Turbojet power system of multi-rotor unmanned aerial vehicle

Abstract

The invention relates to the technical field of unmanned aerial vehicles, in particular to a turbojet power system of a multi-rotor unmanned aerial vehicle, which comprises a turbojet engine, a fuel supply module, an electronic control unit, an air inlet flow guide structure, an exhaust flow guide structure, a cooling module, a damping bracket, a quick disassembly and assembly interface, a redundant backup system and an energy recovery module, according to the multi-rotor unmanned aerial vehicle, the power performance of the multi-rotor unmanned aerial vehicle is improved by replacing a traditional motor with the micro turbojet engine, so that the unmanned aerial vehicle can obtain longer endurance time under the condition of carrying the same load, and particularly, the unmanned aerial vehicle can still keep stable power output in extreme environments such as high altitude or high temperature; the problem of power attenuation caused by thin air or poor heat dissipation of the motor is solved, multi-degree-of-freedom thrust direction adjustment is achieved through the design of the deflectable nozzles, and the unmanned aerial vehicle can still keep accurate flight attitude control under severe environment conditions such as strong wind.

Description

Technical Field

[0001] The present invention relates to the technical field of unmanned aerial vehicles, and particularly to a turbojet power system for multi-rotor unmanned aerial vehicles. Background Art

[0002] Currently, the power systems of multi-rotor unmanned aerial vehicles mainly rely on the method of driving propellers by electric motors. This technology has been widely applied in various commercial and military unmanned aerial vehicle fields. However, with the continuous improvement of the requirements for the load capacity, endurance time, and high-speed maneuverability of unmanned aerial vehicles, the limitations of traditional electric power systems have gradually emerged in terms of performance.

[0003] When existing electric multi-rotor unmanned aerial vehicles operate in high-altitude or high-temperature environments, the electric motors and batteries are prone to significant power attenuation due to thin air or insufficient heat dissipation, thus seriously affecting flight stability. At the same time, the energy density of the electric system is relatively low, resulting in a sharp reduction in the endurance time of the unmanned aerial vehicle when carrying large payloads, making it difficult to meet the requirements of long-endurance missions. Although some studies have attempted to adopt hybrid power or fuel engine replacement solutions, these systems often have problems such as complex structures, slow response speeds, or excessive vibrations, and cannot meet the requirements of high efficiency and high maneuverability.

[0004] Therefore, in view of the deficiencies of the above-mentioned electric power systems in high-performance application scenarios, the present invention proposes a turbojet power system for multi-rotor unmanned aerial vehicles, which improves the high-altitude performance, load capacity, and maneuver response speed of the unmanned aerial vehicle through the high thrust-to-weight ratio characteristics of the micro turbojet engine and vector thrust control technology. Summary of the Invention

[0005] In order to overcome the deficiencies of electric power systems in high-performance application scenarios, the present invention proposes a turbojet power system for multi-rotor unmanned aerial vehicles.

[0006] The technical solution of the present invention is as follows: The turbojet power system for multi-rotor unmanned aerial vehicles includes: Turbojet engine: used to provide the main lift and attitude control ability for the unmanned aerial vehicle; Fuel supply module: used to stably supply fuel to the turbojet engine; Electronic control unit: used to monitor and adjust the operating state of the turbojet engine in real time; Intake air guiding structure: used to optimize the air intake efficiency and prevent foreign objects from entering; Exhaust air guiding structure: used to improve the exhaust thrust efficiency and reduce the infrared signature; Cooling module: used to prevent the engine from overheating; Shock-absorbing bracket: used to suppress the vibration transmission of the turbojet engine; Quick disassembly and assembly interface: used to achieve the quick replacement of the turbojet power unit; Redundant backup system: used to ensure the safe restart of the engine when the main system fails; Energy recovery module: used to generate electricity using waste heat from exhaust gases.

[0007] Preferably, the turbojet engine adopts a micro axial flow design, is integrated at the end of the rotor arm of a multi-rotor unmanned aerial vehicle, and adjusts the thrust direction through a vector nozzle. The nozzle is made of a high-temperature resistant alloy material and is equipped with an electric servo mechanism, which can deflect within ±30° to provide pitch, roll, and yaw control torques. At the same time, the compressor of the turbojet engine adopts a single-stage centrifugal design, the combustion chamber adopts an annular evaporation tube fuel atomization technology, and the turbine part adopts ceramic matrix composite material blades.

