215 results about "Bypass ratio" patented technology

A gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3 that is defined byVGVG⁢B.VG is a gutter volume of the gutter and VGB is a gearbox volume. A gas turbine engine includes the gearbox assembly, a fan, a combustor that generates combustion gases, a turbine, a nacelle that surrounds the fan, and a steam system. The combustor and the turbine define a core air flowpath. Core air flows through the core air flowpath. The steam system that extracts water from the combustion gases, vaporizes the water to generate steam, and injects the steam into the core air flowpath to add mass flow to the core air. A bypass ratio of the gas turbine engine is in a range of 18:1 to 100:1.

The invention provides a method for simulating bypass ratio regulation and control under the complete machine condition of a turbofan engine, and the method is realized based on a double-shaft compressor tester, and belongs to the field of aero-engine tests.The method comprises the steps that control parameters under the complete machine state are determined; determining the control precision of the test state control parameters of the double-shaft compressor tester; constructing a control module which stores an adjusting rule in a whole machine state and is used for enabling the rotating speed of the fan and the rotating speed of the high-pressure compressor to be in a linkage mode in the test process; the fan pressure ratio is monitored, an outer culvert throttling valve is adjusted, then an inner culvert throttling valve is adjusted, the high-pressure compressor pressure ratio is monitored at the same time, and finally the inner culvert throttling valve and the outer culvert throttling valve are adjusted at the same time, so that the high-pressure compressor pressure ratio and the fan pressure ratio meet the requirements; the state of the double-shaft compressor tester is kept stable, and compressor parameters are collected through an inner and outer culvert exhaust collection device; the rotating speed of the high-pressure compressor or the rotating speed of the fan is adjusted according to the rotating speed difference of the fan and the high-pressure compressor in the complete machine state, and parameters of the compressor are collected after stabilization.

A gas turbine engine includes: a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath; a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio during operation of the gas turbine engine in a cruise operating mode; and a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan; wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8.

The present invention provides a design method for a dual-duct fan mixed exhaust test device, comprising the following steps: Step 1, determining the axial position of the outlet of the intermediate casing, the axial length of the mixed exhaust, and the radial height of the outlet of the mixed exhaust; Step 2, determining the shape of the mixed exhaust inner flow path based on the results of Step 1; Step 3, determining the shape of the inner-outer flow path according to the expansion ratio based on the shape of the mixed exhaust inner flow path in Step 2; Step 4, determining the shape of the mixed exhaust outer flow path and the mixing intersection point of the inner and outer ducts according to the outlet extension shape of the mode selection valve. The present invention can achieve the performance measurement of the dual-duct fan variable duct ratio test, improve the overall performance measurement of the fan component under variable duct ratio conditions, and has important guiding significance for the variable cycle whole machine test research.

The invention belongs to the field of aircrafts, relates to an aircraft power system design technology, and provides an aero-engine of a contra-rotating blade configuration and a control method, and the aero-engine comprises a gas turbine engine, contra-rotating blades and a main transmission shaft. The gas turbine engine is located in the main flow channel, the contra-rotating blades are arranged in the additional duct, an inlet adjusting piece is arranged at the inlet position of the additional duct, and an outlet adjusting piece is arranged at the outlet position of the additional duct. One end of the main transmission shaft is selectively connected or disconnected with the gas turbine engine through a clutch, and the other end of the main transmission shaft is connected with the contra-rotating blades through a connecting structure. Through the innovative power architecture design, the core problem that the thermodynamic cycle efficiency and the thrust characteristic of a traditional aero-engine under the complex flight working condition are difficult to dynamically adapt is solved, the bypass ratio is adjusted through the external contra-rotating blades, multi-working-condition self-adaptive optimization is achieved, and the reliability of the aero-engine is improved. Therefore, the aircraft power system can efficiently and reliably operate in a wide-area complex environment.

The invention relates to the technical field of aero-engines, and particularly provides a variable cycle engine duct experiment system and experiment method. A variable cycle engine duct experiment system comprises: a first combined experiment member comprising a main flow valve and a shunt ring; the second combined experiment piece comprises a first outer culvert, a second outer culvert and an ejector; the inlet of the variable capacitor second external culvert experiment piece is connected with the external culvert outlet of the shunt ring, and the outlet is connected with the second external culvert; an inlet of the first gas collection tank is communicated with an inner culvert outlet of the shunting ring; the bypass ratio adjusting valve is used for adjusting the bypass ratio at the shunting ring; an inlet of the second gas collection tank is connected with an outlet of the first gas collection tank through an inner culvert pressurization simulator, and an outlet is communicated with the first outer culvert; an inlet of the variable capacitor confluence duct experiment piece is communicated with the second outer duct; and the outer culvert throttle valve is connected with an outlet of the variable capacitor confluence culvert experiment piece. According to the technical scheme, engine duct experiments in multiple modes can be achieved, and the experiment efficiency is improved.

