80 results about "Bypass ratio" patented technology

A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

The invention discloses a large-bypass-ratio turbofan engine fleet full-working-condition abnormity monitoring method based on a self-attention reconstruction network. According to the method, sliding window sampling is carried out on multivariable flight parameter time sequence data, and normalization is carried out by adopting a mean value standard deviation; and learning the health baseline by using a reconstruction network comprising a self-attention encoder and a decoder, outputting a target parameter reconstruction value, and taking a residual error between a measured value and the reconstruction value as an abnormal measure. Aiming at the problem that the distribution difference of residual errors in different flight stages and different state intervals is obvious, a quantile threshold value is estimated under the three-dimensional conditions of an engine individual, the flight stages and a state sub-box based on healthy residual errors, and a hierarchical backspacing mechanism is set to cope with strange individual or small sample condition intervals of different fleets; and during online monitoring, in combination with a continuous duration consistency rule, an alarm is triggered only when the overrun lasts no less than K sampling periods. According to the invention, robust anomaly monitoring can be realized under all working conditions and individual difference conditions, false alarms are reduced, and the engineering deployability is improved.

A gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10. The thrust to power airflow ratio is a ratio of airflow through a bypass passage over the turbomachine plus airflow through the fan duct to airflow through the core duct. The core bypass ratio is a ratio of airflow through the fan duct to airflow through the core duct. The fan duct includes an exhaust nozzle having a plurality of chevrons disposed at its aft end to define an exhaust outlet.

The application provides a two-stage annular mixer of a booster combustion chamber and a bypass ratio self-adaptive matching method, and relates to the technical field of an aero-engine.The two-stage annular mixer of the booster combustion chamber comprises a casing, a first-stage annular mixer arranged in the casing, an outer bypass formed between the first-stage annular mixer and the casing, a second-stage annular mixer arranged in the casing and located inside the first-stage annular mixer, a split flow channel formed between the first-stage annular mixer and the second-stage annular mixer, a mixing channel formed by the casing and the second-stage annular mixer, and the outer bypass and the split flow channel are communicated with the mixing channel.In the embodiment of the application, a small mixer mixing effect is realized under a small overflow loss and a flow separation risk, a low cooling gas temperature is ensured, and a low flow resistance and high-efficiency stable working requirement of a wide outer / inner bypass range can be met.

The invention provides a small-bypass-ratio aero-engine multi-working-condition low-temperature starting test simulation device and method, and the device comprises a test cabin which is internally provided with an engine; the engine inlet mixer is arranged at the inlet of the test cabin, the inlet of the engine inlet mixer is communicated with the cooling and pressure regulating device, and the outlet of the engine inlet mixer is connected with the inlet of the engine; the air extractor set is communicated with the outlet of the test cabin; an inlet of the adjusting bypass is formed in the side wall of the end, close to the outlet, of the engine inlet mixer, and an outlet of the adjusting bypass communicates with the air extractor unit. According to the invention, the maturity of the air low-temperature starting test of the small-bypass-ratio aero-engine is improved, the accuracy and the test efficiency of low-temperature test simulation are improved, the test cost is reduced, and the requirements of sea level or plateau low-temperature starting test simulation of the engine can be effectively met.

The present disclosure relates to the technical field of aircraft design, and particularly provides a turbo ram combined power and thermal management multi-modal collaborative system for an aero-engine, comprising a first shell, a second shell, a ducted fan unit and a liquid metal cooling unit; the first shell is arranged at the outer periphery of the aero-engine; a turbine engine is arranged in the first shell; the second shell is arranged below the turbine engine, and an air inlet channel is formed between the upper side of the second shell and the first shell, and a ram channel is formed between the lower side of the second shell and the first shell; the ducted fan unit is arranged in the ram channel and can be unfolded and folded, and when unfolded, the ducted fan unit can increase the bypass ratio of the aero-engine, and when folded, the ducted fan unit is withdrawn into the second shell; the liquid metal cooling unit is connected to the first shell through a first pipeline and can cool the first shell. Through the collaborative design of the ducted fan unit, the liquid metal cooling unit and the liquid metal cooling unit, the present disclosure can effectively solve the problems of thrust gap and thermal protection of the aero-engine.