[0008] Preferably, the fuel supply module includes a high-pressure fuel pump, a micro fuel filter, and an adaptive flow regulating valve to ensure stable fuel delivery and adapt to different flight conditions. The high-pressure fuel pump adopts a gear structure to provide a constant fuel pressure. The micro fuel filter is internally provided with multiple layers of sintered metal meshes to intercept tiny particulate impurities. The adaptive flow regulating valve adjusts the opening degree in real time through an electronic control unit. The fuel pipeline is made of lightweight titanium alloy material and is wrapped with a heat insulation layer.

[0009] Preferably, the electronic control unit uses multi-sensor fusion technology to monitor the turbojet speed, temperature, and thrust parameters in real time, and cooperates with the flight control system to achieve dynamic power distribution. The speed monitoring uses a Hall effect sensor, the temperature monitoring is covered by a distributed thermocouple array for the combustion chamber and the turbine section, and the thrust parameters are fed back in real time by a micro piezoelectric force sensor. The electronic control unit is built-in with an adaptive PID algorithm and communicates with the main flight control through the CAN bus.

[0010] Preferably, the air intake guiding structure adopts a combined design of an annular diffuser and an anti-vortex grille. The annular diffuser is provided with a gradually expanding cross-section. The anti-vortex grille is composed of titanium alloy thin sheets arranged radially. A guiding lip is provided at the edge of the air intake, and the surface of the overall structure is coated with an oil-repellent coating.

[0011] Preferably, the exhaust guiding structure includes a high-temperature resistant ceramic coating and guiding fins. The ceramic coating is a zirconia-based composite material, which can withstand temperatures above 1000°C. The guiding fins are distributed in a spiral shape. The end of the nozzle is designed as a convergent-divergent configuration, and micro air film cooling holes are integrated on the outer wall of the exhaust duct.

[0012] Preferably, the cooling module adopts an air-cooling and oil-cooling double circulation circuit, and strengthens heat dissipation through the downwash airflow of the rotor. The air-cooling circuit uses the shunted air from the compressor to impact-cool the turbine blades. The oil-cooling circuit exchanges heat through an aluminum alloy micro radiator and the downwash airflow of the rotor. The cooling oil uses a synthetic ester high-temperature oil product and is equipped with a magnetic filter, and phase change materials are embedded in the temperature-sensitive areas.

[0013] Preferably, the shock-absorbing bracket is composed of a carbon fiber composite material and an elastic damper to suppress the transmission of turbojet vibration to the UAV body. The carbon fiber composite material is provided with an orthogonal ply structure, the elastic damper is a silicone-metal laminated structure, and an inertial balance weight is provided at the connection between the bracket and the rotor arm.

[0014] Preferably, the quick-disassembly interface realizes the second-level replacement of the turbojet power unit through a standardized buckle and an electro-hydraulic plug, supporting rapid battlefield maintenance. The buckle is set as a shape memory alloy locking mechanism, the electro-hydraulic plug has a self-sealing function, the interface positioning pin is set as a magnetic-assisted alignment structure, and the supporting tool integrates an infrared scanning module.

[0015] Preferably, the redundant backup system includes a dual ignition device and an emergency battery pack to ensure the safe restart of the turbojet in case of main power failure. The dual ignition device is provided with independent spark plugs and a capacitor discharge unit for providing redundant ignition sources. The emergency battery pack is a supercapacitor-lithium battery hybrid power supply that can switch power supply within 200 ms. The system is built-in with a self-checking circuit to monitor the ignition state in real time. When an abnormality occurs, it automatically triggers the standby ignition sequence and synchronously closes the fuel valve.

[0016] Advantages of the present invention: 1. By replacing the traditional electric motor with a micro turbojet engine, the power performance of the multi-rotor UAV is improved. The turbojet engine has a higher energy density and thrust-to-weight ratio, enabling the UAV to obtain a longer endurance time when carrying the same load. Especially in extreme environments such as high altitude or high temperature, it can still maintain stable power output, overcoming the problem of power attenuation of the electric motor due to thin air or poor heat dissipation. At the same time, the high-speed exhaust characteristic of the turbojet engine provides the UAV with stronger maneuverability, enabling it to perform more complex flight maneuvers and faster speed responses.

