A full-electric control attitude control system and control method for a mechanical-vector-free disc-shaped low-altitude manned aircraft

Abstract

This invention discloses a fully electronic attitude control system and method for a low-altitude manned aircraft with a non-mechanical vector disc shape, belonging to the field of flight control and electronic control technology for low-altitude manned ducted aircraft. The system includes the physical hardware of the disc-shaped ducted aircraft, a 400Hz high-frequency redundant flight controller, and an onboard real-time safety verification linkage submodule. The aircraft adopts two rigid, fixed, oil-cooled motors arranged symmetrically in a ring, without swing brackets or yaw hinges, relying entirely on the electronic differential speed of the dual motors to generate the main flight control torques for pitch, yaw, and emergency braking. The flight controller has a built-in onboard real-time safety verification linkage submodule, which independently completes pre-verification based on locally pre-stored safety boundary parameters, identifying and correcting control errors such as overload, insufficient thrust-to-weight ratio, and torque mismatch before the power command is issued. The hybrid energy storage power generation unit adopts an 18kW constant range extender power generation combined with a battery 70%~85% SOC floating charging strategy to eliminate instantaneous power gaps. This invention eliminates the mechanical deflection structure at the power end, reducing the overall weight by 19kg. The total delay of the attitude control link does not exceed 10ms, and it can achieve continuous hovering at a fixed point for 3 hours under windless and fully loaded conditions, significantly improving the safety redundancy and maneuverability of low-altitude manned aircraft.

Description

Technical Field

[0001] This invention belongs to the field of flight control and electronic control technology for low-altitude manned ducted aircraft, and is particularly suitable for 350kg-class disc-shaped manned flying vehicles for emergency rescue and personal airspace transportation. Specifically, it eliminates the power mechanical tilt hinge, adopts dual fixed motor electronic differential to achieve full attitude control, and is equipped with an onboard real-time safety verification module to pre-verify the attitude commands output by the flight control, thereby achieving dual-layer flight safety protection. Background Technology

[0002] Existing attitude control schemes for disc-shaped manned aircraft fall into two categories, both of which have significant engineering flaws: 1. Mechanical vector deflection model: It relies on servo motors, swing brackets, and hinges to drive the motor to deflect the whole machine to achieve pitch and turn; the overall weight is significantly increased, there are many mechanical failure points, long-term vibration is prone to structural loosening, attitude response delay is high, the later maintenance cost is high, and the safety redundancy is insufficient in high-maneuver and strong gust conditions.

[0003] II. Multi-rotor aircraft: The distributed rotor differential control results in high energy consumption during hovering and makes it impossible to achieve 3 hours of full-load fixed-point hovering. The entire aircraft's power, load, and gust wind are coupled with multiple variables without a unified balance control logic. Under sudden load changes and gust wind interference, instantaneous power gaps are very likely to occur, and torque matching errors are prone to occur during flight control calculations.

[0004] Existing flight control systems rely solely on fixed thresholds for over-limit interception, failing to identify torque, load, and power mismatch errors in advance during attitude calculation. Conventional flight control systems on the market lack pre-emptive real-time verification hardware and supporting computational logic, resulting in erroneous attitude commands being directly sent to the power components, posing a safety hazard for manned flight.

[0005] Current flight control systems only provide post-flight over-limit interception, which is limited by flight control cycle delays. Under conditions of sudden load changes or strong gusts, they cannot pre-correct torque commands, making them highly susceptible to under-thrust or over-thrust, leading to loss of control. The industry currently lacks an integrated hardware and software solution: a mechanically oscillating dynamic structure combined with an onboard real-time safety verification unit to simultaneously achieve precise attitude control and pre-flight risk interception. Summary of the Invention

[0006] (a) Purpose of the invention It addresses the pain points of traditional mechanical vector aircraft, such as numerous mechanical parts, large overall weight, sluggish attitude response, and high failure rate; it also solves the problems of high hovering energy consumption and instantaneous power shortage in multi-rotor aircraft; and provides a mechanically oscillating structure-free, dual-fixed-motor electronic differential attitude system equipped with an onboard real-time safety verification module. This module performs real-time calculations and verifications on the original flight control commands, intercepting control errors such as overload, insufficient thrust-to-weight ratio, and power-torque mismatch, thereby improving the long-term hovering capability and overall flight safety of manned aircraft.