The application belongs to the technical field of aero-engine mounting structure design, and particularly relates to a split type connecting structure of an engine thrust joint and a torque mounting joint, and especially to a split type design which completely decouples the engine thrust transmission function and the torque and moment balance function in structure. The thrust transmission function and the torque balance function are completely separated, the problems of complex load path, large structure weight and poor maintainability of a traditional integrated engine rear mounting joint are solved, the split type connecting structure has the advantages of clear force transmission path, structure optimization and weight reduction, convenient maintenance, high reliability and the like, and is particularly suitable for the installation of a modern high-bypass-ratio turbofan engine.

The invention belongs to the technical field of aero-engine installation structure design, and particularly relates to a split type connection structure of an engine thrust connector and a torque installation joint, in particular to a split type design for structurally and completely decoupling an engine thrust transmission function and a torque and moment balance function. A thrust transmission function and a torque balance function are completely separated, the problems that a traditional integrated engine rear installation joint is complex in load path, large in structural weight, poor in maintainability and the like are solved, and the rear installation joint has the advantages of being clear in force transmission path, optimized in structure, capable of reducing weight, convenient to maintain, high in reliability and the like and is particularly suitable for installation of a modern large-bypass-ratio turbofan engine.

This invention discloses a porous rectifying structure for the afterburner of a variable cycle engine, comprising a rectifying orifice plate installed in an annular channel between the turbine outlet and the flame stabilizer. The orifice plate is symmetrically arranged along its radial central axis, with the clogging area facing the turbine inflow and axially aligned with the heat shield inlet. The orifice plate height is 2 / 3 of the channel height between the central cone and the heat shield. The plate has uniformly distributed pores with an equivalent diameter of 10-20 mm along the radial direction, with a clogging ratio of 5%-25%. This invention forms a multi-stage mixing interface through the radially gradient porosity pore arrangement, promoting shear mixing of the high-temperature internal combustion gas and the external bypass gas. Simultaneously, the static pressure difference across the orifice plate drives the external bypass gas flow into the heat shield cooling channel, increasing the cooling flow rate, improving the uniformity of the upstream velocity field of the flame stabilizer, eliminating the recirculation zone inside the heat shield, reducing the thermal load on the casing wall, and controlling the total pressure loss to within 3%, achieving high-stability combustion and reliable thermal protection under a wide bypass ratio.

A gas turbine engine for an aircraft includes an engine core with a turbine, combustor, compressor, and core shaft connecting the turbine to the compressor. The engine includes a fan upstream of the core and driven by the shaft. The engine includes a nacelle surrounding the fan and core and defining a bypass duct radially outside of the core. The bypass ratio, defined as the ratio of mass flow rate through the bypass duct to mass flow rate through the core at cruise conditions, is greater than or equal to 4. The engine includes a plurality of actuators. At least one of the actuators is configured to actuate a bleed air valve. The engine includes a fuel supply system that supplies fuel for combustion in the combustor, and supplies fuel to fueldraulically drive the at least one actuator configured to actuate the bleed air valve.