[0017] 2. Through the deflectable nozzle design, the adjustment of the thrust direction in multiple degrees of freedom is realized. This design not only provides the direct thrust vector control ability that traditional multi-rotor UAVs do not have, but also greatly improves flight stability, enabling the UAV to maintain precise flight attitude control even in harsh environmental conditions such as strong winds. At the same time, the deep integration of the vector thrust system and the flight control enables the UAV to perform more complex maneuvering actions, providing more possibilities for special operations and military applications.

[0018] 3. Adopting modular design, the standardized quick-disassembly and assembly interface enables the turbojet power unit to be replaced within seconds under battlefield or field conditions, greatly enhancing the attendance rate and mission reliability of the UAV. This design not only reduces the maintenance difficulty and cost but also supports the quick replacement of different specifications of power modules according to different mission requirements, providing great convenience for the multi-functional application of the UAV. Meanwhile, the integrated self-check and positioning system ensures the accuracy and safety of the replacement process.

[0019] 4. The design of the dual ignition device and emergency power supply ensures that the engine can still be safely restarted or smoothly stopped even in the case of main system failure. This redundant design reduces the risk of complete failure of the power system. Meanwhile, the intelligent self-check and switching mechanism realizes the instant detection and automatic recovery of faults, thus enhancing the fault tolerance of the entire power system. Description of the Drawings

[0020] Figure 1 Shows the schematic diagram of the specific implementation process of the present invention; Figure 2 Shows the schematic diagram of the emergency handling process of the present invention. Detailed Embodiment

[0021] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are some but not all of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the scope of protection of the present invention.

[0022] The present invention provides an embodiment: a turbojet power system for a multi-rotor UAV, including: Turbojet engine: used to provide the main lift and attitude control ability for the UAV; Fuel supply module: used to stably supply fuel to the turbojet engine; Electronic control unit: used to monitor and adjust the operating state of the turbojet engine in real time; Air intake diversion structure: used to optimize the air intake efficiency and prevent foreign objects from entering; Exhaust diversion structure: used to improve the exhaust thrust efficiency and reduce the infrared signature; Cooling module: used to prevent the engine from overheating; Shock-absorbing bracket: used to suppress the vibration conduction of the turbojet engine; Quick-disassembly and assembly interface: used to realize the quick replacement of the turbojet power unit; Redundant backup system: used to ensure the safe restart of the engine when the main system fails; Energy recovery module: used for generating electricity by utilizing the waste heat of exhaust gas.

[0023] Furthermore, the turbofan engine of the present invention adopts a micro axial-flow design and is integrated at the end of the rotor arm of a multi-rotor UAV. Specifically, the compressor of the engine adopts a single-stage centrifugal structure and is made of high-strength titanium alloy to improve air compression efficiency and reduce weight. The combustion chamber adopts an annular evaporation tube type fuel atomization technology to ensure full mixing of fuel and air and efficient combustion. The turbine part adopts ceramic matrix composite (CMC) blades, and the temperature resistance can reach over 1200 °C, while reducing the influence of thermal expansion on performance. The nozzle adopts a high-temperature resistant nickel-based alloy and is equipped with an electric servo mechanism, which can perform vector adjustment within the range of ±30° to provide pitching, rolling and yaw control torques.

[0024] The combustion chamber adopts an annular evaporation tube type fuel atomization technology, and its core structure consists of multiple evaporation tubes evenly distributed circumferentially. Each evaporation tube is internally provided with a spiral guide vane to enhance the fuel atomization effect. The fuel is sprayed into the inner wall of the evaporation tube through a high-pressure nozzle with a particle size of 15 - 20 μm to form an oil film. At the same time, the 300 - 400 °C high-temperature air from the compressor enters the evaporation tube tangentially, and a strong swirl is formed under the action of the spiral guide vane, causing the oil film to evaporate on the tube wall and fully mix with the air. The inner surface of the evaporation tube adopts a porous sintered metal material to increase the heat transfer area and promote secondary fragmentation of fuel particles. The mixed combustible gas enters the main combustion zone with a uniform velocity through the porous injection plate at the end of the evaporation tube. A pilot flame stabilizer is provided at the head of the combustion chamber to ensure ignition reliability.