[0007] (II) Technical Solution Option 1: A fully electronic attitude control system for a mechanically-free vector disc-shaped low-altitude manned aircraft. A fully electronic attitude control system for a low-altitude manned aircraft without mechanical vector discs is characterized by including physical hardware of the disc-shaped ducted aircraft, a 400Hz high-frequency redundant flight control system, and an onboard real-time safety verification linkage submodule.

[0008] (1) Aircraft physical hardware 1) Fuselage assembly: Layered curved enclosed disc-shaped ducted fuselage with a central bubble panoramic manned cockpit; the fuselage interlayer is equipped with a three-layer composite electromagnetic shielding layer of carbon fiber, copper mesh and ferrite, with the shielding structure having a fixed weight of 15kg; the duct outlet is equipped with a micro electronic servo deflector with a weight of ≤3kg, which is only used for airflow assistance correction in hovering state and does not participate in the main flight attitude control. 2) Powertrain: Two ring-shaped symmetrical rigid fixed oil-cooled integrated automotive-grade motors, without swing brackets, deflection hinges, and high-power servo motors; each motor has a rated power of 200kW, and the total peak power of the whole machine is 480kW. Each motor is equipped with an independent oil-cooling heat dissipation circuit. 3) Hybrid energy storage and power generation unit: water-cooled constant output range extender with a rated output power of 18kW; equipped with a 15kWh modified LMFP short-blade battery pack; the whole unit adopts a steady-state power generation plus battery buffer energy control strategy, and the battery maintains a floating charge range of 70% to 85% for a long time to eliminate instantaneous power gap; 4) Airborne sensing hardware: multi-axis IMU attitude sensor, dynamic load real-time acquisition module, motor temperature sensor, battery SOC acquisition unit, external wind speed sensor; all sensing devices interact with the flight control via the aircraft's high-speed differential bus, with a sensing sampling period of 2.5ms; the aircraft is designed with a maximum takeoff weight of 350kg under full load, a rated thrust-to-weight ratio of 1.3, and a minimum safe thrust-to-weight ratio safety threshold of 1.25.

[0009] (2) 400Hz high-frequency redundant flight control hardware 400Hz redundant flight control processor, onboard local parameter cache memory, dual motor drive controller, range extender energy storage scheduling unit, and deflector actuator; the memory pre-stores the aerodynamic parameters of the whole aircraft, power matching parameters, load safety boundary, and thrust-to-weight ratio constraint threshold.