The invention discloses a multidisciplinary coupling adjoint optimization method for a large-bypass-ratio fan blade. The method comprises the following steps of: 1, extracting two-dimensional blade profiles of different blade heights of a large-bypass-ratio fan blade and blade modeling parameters of each blade height; 2, obtaining optimization design parameters; 3, obtaining an unsteady harmonic balance equation according to the unsteady flow field equation, and calculating a pneumatic-aeroelastic multidisciplinary coupling objective function; 4, establishing an unsteady adjoint equation, and performing iterative solution to obtain an unsteady adjoint variable; 5, obtaining sensitivity information of the aerodynamic-aeroelastic multidisciplinary coupling objective function about design parameters; step 6, updating and optimizing design parameters; 7, updating blade modeling parameters and blade geometry, and updating the computational grid by adopting a linear elasticity method; and 8, when the two norms of the sensitivity information meet the convergence standard, outputting the optimized blade modeling parameters and blade geometry. The method has the advantages of being low in calculation cost, high in optimization design automation degree and high in optimization efficiency.

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.

A gas turbine engine for an aircraft includes an engine core including a turbine, combustor, compressor, and core shaft connecting the turbine to the compressor. The engine includes a fan located upstream of engine core and arranged to be driven by core shaft. The engine includes a nacelle surrounding the fan and engine core and defining bypass duct located radially outside of engine core, where bypass ratio, defined as the ratio of the mass flow rate of the flow through bypass duct to the mass flow rate of the flow through the core at cruise conditions, is at least 4. The engine includes plurality of actuators and fuel supply system arranged to supply fuel to fueldraulically drive at least one actuator of the plurality of actuators. At cruise, fuel temperature on entry into the at least one actuator is at least 5° C. greater than fuel temperature on entry to combustor.

A gas turbine engine includes: a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath; a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio during operation of the gas turbine engine in a cruise operating mode; and a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan; wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8.

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 for an aircraft comprises an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. The engine comprises a fan located upstream of the engine core and arranged to be driven by the core shaft. The engine comprises 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 at least 4. The engine comprises a plurality of actuators and a fuel supply system arranged to supply fuel to fueldraulically drive at least one actuator of the plurality of actuators. The engine comprises two fuel-oil heat exchangers arranged to have oil and the fuel flow therethrough, the heat exchangers comprising a primary fuel-oil heat exchanger arranged to heat at least the majority of the fuel, and a secondary fuel-oil heat exchanger arranged to provide additional heat to the fuel to be supplied to fueldraulically drive the at least one fueldraulic actuator. The heat exchangers are controlled such that, under cruise conditions, a heat transfer ratio of: rateofheattransferfromoiltofuelinthesecondaryfuel−oilheatexchangerrateofheattransferfromoiltofuelintheprimaryandsecondaryfuel−oilheatexchangers has a maximum value of at least 0.35.

A gas turbine engine includes: a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath; a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio during operation of the gas turbine engine in a cruise operating mode; and a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan; wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8.

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.

This invention provides a dual-angle adjustable inlet guide vane for expanding the stability margin of an adaptive fan, comprising: a main blade D1, a main blade D2, an auxiliary blade A, a profile compensation block B, a profile compensation block C, and a connecting mechanism. The main blade D2 is mounted on the main blade D1. The leading edge of the auxiliary blade A contacts an arc-shaped groove on one side of the main blade D1. The auxiliary blade A, profile compensation block B, profile compensation block C, and main blade D2 are sequentially connected to each other through the connecting mechanism. A hollow main shaft is also provided on the tip of the main blade D1, and a secondary shaft is provided inside the hollow main shaft. The secondary shaft is connected to the tip of the auxiliary blade A. The hollow main shaft and the secondary shaft drive the blades D1, D2, and A to rotate, adjusting the pre-rotation degree at different blade height positions, significantly expanding the stability margin and working performance of the adaptive fan under full bypass ratio conditions. When the auxiliary blade A rotates, it drives the profile compensation blocks B and C to move, thereby compensating for the area loss of the entire guide vane's outer surface caused by local deflection.

The invention relates to the technical field of aero-engines, and discloses a variable camber ducted fan and an adjusting method thereof.The variable camber ducted fan comprises a first row of rotor blades and a second row of rotor blades which are installed on a hub, the first row of rotor blades conduct angle adjustment relative to the fixed second row of rotor blades, and the second row of rotor blades conduct angle adjustment relative to the fixed second row of rotor blades; the ducted ratio is dynamically changed according to the flight state by adjusting the change of the inlet angle of the first row of rotor blades. Through the variable camber ducted fan technology, a large bypass ratio is adopted during vertical take-off and landing, the lift energy efficiency is improved, and fuel consumption is reduced; and during high-speed level flight, a small bypass ratio is switched, so that efficient thrust is ensured. Besides, the ducted fan with the variable camber can dynamically adjust the ducted ratio according to the flight state to adapt to different task requirements such as vertical take-off and landing, hovering and flat flight, and it is ensured that high efficiency and low energy consumption are kept in various flight modes.

