246 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 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 small bypass ratio turbofan engine performance evaluation method based on model self-adjustment, comprising the following steps: defining a parameter structure for engine performance evaluation, the parameter structure comprising parameters and parameter units selected from engine test flight test data; determining a sampling moment, traversing the test flight test data, and extracting a steady-state section at the sampling moment; updating the sampling time of the test flight stage, repeatedly executing the steady-state segment extraction of the sampling time, and outputting the evaluation coefficient of the actual part characteristic recession amount of the engine; and generating an engine component performance degradation trend according to the sampling time and the engine actual component characteristic degradation amount evaluation coefficient corresponding to the sampling time. According to the technical scheme, the gas path component characteristic characterization parameters of the engine can be corrected by applying the engine self-adjusting model, table conversion processing of real sensor measurement values can be achieved, and then the gas path component performance of the engine can be effectively evaluated under multiple working conditions in the full envelope of the aero-engine.

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 turbine engine for an aircraft includes a fan that rotates to generate a volume of air, a core turbine engine, a nacelle, and a steam system. The core turbine engine includes a combustor that generates combustion gases, and a turbine including a shaft. The combustor and the turbine define a core air flowpath. The fan is drivingly coupled to the shaft. The nacelle circumferentially surrounds the fan and defines a bypass airflow passage between the nacelle and the core turbine engine. The volume of air flows into the bypass airflow passage as bypass air and flows into the core air flowpath as core air. The steam system 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 turbine engine is greater than 18:1.

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 provides a double-adjusting-angle inlet-adjustable guide vane for widening the bypass ratio working range of a fan, which comprises a first main vane, a second main vane, an auxiliary vane, a main shaft A and an auxiliary shaft B. One side of the first main vane is provided with the second main vane and the auxiliary vane, and the auxiliary vane is positioned below the bottom of the second main vane. The main shaft A is arranged at the blade bottom of the first main blade, and the auxiliary shaft B is arranged at the blade bottom of the auxiliary blade and located in the hollow cylinder of the main shaft A. The main shaft A is used for driving the first main blade and the second main blade to rotate. The auxiliary shaft B is used for driving the auxiliary blade to rotate. The first main blade, the second main blade and the auxiliary blade are adjusted through the main shaft A and the auxiliary shaft B, the flow field of an inlet of the fan is adjusted in a partitioned mode, pre-rotation of different degrees is provided for inlet airflow with different blade heights of the fan, the quality of the inlet airflow of the fan is improved, the working efficiency of self-adaptive starting is improved, and the bypass ratio change range of the self-adaptive fan is widened.

The invention provides a three-degree-of-freedom variable mounting angle inlet guide vane suitable for a variable circulating fan, which comprises 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 arranged on the main shaft, the first hollow auxiliary shaft is arranged on the first auxiliary blade, the second hollow auxiliary shaft is arranged on the second auxiliary blade, and the first auxiliary blade is arranged on the main shaft. The first hollow auxiliary shaft is sleeved with the second hollow auxiliary shaft, the main shaft is sleeved with the first hollow auxiliary shaft, a three-degree-of-freedom rotating mechanism is formed, and the main blade, the first auxiliary blade and the second auxiliary blade form the whole inlet guide blade. The main blade is used for adjusting the flowing direction in the whole flow channel, the second auxiliary blade is used for adjusting and controlling flow field adverse factors generated due to change of the bypass ratio of the variable cycle engine, and meanwhile the first auxiliary blade is used for adjusting and controlling various flow field adverse factors generated due to contact air inlet distortion. Therefore, the performance of the variable circulation fan under the action of different distortion forms under different variable circulation working conditions is improved.

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.

The invention discloses a design method for oil supply of a small and medium bypass ratio afterburner, which comprises the following steps of: acquiring environment index distribution of a flow field behind a mixer of the small and medium bypass ratio afterburner, and dividing a flow field area according to the environment index distribution to form n areas; the optimal residual air coefficient alpha optim and the air flow of the n areas are obtained; the optimal residual gas coefficient is the residual gas coefficient corresponding to the highest combustion efficiency under the conditions of certain temperature, pressure and oxygen content; according to the optimal residual air coefficient alpha i of the n areas and the air flow of the n areas, the stress application oil supply amount Wfi is distributed; wherein the optimal residual gas coefficient is close to the optimal residual gas coefficient alpha < optimal >. According to the technical scheme, the problems that the oil supply distribution plan of the afterburner with the small and medium bypass ratio cannot fully consider different environments of all areas of the combustion chamber, the optimal residual gas coefficients corresponding to combustion are different, and the oil supply amount of each area cannot be adaptively adjusted according to the change of the bypass ratio can be solved.

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.

A low-bypass-ratio turbofan engine comprises a fan, an outer duct, a gas conveying valve, a bypass pipe, an external valve, an axial-flow type gas compressor, a combustion chamber, a turbine, an exhaust nozzle and the like. During operation after starting, the air delivery valve is opened, a part of compressed air in the outer duct is sprayed into an air inlet of the engine through the bypass pipe, the air inlet pressure is increased under the condition that surge does not occur, and through acting of the engine, the supercharge ratio and the total supercharge ratio of the fan are continuously increased; in the safe operation range of mechanical strength and high-temperature resistance of the engine, the total supercharge ratio and the thrust-weight ratio are increased compared with those of the engine without a gas transmission valve, a bypass pipe and an external valve, then the gas transmission valve is adjusted, the flow of the bypass pipe is reduced, the external valve adjusts the pressure in the bypass pipe, the engine continues to operate after the increased total supercharge ratio is kept unchanged, and the engine is safe. At the moment, the total supercharge ratio and the thrust-weight ratio are increased, but the stage number of the engine and the turbine front temperature are not changed, and the air inlet pressure is not changed.

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.