[0025] Furthermore, the fuel supply module includes a high-pressure gear fuel pump, a multi-layer metal sintered mesh micro filter and an electronic control flow regulating valve. The fuel pump is driven by a motor to provide a stable fuel pressure (typical value 3 - 5 MPa). The fuel filter adopts a filtration accuracy of 5 μm level to prevent tiny particles from entering the combustion chamber and causing wear. The flow regulating valve is controlled by the ECU and dynamically adjusts the fuel flow according to the real-time working conditions of the engine to ensure maximum combustion efficiency. The fuel pipeline adopts titanium alloy material and is covered with a silicon-based heat insulation layer to prevent fuel vaporization in a high-temperature environment.

[0026] The flow control valve dynamically adjusts the spool opening based on real-time data from the engine speed sensor, intake air temperature sensor, combustion chamber pressure sensor, and exhaust gas temperature sensor by an in-built adaptive fuel control algorithm in the electronic control unit (ECU). This algorithm first calculates the theoretical fuel demand according to the current flight altitude and airspeed, and then performs closed-loop correction in combination with the oxygen content in the exhaust gas feedback by the oxygen sensor. When an acceleration command is detected, feed-forward control is adopted to increase the fuel supply in advance to prevent turbo lag. If in a steady-state condition, the air-fuel ratio is maintained near the optimal 14.7:1 through PID regulation. At the same time, the ECU monitors the combustion chamber pressure fluctuation frequency and fine-tunes the spool position through high-frequency pulse width modulation (PWM) signals to suppress combustion oscillation.

[0027] Furthermore, the ECU monitors the engine state in real time through multiple sensors, including Hall effect speed sensors, thermocouple temperature sensors, and piezoelectric thrust sensors. The control algorithm adopts adaptive PID regulation to optimize the engine response speed. The ECU communicates with the main flight control of the UAV through the CAN bus, receives flight commands, and coordinates the thrust distribution of multiple turbojet engines to achieve precise flight control.

[0028] The PID algorithm establishes a multi-parameter closed-loop control model of the engine operating conditions by collecting the speed signal provided by the Hall effect sensor, the data of the thermocouple temperature sensor, and the feedback of the piezoelectric thrust sensor in real time. Among them, the proportional link dynamically adjusts the gain coefficient according to the instantaneous deviation between the set speed and the actual speed to quickly eliminate the error. The integral link automatically compensates the steady-state error of the system by accumulating the historical deviation. The derivative link predicts the speed change trend and applies the control quantity in advance to suppress overshoot. The ECU conducts data interaction with the main flight control of the UAV through a high-speed CAN bus (transmission rate 1 Mbps), receives the pitch, roll, and throttle commands sent by the flight control in the form of a timing interruption with a period of 10 ms, and converts them into the target thrust parameters of each turbojet engine. At the same time, based on the current attitude information of the UAV and the aerodynamic characteristic model, the weighted least squares method is used to calculate the optimal thrust distribution scheme, coordinate the vector nozzle deflection angles (±30° adjustable) and fuel flow rates (regulation accuracy ±0.5%) of multiple turbojet engines, and achieve the three-axis moment balance of the UAV through the differential control of the thrust vector on the premise of ensuring that the total lift meets the flight requirements. Among them, the roll control is achieved by the thrust difference between the left and right engines, the pitch control is adjusted by the thrust difference between the front and rear engines, and the yaw control relies on the lateral component of the vector nozzle to generate a yaw moment, ultimately enabling the UAV to maintain an attitude control accuracy of ±0.5° and a speed tracking error of 0.1 m / s in various flight states.

