154 results about "Bypass ratio" patented technology

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 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 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 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.

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.

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).

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 present invention discloses a thermal management system for a hydrogen-fueled aviation engine that coordinates multiple thermodynamic processes. The system includes a high-bypass-ratio turbofan engine structure and an intercooling and regenerative cycle structure. The intercooling and regenerative cycle structure includes a hydrogen storage tank, a boost pump, a metering valve, a regenerator, an intercooler, a first-stage radiator for bleed air, and a second-stage radiator. The system coordinates and optimizes multiple thermodynamic processes such as high-temperature fuel gas reheating, compressor interstage air cooling, turbine blade cooling air precooling, and hydrogen fuel preheating. It rationally utilizes low-temperature hydrogen fuel and the cooled combustion product - water as a heat sink, overcomes the problems of excessive pressure and relatively complex system caused by the complex bleed air structure in the traditional intercooling and regenerative layout, effectively reduces compressor power consumption, fuel consumption and overall machine waste heat loss, improves the cooling quality of turbine blade cooling air, and meets the low-energy consumption and high-reliability thermal management requirements of civil aviation engines.

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.

The invention belongs to the technical field of aircraft structural strength tests, and particularly relates to a large-bypass-ratio engine force transmission simulation device which analyzes the connection condition of a large-bypass-ratio engine and an aircraft according to the structural characteristics of the large-bypass-ratio engine. The large-bypass-ratio engine force transmission simulation device is designed to be composed of a course loading point, loading points on the two sides of the gravity center, a main force transmission shaft simulation piece, a fan casing front installation edge simulation piece, an intermediate casing inner ring simulation piece, an intermediate casing rear installation plane simulation piece, a turbine rear casing simulation piece, an installation joint simulation piece, a lifting lug simulation piece and a reverse thrust torque box front cover simulation piece. The test device is composed of 11 types such as a main body, an auxiliary body and reinforcing ribs, the mass, inertia, rigidity and installation interfaces of the engine are simulated, loading points at all positions are provided, the simulation of the load of the aero-engine is realized through the combination of single-point simple loading, and the loading problem of a large-bypass-ratio aero-engine test in an aircraft strength test is solved.

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.