The invention discloses a static performance testing device for a grating type thrust-reverser nacelle test model, belongs to the technical field of wind tunnel special tests, and aims to solve the problem that an existing testing device can only test the influence of external culvert pressure ratio changes on the performance of a thrust-reverser nacelle. Comprising an external balance, a gas supply rectifying device and a pressure measuring device, the gas supply rectifying device is provided with an outer culvert rectifying cavity and an inner culvert rectifying cavity, the pressure measuring device is provided with an outer ring of an inner ring, the inner ring and the outer ring are connected through a plurality of outer culvert total pressure measuring rakes, and a plurality of inner culvert total pressure measuring rakes are arranged on the inner circumference of the inner ring. The outer culvert flow meter is communicated with the outer culvert rectifying cavity through the external balance, an outlet of the outer culvert rectifying cavity is connected with the outer ring through a pipeline, the inner culvert flow meter is communicated with the inner culvert rectifying cavity through the external balance, and an outlet of the inner culvert rectifying cavity is connected with the inner ring through a pipeline. According to the invention, the thrust performance of the inner culvert and the thrust performance of the outer culvert nozzle can be measured independently, and a powerful design verification means is provided for the design of the large-bypass-ratio thrust reverser nacelle.

The invention belongs to the technical field of engine structure design, and particularly discloses a bird-impact-resistant layered hybrid asymmetric layering structure for the tail edge of a composite material blade, and the bird-impact-resistant layered hybrid asymmetric layering structure is used for forming a fan blade of a large-bypass-ratio turbofan engine. The bird-impact-resistant layered hybrid asymmetric layering structure for the tail edge of the composite material blade comprises a bullet-facing surface, a central area and a bullet-backing surface, a hybrid region is arranged in the central region, and fibers which are different from other regions and have high toughness are arranged in the hybrid region, so that a hybrid fiber laying layer structure is formed; the mixed area is not symmetrical about the geometric middle plane of the layering structure in the thickness direction, so that an asymmetric layering structure is formed. The problems that the rigidity and toughness of a single material system of an existing composite material fan blade laying layer are contradictory, and the design freedom degree is limited by symmetrical laying layer design are solved.

The application belongs to the technical field of aero-engine safety design and whole machine vibration, and provides a whole machine dynamic response analysis method after a blade fly-off of an aero-engine. When an aero-engine has a blade fly-off and other inclusive accidents, the dynamic response of the whole engine is complex and changeable, and in serious cases, the safety of the engine and even the airplane will be threatened. Therefore, the application provides a whole machine dynamic response analysis method after a blade fly-off of a high-bypass-ratio turbofan engine. The application mainly considers the influence of the collision and friction between the low-pressure rotor and the fan case after the fly-off of a large-diameter fan blade on the whole engine, and the influence of the large unbalanced rotating inertia excitation on the fulcrum dynamic load during the reduction process after the fly-off of the engine blade.

A turbofan engine capable of reducing the air flow of an outer duct comprises multiple stages of fans, the outer duct with the reduced cross section area, an axial-flow air compressor, a combustion chamber, a turbine, an exhaust nozzle and the like, and the duct ratio does not exceed 1.0 or about 1.0; during operation after starting, the air flow of the outer duct with the cross section area reduced is reduced compared with the air flow of the outer duct with the cross section area unchanged, the air flow entering the core engine is increased due to the fact that the flow of the fan is unchanged, and the air flow is increased through operation of the engine. Compared with a conventional turbofan engine, the total supercharge ratio and the thrust-weight ratio of the turbofan engine are improved when the number of stages of a fan, an axial-flow compressor and a turbine, the fan diameter, the temperature in front of the turbine and the cross sectional area of an outer duct of the turbofan engine are not changed within the safe operation range of the mechanical strength and the high-temperature resistance of the turbofan engine, and compared with the conventional turbofan engine with the outer duct with the reduced cross sectional area, the bypass ratio of the outer duct is the same; however, the stages of the fan, the axial-flow compressor and the turbine are not changed, and the front temperature of the turbine is not changed.