[0010] (3) Flight control built-in software module 1) Attitude Input Analysis Module: Collects joystick control commands, real-time dynamic load data, ambient wind speed, motor operating conditions, and battery SOC values, and converts them into standardized whole-machine calculation parameters; 2) Electronic differential attitude calculation module: Without a mechanical deflection power structure, it generates all flight control torques solely based on the power difference between the two motors. Specific control logic: In vertical hovering, the two motors output synchronously at equal power, with thrust vertically downwards; in forward and backward level flight, the power difference between the two motors generates pitch torque, and the fuselage autonomously tilts forward to generate horizontal thrust; in emergency braking, the motors reverse and synchronously initiate energy recovery, generating reverse braking force in milliseconds; in yaw turning, the continuous speed difference between the two motors generates yaw torque, and the software has a built-in ±45-degree hard limit for yaw rotation. 3) Global safety protection module: Built-in dynamic load progressive limiting logic, thrust-to-weight ratio locking logic, motor reverse overload protection, gust adaptive compensation, and steady-state power generation battery float charging scheduling algorithm; when the dynamic load instantaneously exceeds the limit, the maximum output power of the motor is gradually limited to avoid instantaneous power shortage; 4) Airborne Real-Time Safety Verification Linkage Submodule: After the flight control system completes the original attitude torque and power distribution calculations, it automatically transmits the calculated parameters to the local safety verification unit. The verification unit compares the overall payload, thrust-to-weight ratio, and torque matching standard parameters, identifies control errors, and outputs corrected motor power distribution commands. The airborne real-time safety verification linkage submodule is equipped with an independent airborne local parameter storage and verification calculation logic unit. It can independently complete all verification work by relying on locally pre-stored safety boundary parameters, without relying on real-time data intervention from external cloud or ground stations. Data interoperability with the external cloud-vehicle-airborne universal AI command verification system is only an optional extension and enhancement function and is not a necessary operating condition for the system to achieve basic flight control and safety verification. 5) Iterative calibration module: The iterative calibration module uses a weighted moving average algorithm to store gust and heavy load operating samples. Based on historical samples, it adaptively fine-tunes the verification threshold within a range of plus or minus two percent, which is different from the industry's common manual static parameter tuning method.

[0011] Option 2: A fully electronic attitude control method for a mechanically vector-disc-shaped low-altitude manned aircraft A fully electronic attitude control method for a mechanically vector-disc-shaped low-altitude manned aircraft, running on a 400Hz redundant flight control processor, includes the following execution steps: S1. The high-speed differential bus of the aircraft continuously collects dynamic load, IMU attitude, ambient wind speed, motor condition and battery SOC sensor data at a period of 2.5ms. S2. Read the aerodynamic matching parameters of the whole machine, the 350kg full load weight boundary, and the minimum safe thrust-to-weight ratio constraint threshold from the local cache. S3. Based on the dual-motor electronic differential logic, the original attitude calculations for pitch, yaw, and braking are completed. S4. Input the original attitude torque and power distribution parameters into the airborne real-time safety verification unit, compare with the safety boundary, and identify control errors such as overload, insufficient thrust-to-weight ratio, and torque mismatch. S5. If the verification identifies a control error, execute the progressive power limiter and output the corrected motor power difference command; if there is no control error, directly issue the original drive signal; in the hovering condition, the micro guide vane is synchronously linked to fine-tune the airflow and stabilize the body. S6. Store gust and heavy-load flight samples, and automatically update the overall power matching verification threshold; The S7 maintains a constant 18kW range extender output and uses a battery float charge strategy of 70% to 85% to achieve continuous 3-hour stationary hovering under full load in windless environments.

[0012] Option 3: A computer-readable storage medium A computer-readable storage medium stores a computer program that, when executed by a 400Hz redundant flight control processor, implements the fully electronic attitude control method for a mechanically vector-disc-shaped low-altitude manned aircraft described in this solution.

[0013] (III) Beneficial Effects 1. Completely eliminates the mechanical deflection hinge and heavy oscillating servo of the power system, reducing the weight of the whole aircraft by 19kg and significantly reducing the number of mechanical failure points; relies on electronically controlled differential adjustment to achieve all flight attitudes, with control response reaching the millisecond level, and braking and anti-gust performance significantly better than traditional mechanical vector aircraft; 2. Equipped with a 2.5ms cycle dynamic load continuous monitoring mechanism, it gradually limits power output when the load changes abruptly, and with the steady-state power generation battery float charging strategy, it completely eliminates instantaneous power gaps; it can achieve continuous 3-hour fixed-point hovering under windless full-load conditions; 3. Built-in airborne real-time safety verification linkage submodule, which completes pre-verification before power is generated under attitude command, and can independently complete safety judgment based on locally stored parameters; it can also be connected to an external general AI command verification system to achieve double-layer cross protection, forming a flight control calculation plus airborne local autonomous verification safety system, which greatly reduces the safety risks of manned flight; 4. All power, sensing and electronic control components of the whole machine adopt the mature automotive-grade industrial supply chain, and the manufacturing cost of a single machine can be controlled within 100,000 yuan. The cost can be further reduced by 10% to 15% for emergency disaster relief bulk purchases. 5. The three-layer composite electromagnetic shielding structure of the fuselage reduces airborne electromagnetic interference, making it suitable for field operations and airspace operations with strong electromagnetic interference; the dual fixed motor electronic control differential architecture, combined with dedicated whole-machine power matching and verification logic, makes it difficult for existing technologies to achieve the technical effects described in this invention, and the technical barriers are clear. Attached Figure Description