[0029] Furthermore, during the intake process, the high-speed incoming air is first decelerated and pressurized through the divergent section of the annular diffuser. The anti-vortex titanium alloy grille installed at the diffuser inlet is arranged radially at 15° to eliminate turbulence and improve the airflow uniformity. The intake lip is designed with the NACA series airfoil profile to minimize the boundary layer separation loss. The inner surface of the entire intake passage is coated with a fluoropolymer oil-repellent coating to prevent pollutant adhesion. During the exhaust process, the spiral guide fins are used to make the high-temperature exhaust gas generate swirl and accelerate the exhaust. The inner wall of the exhaust duct is coated with a plasma-sprayed 8% yttria-stabilized zirconia (8YSZ) ceramic coating to withstand the high-temperature environment of 1000°C continuously. The nozzle end is designed as a convergent-divergent Laval nozzle structure, and the ratio of its throat cross-sectional area to the outlet area is 1:1.8 to optimize the supersonic exhaust expansion efficiency. The outer wall of the exhaust duct is arranged with a microporous array with a diameter of 0.5 mm. By introducing the bypass air from the third stage of the compressor to form a gas film cooling layer, the outer wall temperature is reduced to below 300°C. At the same time, a flexible graphite sealing ring is set at the connection between the exhaust system and the airframe to compensate for the thermal expansion deformation. The overall intake and exhaust system makes the total pressure loss controlled within 5% through the optimized flow path profile design by computational fluid dynamics (CFD).

[0030] Furthermore, the cooling system adopts a dual-cycle design: Air-cooling circuit: 5%-10% of the air is diverted from the compressor and impacts the turbine disk through the guide vanes for cooling; Oil-cooling circuit: The synthetic ester cooling oil is cooled by the aluminum alloy radiator, and the downwash airflow of the rotor is utilized to enhance the heat transfer. Phase change materials (such as paraffin-based composites) are embedded in the key high-temperature areas (such as the roots of the turbine blades) to absorb the transient thermal shock.

[0031] Furthermore, the shock-absorbing bracket is composed of carbon fiber composite materials and elastic dampers to inhibit the transmission of the turbojet vibration to the UAV airframe. The carbon fiber composite materials are provided with an orthogonal ply structure, the elastic damper is a silicone-metal laminated structure, and an inertial balance block is set at the connection between the bracket and the rotor arm.

[0032] Furthermore, the quick-disassembly interface realizes the second-level replacement of the turbojet power unit through standardized buckles and electro-hydraulic plugs, supporting rapid battlefield maintenance. The buckle is set as a shape memory alloy locking mechanism, the electro-hydraulic plug has a self-sealing function, the interface positioning pin is set as a magnetic-assisted alignment structure, and the supporting tool integrates an infrared scanning module.

[0033] Furthermore, the redundant backup system includes a dual ignition device and an emergency battery pack to ensure the safe restart of the turbojet in case of main power failure. The dual ignition device is provided with independent spark plugs and a capacitor discharge unit to provide redundant ignition sources. The emergency battery pack is a supercapacitor-lithium battery hybrid power supply that can switch power supply within 200 ms. The system is built-in with a self-checking circuit to monitor the ignition status in real time. When an abnormality occurs, it automatically triggers the standby ignition sequence and synchronously closes the fuel valve.

[0034] Please refer to Figure 1 and Figure 2 , furthermore, the specific implementation manners of the present invention will be described: When the drone starts, the electronic control unit (ECU) sends an instruction to the fuel pump to establish fuel pressure, and at the same time activates the capacitor discharge unit of the dual ignition system. The spark plugs ignite the atomized fuel in the combustion chamber, causing the turbojet engine to enter the idle state. At this time, the rotational speed is maintained within the lowest stable operating range, and the exhaust gas temperature sensor monitors the combustion state in real time to ensure a smooth start process.

[0035] When the drone receives the takeoff instruction, the ECU gradually increases the rotational speed of the turbojet engine according to the required lift calculated by the flight control system. The fuel flow regulating valve synchronously increases the opening degree to match the power demand. The vector nozzle remains vertically downward in the initial stage to provide the maximum lift. At the same time, the air-cooling and oil-cooling circuits of the cooling module start to operate at full load. The phase change material absorbs the transient high temperature in the turbine area to prevent the accumulation of thermal stress. The rotor downwash airflow is actively guided to the radiator surface to enhance the heat transfer efficiency.