Gas turbine engine (10) for an aircraft. The gas turbine engine comprising: a rich-burning, fast-cooling, lean-burning (RQL) combustor device (16) having a number of fuel spray nozzles (403, 404) in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6.A depleted-PMD cruise nvPM emission index ratio is defined as: where: EIcruise (depleted) is defined as: EImaxTO is the loss-corrected system nvPM emission index in mg / kg of the gas turbine engine (10) when operating at approximately 100% of available thrust under given operating conditions; EIclimb is the loss-corrected system nvPM emission index in mg / kg of the gas turbine engine (10) when operating at approximately 85% of available thrust under given operating conditions; and BPR is the bypass ratio of the gas turbine engine (10). The depleted-PMD cruise nvPM emission index ratio is less than 0.2. The gas turbine engine (10) is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles (403, 404).The invention also discloses a method of operating the gas turbine engine. Figure for the abstract: [Figure 6].

A gas turbine engine for an aircraft includes an engine core including a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan upstream of the engine core and driven by the core shaft. The engine includes a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, where the bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions, is greater than or equal to 4. The engine includes a plurality of actuators and a fuel supply system. The fuel supply system is arranged to supply fuel for combustion in the combustor, and to supply fuel to fueldraulically drive at least ten of the plurality of actuators.

The application belongs to the technical field of engine structure design, and particularly discloses a hybrid asymmetric layered structure of a composite blade tail edge for resisting bird strikes, which is used for forming a fan blade of a large-bypass-ratio turbofan engine, and comprises a bullet-facing surface, a central region and a bullet-facing back surface; a hybrid region is arranged in the central region, and high-toughness fibers different from those in other regions are arranged in the hybrid region to form a hybrid fiber layered structure; the hybrid region is not symmetric about the geometric median plane of the layered structure in the thickness direction, thereby forming an asymmetric layered structure. The application solves the contradiction between the stiffness and toughness of the single material system of the existing composite fan blade layered structure and the problem that the symmetric layered design limits the design freedom.

The invention discloses an automatic butt joint assembly system for a large-bypass-ratio aero-engine, and relates to the technical field of aero-engine assembly, the automatic butt joint assembly system comprises a six-degree-of-freedom platform module, a winch module, a precise posture adjusting module, a travelling crane trolley module and a binocular vision detection system; the upper portion of the traveling crane trolley module is erected above a steel rail through a force bearing wheel and a driving wheel, the two ends of the lower portion of the traveling crane trolley module are connected with the precise posture adjusting module module through hinges, the precise posture adjusting module module is connected with the winch module through a linear guide rail, and the winch module is connected with the six-degree-of-freedom platform static platform through a steel wire rope. The movable platform of the six-degree-of-freedom platform is connected with a product to be butted through a tool; the travelling crane trolley module and the winch module are coarse adjustment mechanisms, the precise posture adjustment module and the six-degree-of-freedom platform module are fine adjustment mechanisms, and the binocular vision detection system is a detection feedback module. The problems that an existing industrial robot and a laser guiding system are insufficient in positioning precision, large in error accumulation and the like in the large-bypass-ratio aero-engine butt-joint assembly process are solved, and automatic butt-joint assembly with high precision, high stability and high adaptability is achieved.

A gas turbine engine includes: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct, the gas turbine engine defining a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

The application provides an inner-content nozzle and a large-bypass-ratio aero-engine, and relates to the technical field of aero-engine structures. The inner-content nozzle comprises a fixed part and a moving part, and the fixed part forms a first airflow channel. The moving part is distributed along an axial direction with the fixed part, and the moving part can slide along the axial direction relative to the fixed part, thereby realizing switching of the inner-content nozzle between a nozzle mode and a reverse thrust mode. The moving part is provided with a cascade assembly. When the inner-content nozzle is in the nozzle mode, the cascade assembly is shielded by a fixed pipe wall, and airflow in the first airflow channel enters a second airflow channel and is sprayed out. When the inner-content nozzle is in the reverse thrust mode, the cascade assembly moves downstream of the fixed pipe wall, and the airflow in the first airflow channel is sprayed out through a reverse thrust airflow channel in the cascade assembly, so as to generate a reverse thrust, thereby meeting the requirements of ground reverse driving or rapid landing in the air.