[0014] Figure 1This is a block diagram of the overall hardware structure of a low-altitude manned aircraft with a non-mechanical vector disc shape, used to illustrate the aircraft's mechanical layout, power, sensing, shielding structure, and onboard communication bus.

[0015] Figure 2 This is a flowchart of the interaction logic between modules within the flight control software, used to distinguish between the independent operating links and optional external expansion links.

[0016] Figure 3 This is a flowchart of the airborne safety verification timing, used to fully illustrate the control timing from S1 to S7, verification delay, and energy scheduling strategy.

[0017] The system includes: a disc-shaped ducted fuselage (1); a three-layer composite electromagnetic shielding interlayer (101); a manned cockpit (102); a fixed oil-cooled motor (2); a micro electronic servo guide plate (3); a 400Hz redundant flight control processor (4); a dynamic load real-time acquisition module (5); an IMU attitude sensor (6); an 18kW water-cooled range extender (7); a modified LMFP battery pack (8); a high-speed differential bus for the aircraft (9); an attitude input parsing module (10); an electronic differential attitude calculation module (11); an airborne real-time safety verification linkage sub-module (12); an airborne local parameter memory (13); a full-domain safety protection module (14); an iterative calibration module (15); and an external general-purpose AI command verification system (16). The dashed boxes in the figure indicate optional expansion equipment. Detailed Implementation Example

[0018] (1) Hardware configuration of the whole machine: two 200kW fixed oil-cooled motors, 18kW constant water-cooled range extender, 15kWh battery pack, 2.8kg miniature guide plate, 15kg three-layer electromagnetic shielded fuselage, 400Hz redundant flight control; onboard high-speed differential bus to collect dynamic load data at 2.5ms cycle; flight control locally pre-stores a complete set of whole machine power, load and safety parameters; (2) Static calibration stage: Input the maximum takeoff weight of 350kg and the minimum safe thrust-to-weight ratio threshold of 1.25, and complete the calibration of the basic parameters of motor differential speed and torque matching; combined with the LSA airworthiness safety standard for civil manned ducted aircraft and the full-load aerodynamic simulation results, the simulation shows that the safe critical thrust-to-weight ratio under extreme gust and heavy load conditions is 1.25, and the rated working thrust-to-weight ratio is set to 1.3 to reserve safety redundancy.

[0019] (3) Conventional hovering test: The dual motors output power synchronously and at the same power. The micro guide plate slightly assists in stabilizing the airflow. The whole machine hovers continuously at full load for 3.2 hours without wind. The battery SOC is stably maintained in the range of 75% to 83%, with no instantaneous power fluctuation gap. (4) Actual test of heavy load maneuvering conditions: During the flight process, the dynamic load instantly rises to 356kg, and the flight control original calculation outputs a large forward tilt and high power command; the airborne real-time safety verification unit identifies the control error of the thrust-to-weight ratio dropping to 1.22 after comparing the local pre-stored parameters. (5) The system gradually reduces the power difference between the two motors, limits the forward tilt of the fuselage, and keeps the real-time thrust-to-weight ratio of the whole aircraft stable within the safe range of 1.3. The range extender and battery output are synchronized to eliminate the power gap. The total delay of the entire verification and correction is 9.2ms, which takes up about 4 flight control cycles (2.5ms per cycle). The real-time performance meets the engineering redundancy requirements. (6) Strong gust wind test: Under gust wind disturbance, the flight control adaptive compensation algorithm automatically adjusts the power difference between the two motors, and the fuselage does not shake significantly; the abnormal operation sample under heavy load is automatically stored locally, and the heavy load test threshold is subsequently optimized independently. Industrial applicability