[0036] In the climbing stage, the flight control system dynamically adjusts the thrust distribution of each turbojet engine according to the attitude sensor data. The vector nozzle makes a slight deflection (within the range of ±5°) according to the flight attitude change to compensate for the rotor torque and external wind disturbance. The anti-vortex grille of the air intake guiding structure ensures that the compressor always obtains a stable air flow to avoid surging. The exhaust guiding fins force the exhaust gas to spiral out to improve the thrust efficiency. At the same time, the thermoelectric generator starts to use the waste heat of the exhaust gas to supply power to the avionics equipment.

[0037] After entering the cruise flight, the rotational speed of the turbojet engine stabilizes within the optimal fuel efficiency range. The vector nozzle turns to the horizontal direction to provide the forward flight thrust. At this time, the flight control system realizes efficient level flight by coordinating the rotational speed of the rotor and the angle of the nozzle. The ECU continuously optimizes the fuel injection volume to keep the temperature in the combustion chamber within the safe range. The cooling module dynamically adjusts the air-cooling diversion ratio and the oil-cooling circulation rate according to the real-time temperature data to ensure thermal balance.

[0038] When the UAV performs high-speed maneuvers, the ECU instantaneously increases the thrust of the turbofan engine in the corresponding direction. The vector nozzle completes the maximum deflection (±30°) within 200 ms under the drive of the servo motor, directly generating pitching, rolling or yawing moments. The flight control system synchronously reduces the rotational speed of the rotor to reduce aerodynamic interference. The carbon fiber-damping silicone composite structure of the shock-absorbing bracket effectively absorbs high-frequency vibrations, thereby preventing the resonance of the airframe structure. The redundant backup system continuously monitors the ignition state and immediately switches to the standby ignition unit once an anomaly is detected.

[0039] During the landing phase, the ECU gradually reduces the turbofan speed. The vector nozzle returns vertically downward to provide buffer lift. The fuel supply module enters the low-flow mode to reduce carbon deposition in the combustion chamber. The cooling system keeps running until the engine temperature drops to the safety threshold. The magnetic-assisted positioning function of the quick-disassembly interface is ready, facilitating rapid maintenance or replacement of the power unit after landing.

[0040] In case of an emergency (such as the failure of the main power supply), the supercapacitor-lithium battery hybrid power supply of the redundant system immediately takes over the power supply. The dual ignition devices ensure that the engine does not stall. If the fault persists, the ECU automatically closes the fuel valve and initiates the safe shutdown procedure. The ceramic coating of the exhaust diversion structure delays heat radiation to avoid scalding ground personnel.

[0041] After the mission is completed, the maintenance personnel press the release button of the quick-disassembly interface. The shape memory alloy buckle is heated and unlocked. The electro-hydraulic plug cuts off the oil circuit from the self-sealing valve. The entire turbofan power unit can be disassembled within 30 seconds. After replacing the new unit, the infrared scanner automatically calibrates the installation accuracy, thus ensuring the reliable execution of the next mission.

Claims

1. Turbojet power system for multi-rotor UAV, characterized in that, Includes: Turbojet engine: used to provide the main lift and attitude control capabilities for the UAV; Fuel supply module: used to stably deliver fuel to the turbojet engine; Electronic control unit: used to monitor and adjust the operating status of the turbojet engine in real time; Air intake guide structure: used to optimize air intake efficiency and prevent foreign matter from entering; Exhaust guide structure: used to improve exhaust thrust efficiency and reduce infrared characteristics; Cooling module: used to prevent the engine from overheating; Shock absorber bracket: used to suppress the vibration transmission of turbojet engine; Quick disassembly interface: used to realize the quick replacement of turbojet power unit; Redundant backup system: used to ensure safe restart of the engine when the main system fails; Energy recovery module: used to generate electricity using waste heat from exhaust gas.

2. The turbojet power system of the multi-rotor unmanned aerial vehicle according to claim 1, wherein: The turbojet engine adopts a micro axial flow design and is integrated into the end of the rotor arm of a multi-rotor UAV. The thrust direction is adjusted through a vector nozzle. The nozzle is made of high-temperature resistant alloy material and is equipped with an electric servo mechanism. It can be deflected within a range of ±30° to provide pitch, roll and yaw control torque. At the same time, the compressor of the turbojet engine adopts a single-stage centrifugal design, the combustion chamber adopts annular evaporator tube fuel atomization technology, and the turbine part adopts ceramic-based composite blades.