This invention provides a three-degree-of-freedom variable installation angle inlet guide vane suitable for variable cycle fans, comprising: a main blade, a first auxiliary blade, a second auxiliary blade, a main shaft, a first hollow auxiliary shaft, and a second hollow auxiliary shaft. The main blade is mounted on the main shaft, the first hollow auxiliary shaft is mounted on the first auxiliary blade, and the second hollow auxiliary shaft is mounted on the second auxiliary blade. The second hollow auxiliary shaft is sleeved on the first hollow auxiliary shaft, and the first hollow auxiliary shaft is sleeved on the main shaft, forming a three-degree-of-freedom rotating mechanism. The main blade, the first auxiliary blade, and the second auxiliary blade constitute the entire inlet guide vane. By using the main blade to adjust the flow direction within the entire flow channel, and by using the second auxiliary blade to control the adverse flow field factors caused by changes in the bypass ratio of the variable cycle engine, and by using the first auxiliary blade to control various adverse flow field factors caused by contact intake distortion, the performance of the variable cycle fan subjected to different distortion forms under different variable cycle operating conditions is improved.

The present invention discloses a control method and system for a scaled-down civil aviation turbofan engine. The scaled-down civil aviation turbofan engine includes a turbofan, a core engine, a main channel located between the turbofan and the core engine, an inner channel, and an outer channel. The turbofan uses an independent drive source. The control method for the scaled-down civil aviation turbofan engine includes the following steps: obtaining first temperature data and first pressure data at a main channel cross section of the main channel; or / and obtaining second temperature data and second pressure data at an outer channel cross section of the outer channel; calculating a bypass ratio based on the first temperature data and first pressure data, or / and the second temperature data and second pressure data; and independently adjusting the turbofan speed and the core engine throttle to ensure that the thrust operating conditions and bypass ratio of the scaled-down civil aviation turbofan engine are consistent with those of the original civil aviation turbofan engine. The present invention can provide a wide adjustment range for the bypass ratio without causing a loss in total pressure.

Fuel-driven hydraulic actuation A gas turbine engine for an aircraft includes an engine core including a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan located upstream of the engine core and configured to be driven by the core shaft. The engine includes a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outward 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 delivery system.The fuel supply system is configured to supply fuel for combustion in the combustion chamber, and to supply fuel for hydraulically fuel-driving at least ten of the plurality of actuators. Figure for abstract: Fig. 31.

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 invention discloses an airflow circulating device utilizing high-speed incoming flow in a jet flow precooling system, and belongs to the field of aero-engines. The device comprises an air inlet device, a compression device and a cooling device. The air inlet device is arranged at the front part of the jet flow precooling device or an air inlet channel of the aircraft, and controls the flow of incoming flow entering the device according to working requirements; and secondly, compression of incoming flow is achieved through a small-bypass-ratio gas compressor in the compression device, and finally high-pressure gas is released through a gas outlet of the compression device and enters the cooling device to be cooled. According to the working requirement of the aircraft, when the jet flow precooling device works, the whole airflow circulating device converts high-speed incoming flow into high-pressure low-temperature gas, and the high-pressure gas serves as a gas phase in the two-phase flow nozzle to be released. In a jet flow pre-cooling device of an original two-phase flow nozzle, the situation that the internal space of an engine is occupied and the weight is redundant due to the fact that a gas phase device is carried is effectively avoided, and the jet flow pre-cooling device has better working condition applicability.

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.

A gas turbine engine comprises: a combustion chamber (16) comprising a plurality of fuel spray nozzles (124), wherein the fuel spray nozzles (124) comprise a first subset (124A) and a second subset (124B) of fuel spray nozzles (124), wherein the combustion chamber (16) is operable in a state in which each of the fuel spray nozzles of the first subset (124A) is supplied with a higher fuel flow rate than each of the fuel spray nozzles of the second subset (124B), wherein the ratio of the number of fuel spray nozzles (124) in the first subset (124A) to the number in the second subset (124B) is in the range of 1:3 to 1:6. An nvPM emission index ratio in lean-burn cruise flight to MTO is defined as: EIR (lean) EI max TOBPR where: EI Reiseflug (mager) is defined as: EI max TO + EIA approach 2 EGG maxTOEI is the system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the given operating conditions; Steigflug is the nvPM emission index of the gas turbine engine in mg / kg, corrected for system loss, at operation with approximately 85% of the available thrust under the given operating conditions; BPR is the bypass ratio of the gas turbine engine. The nvPM emission index ratio in lean-burn cruise flight to MTO is less than 0.2. The gas turbine engine is configured to supply a sustainable aviation fuel (SAF) to the multiple fuel spray nozzles (124).

A gas turbine engine for an aircraft comprises 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 located upstream of the engine core and arranged to be driven by the core shaft. The engine includes a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outward of the engine core, wherein a 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 at least 4.The engine includes a plurality of actuators and a fuel supply system configured to supply fuel for combustion in the combustion chamber and to supply fuel for fuel-hydraulic drive of at least one of the plurality of actuators. The fuel supply system is configured to be controllable to select between: causing the fuel to actuate at least one actuator; and causing the fuel to bypass the at least one actuator based on the Sustainable Aviation Fuel (SAF) content of the fuel.

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.