The invention relates to the technical field of aero-engines, and discloses an adjustable mixing stabilizer structure, a control method and an applied afterburner. The adjustable mixing stabilizer structure comprises a stabilizer main body, an air entraining cover plate and a first driving mechanism; when the duct is large, the first driving mechanism is controlled to drive the air entraining cover plate to rotate along the hinge shaft, so that the throttling area of the mixing channel is increased, the flow of cooling air passing through the mixing channel is increased, the flow of the cooling air of the heat insulation screen is reduced, the content of oxygen in the duct is increased, and the combustion efficiency of the afterburner is improved; when the duct is small, the first driving mechanism is controlled to drive the air entraining cover plate to rotate along the hinge shaft so as to reduce the throttling area of the mixing channel, so that the flow of cooling air passing through the mixing channel is reduced, the flow of cooling air of the heat screen is increased, the flow of cooling air is guaranteed, and it is guaranteed that the relative cold air amount entering the heat screen is within a stable and controllable range; and safe operation requirements of the afterburner are met.

The application belongs to the field of aircraft design and is a method for subsonic cruise state fuel consumption of a small-bypass-ratio turbofan engine. The method comprises the following steps: obtaining initial performance record results of the whole machine in a subsonic cruise state, analyzing and judging whether the working state of the small-bypass-ratio turbofan engine is in an optimal state, and optimizing a preliminary design of a cruise fuel consumption debugging scheme to determine a preliminary adjustment scheme of an optimal fan guide vane angle a1 and a preliminary adjustment scheme of an optimal compressor guide vane angle a2; further adjusting a nozzle throat area A8 and a nozzle outlet area A9 to different areas to determine an optimal nozzle throat area A8 and an optimal nozzle outlet area A9; and finally judging installation fuel consumption under the condition of the whole aircraft to obtain an optimal installation fuel consumption state when the installation fuel consumption is optimal. The method systematically and normatively proposes a debugging step for optimizing subsonic cruise state fuel consumption, which can effectively reduce engine installation fuel consumption.

The afterburner hydrogen supply structure comprises a fuel header pipe assembly and a stabilizer, and the fuel header pipe assembly is connected with the stabilizer through a first connecting supporting pipe and a second connecting supporting pipe. The multi-path fuel header pipe with the rigid connecting rod of the structure can flexibly adjust fuel supply in the transient state of ignition, the middle working condition and the full working condition and in the variable working condition, and the Y-shaped flame stabilizer can achieve low-blockage-ratio, efficient and stable combustion under the condition of the large bypass ratio. The fuel header pipe and the flame stabilizer are rigidly connected through the ignition connection supporting pipe, and meanwhile the ignition fuel multi-point distribution supply and integrated connection requirements are met. And the combustion stability and the temperature distribution uniformity are good, the structural manufacturability is good, violent mechanical vibration is not prone to occurring, the assembly precision is high, and array arrangement can be achieved according to requirements.

The invention belongs to the technical field of engine compression part adjustable stator blade design, and particularly relates to an engine multi-duct front compression part outlet adjustable stator blade and a design method thereof.The design method comprises the steps that numerical simulation is conducted on a front compression part and a rear compression part, and flow dividing streamlines of the front compression part under the large-bypass-ratio working condition and the small-bypass-ratio working condition are obtained; the blade type of an outlet-adjustable stator blade of a front compression component is improved, and the outlet pre-rotation angle between a flow dividing streamline of the front compression component under the working condition of a large bypass ratio and a flow dividing streamline of the front compression component under the working condition of a small bypass ratio of the outlet-adjustable stator blade is changed, so that the flow dividing streamline of the front compression component under the working condition of the small bypass ratio is improved; an outlet of the front compression component can adjust the angle of a stator blade, and sensitivity to air flow entering an inner duct is achieved.