3. The turbojet power system of the multi-rotor unmanned aerial vehicle according to claim 1, characterized in that: The fuel supply module includes a high-pressure fuel pump, a micro fuel filter and an adaptive flow control valve to ensure stable fuel delivery and adapt to different flight conditions. The high-pressure fuel pump adopts a gear structure to provide constant fuel pressure. The micro fuel filter has a built-in multi-layer metal sintered mesh to intercept tiny particle impurities. The adaptive flow control valve adjusts the opening in real time through the electronic control unit. The fuel pipeline is made of lightweight titanium alloy and wrapped with an insulation layer.

4. The turbojet power system of the multi-rotor unmanned aerial vehicle according to claim 1, characterized in that: The electronic control unit uses multi-sensor fusion technology to monitor the turbojet speed, temperature and thrust parameters in real time, and cooperates with the flight control system to achieve dynamic power distribution. The speed monitoring uses a Hall effect sensor, and the temperature monitoring covers the combustion chamber and turbine section through a distributed thermocouple array. The thrust parameters are fed back in real time by a miniature piezoelectric force sensor. The electronic control unit has a built-in adaptive PID algorithm and communicates with the main flight control through the CAN bus.

5. The turbojet power system of the multi-rotor unmanned aerial vehicle according to claim 1, characterized in that: The air intake guide structure adopts a combined design of an annular diffuser and an anti-vortex grille, wherein the annular diffuser is provided with a gradually expanding cross-section, the anti-vortex grille is composed of titanium alloy sheets arranged radially, a guide lip is provided at the edge of the air intake, and the surface of the entire structure is coated with an oleophobic coating.

6. The turbojet power system of the multi-rotor unmanned aerial vehicle according to claim 1, characterized in that: The exhaust guide structure includes a high-temperature resistant ceramic coating and guide fins, wherein the ceramic coating is a zirconium oxide-based composite material that can withstand high temperatures above 1000°C. The guide fins are distributed in a spiral shape, the nozzle end is designed as a convergent-divergent configuration, and the outer wall of the exhaust duct is integrated with micro air film cooling holes.

7. The turbojet power system of the multi-rotor unmanned aerial vehicle according to claim 1, characterized in that: The cooling module adopts an air-cooling and oil-cooling dual-circulation circuit, and enhances heat dissipation through the rotor downwash airflow. The air cooling circuit uses the compressor diverted air to impact cool the turbine blades, and the oil cooling circuit exchanges heat with the rotor downwash airflow through an aluminum alloy micro-radiator. The cooling oil uses synthetic ester high-temperature oil and is equipped with a magnetic filter. Phase change material is embedded in the temperature-sensitive area.

8. The turbojet power system of the multi-rotor unmanned aerial vehicle according to claim 1, characterized in that: The shock-absorbing bracket is composed of carbon fiber composite material and an elastic damper to inhibit the transmission of turbojet vibration to the UAV airframe. The carbon fiber composite material is provided with an orthogonal ply structure, the elastic damper is a silicone-metal laminated structure, and an inertial balance block is provided at the connection between the bracket and the rotor arm.

9. The turbojet power system of the multi-rotor unmanned aerial vehicle according to claim 1, wherein: The quick-disassembly interface realizes the second-level replacement of the turbojet power unit through a standardized buckle and an electro-hydraulic plug, supporting rapid battlefield maintenance. The buckle is set as a shape memory alloy locking mechanism, the electro-hydraulic plug has a self-sealing function, the interface positioning pin is set as a magnetic-assisted alignment structure, and the supporting tool integrates an infrared scanning module.

10. The turbojet power system of the multi-rotor unmanned aerial vehicle according to claim 1, characterized in that: The redundant backup system includes a dual ignition device and an emergency battery pack to ensure the safe restart of the turbojet in case of main power failure. The dual ignition device is provided with independent spark plugs and a capacitor discharge unit to provide redundant ignition sources. The emergency battery pack is a supercapacitor-lithium battery hybrid power supply that can switch power supply within 200 ms. The system is built-in with a self-checking circuit to monitor the ignition state in real time. When an abnormality occurs, it automatically triggers the standby ignition sequence and synchronously closes the fuel valve.

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