUI, LI KUN, WU JIAN, ZHANG SHAOPING, HE JIN, PANG YANLONG, TIAN HONGYU, YANG YUANLONG
Принадлежит:

The invention provides a hole edge stress calculation method for a turbine blade film hole. The hole edge stress calculation method comprises the steps of obtaining a nominal stress component of a film hole position under a blade overall coordinate system; the nominal stress component is converted to the constructed film hole local standard coordinate system; calculating a two-dimensional elliptical hole theoretical solution and a three-dimensional elliptical straight hole finite element simulation solution under the converted nominal stress component unit load; defining a three-dimensional stress influence coefficient; constructing a three-dimensional stress influence coefficient mathematical model; and correcting a two-dimensional elliptical hole theoretical solution under a unit load by adopting the nominal stress component obtained after conversion and the three-dimensional stress influence coefficient, so as to calculate the stress at different hole edge positions of the turbine blade ...

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Номер записи: 74
11-04-2023 дата публикации

Thermal insulation tile casting shape righting mold

Номер: CN115945541A
Автор: CHENG RONGHUI, TIAN WEI, FU YU, HE JIN, WANG QING, ZHANG SHAOPING, ZHONG YAN, LIU YANFEI, LU JING, ZHANG HANG, WU JUNJIE, BAI YUNRUI
Принадлежит:

The invention provides a heat insulation tile casting shape righting mold, and belongs to the technical field of heat insulation tile machining, the heat insulation tile casting shape righting mold specifically comprises an upper mold, a lower mold, a positioning column and a wedge-shaped block, the top surface of the lower mold is provided with a cavity surface used for being matched and attached to the front face and the side edge of a heat insulation tile, and the bottom surface of the upper mold is provided with an arc-shaped surface used for being attached to the back face of the heat insulation tile; through holes penetrating through the top face and the bottom face are formed in the two ends of the upper die and the two ends of the lower die respectively, a flange protruding outwards in the radial direction is arranged at one end of the positioning column, a wedge-shaped hole is formed in the end, away from the flange, of the positioning column, and the wedge-shaped block is matched ...

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Номер записи: 75
04-07-2023 дата публикации

Method for determining radar stealth performance after damage of aero-engine radar stealth material

Номер: CN116381640A
Автор: CHEN HANZE, CHENG RONGHUI, SHANG SHOUTANG, CAO MAOGUO, YANG SHENGNAN, WANG QUN, SUN XU, WU FEI
Принадлежит:

The invention relates to a method for determining radar stealth performance after damage of an aero-engine radar stealth material, and the method comprises the steps: carrying out the simulation calculation on the basis of setting that the radar stealth material on each part has a plurality of different damage areas; the radar stealth performance of the aero-engine under the combination of different damage areas of the radar stealth material on each component is obtained, a regression equation associated between the radar stealth performance of the aero-engine and the damage area of the radar stealth material on each component is trained, a regression coefficient is obtained, and the aero-engine is selected; according to the method, the damage area of each part is measured, the radar stealth performance is obtained through simulation calculation, the accuracy of a regression equation is checked, and after the check is passed, the radar stealth performance of a target aero-engine is quickly ...

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Номер записи: 76