1. Name of the designed product: model of a high-bypass turbofan engine (large-bypass counter-impeller). 2. Use of the designed product: for a model of a high-bypass turbofan engine (large-bypass counter-impeller). 3. Design points of the designed product: in shape. 4. Picture or photo best indicating the design points: perspective view 1.

The application belongs to the field of aero-engine design, and is a high-low pressure turbine matching design method for a cross-generation small-bypass-ratio turbofan engine. The method comprises the following steps: firstly, determining the initial constraint conditions of the low-pressure turbine matching design of a high-performance advanced engine; then, determining the constraint conditions of the high-pressure turbine matching design of the high-performance advanced engine through the high-low pressure turbine matching design; after the matching is completed, carrying out component processing and assembly; after the assembly is completed, carrying out high-low pressure turbine component tests and whole-process engine tests to obtain whole-process engine measurement test results; according to the whole-process engine measurement results, evaluating the high-low pressure turbine matching work, determining the problems existing in the high-low pressure turbine matching, and carrying out targeted structural improvement on the high-low pressure turbine and whole-process tests again; through repeated correction, the design requirements are met. The method can quickly and accurately realize good matching work of the high-low pressure turbine, reduce the exhaust temperature, and solve the problem of ablation of high-temperature components.

The invention discloses a wide-range working condition jet stabilizing structure and an oil supply method.The structure comprises a jet stabilizing system case, a first jet stabilizing component, a second jet stabilizing component, an inner cone, a concave cavity ignition oil injection rod, a mixer adjusting piece and a mixer case, and the first jet stabilizing component, the second jet stabilizing component and the mixer case are connected with the inner wall of the jet stabilizing system case; the inner cone is installed on an engine turbine support, the inner cone and the jet stabilizing system casing are coaxially arranged, the concave cavity ignition fuel injection rod is installed in an inner cavity of the inner cone, and the mixer adjusting piece is connected with the mixer casing. Five oil injection rods are arranged, II, III and IV header pipe oil injection rods, a first stable injection component and a second stable injection component are integrally designed, and the arranged I oil injection rod determines whether to work or not according to the range of the inner culvert total temperature and the bypass ratio; and the requirements for oil supply, efficient combustion organization and stable work of the afterburner of the variable cycle aero-engine under the wide-range conditions that the bypass ratio is 0.2-1, the total temperature of the inner culvert is 950-1500 K and the pressure is 50-600 kPa can be met.

A gas turbine engine 10 for an aircraft comprising a rich burn quick quench, lean burn (RQL) combustor 16 having a number of spray nozzles 403, 404, in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range of 2-6, with the engine configured to provide fuel comprising sustainable aviation fuel to the spray nozzles. The engine operation is characterised by specific non-volatile particulate emission ratio beign related to the bypass ratio. The fuel spray nozzles may comprise one or more duplex nozzles and one more single flow nozzles and preferably the combustor comprises 10-14 duplex fuel spray nozzles and 4-8 signle flow fuel spray nozzles. The duplex fuel spray nozzles are arranged in groups about the circumference of the combustor and optionally wherein the groups duplex fuel spray nozzles comprise at least two groups arraned dimametrically opposite each other. A corresponding method is also claimed. Fig 6

The application belongs to the field of engine design, and particularly relates to a variable cycle engine main fuel regulation method, comprising: designing a main fuel control law architecture based on critical Mach numbers of each characteristic height, and designing main fuel control laws of different bypass ratio states at each characteristic height; carrying out optimization design of the main fuel control law during state switching to obtain a main fuel design law during state switching; carrying out matching design of the main fuel control law during fan speed limitation and flow rate drop to obtain a main fuel design law during flow rate drop; and adding a correction quantity determined by test to the main fuel design law of the whole working envelope to obtain a main fuel adaptive control law. The method fills the blank in the field of main fuel regulation design of the variable cycle engine in the whole envelope, multiple working modes and full thrust states, realizes main fuel adaptive regulation of the variable cycle engine, and improves the design capability of the aero-engine.