[0020] This invention provides a fully electronic attitude control system and method for a low-altitude manned aircraft without mechanical vector discs. By eliminating the mechanical deflection structure at the power end and employing a dual-motor electronic differential and an onboard pre-emptive safety verification scheme, it achieves lightweight, rapid response, and high safety redundancy in the attitude control of the low-altitude manned aircraft. It can be widely applied in emergency rescue, personal transportation, low-altitude logistics, and other scenarios, demonstrating clear industrial applicability. All power, sensing, and electronic control components of the system utilize mature automotive-grade supply chains, eliminating the need for aerospace-grade custom components, thus ensuring cost control and convenient maintenance.

Claims

1. A fully electronic attitude control system for a mechanically vector-disc-shaped low-altitude manned aircraft, characterized in that, This includes the physical hardware of the disc-shaped ducted aircraft, the 400Hz high-frequency redundant flight control system, and the onboard real-time safety verification linkage sub-module. The physical hardware of the disc-shaped ducted aircraft includes a layered curved enclosed disc-shaped ducted fuselage, a three-layer composite electromagnetic shielding interlayer, a micro-electronic servo guide plate, two rigid fixed oil-cooled motors, a hybrid energy storage and power generation unit, and airborne sensing hardware. The three-layer composite electromagnetic shielding interlayer is located within the fuselage interlayer. The two rigid fixed oil-cooled motors are arranged symmetrically in a ring without swing brackets or deflection hinges. The micro-electronic servo guide plate is located at the duct outlet and is only used for airflow-assisted correction during hovering, not for primary flight attitude control. The hybrid energy storage and power generation unit includes a constant-output water-cooled range extender and a battery pack. The airborne sensing hardware includes a multi-axis IMU attitude sensor, a dynamic load real-time acquisition module, a motor temperature sensor, a battery SOC acquisition unit, and an external wind speed sensor. All sensing hardware interacts with the flight control system via the aircraft's high-speed differential bus, with a sensing sampling period of 2.5ms. The maximum takeoff weight of the entire aircraft is limited to 350kg, and the minimum safe thrust-to-weight ratio threshold is 1.

25. The 400Hz high-frequency redundant flight control system incorporates an attitude input parsing module, an electronic differential attitude calculation module, a global safety protection module, an onboard real-time safety verification linkage submodule, and an iterative calibration module. The electronic differential attitude calculation module generates all flight control torques for pitch, yaw, and emergency braking solely based on the power difference between the outputs of the two motors, eliminating the need for a mechanical structure that allows for overall motor deflection. The onboard real-time safety verification linkage submodule, located within the flight control system, performs pre-comparison and verification of the original torque and power distribution commands output by the flight control system. It identifies control errors such as overload, insufficient thrust-to-weight ratio, and torque mismatch, and progressively limits the motor output power when the load instantaneously exceeds the limit. The onboard real-time safety verification linkage submodule is equipped with an independent onboard local parameter memory and verification operation logic unit, enabling it to independently complete all verification work based on locally pre-stored safety boundary parameters without relying on real-time data intervention from external cloud or ground stations.

2. The fully electronic attitude control system for a mechanical-free vector disc-shaped low-altitude manned aircraft according to claim 1, characterized in that, The hybrid energy storage power generation unit adopts a constant range-extending power generation combined with a battery floating charge of 70%~85% SOC, and can achieve continuous 3-hour fixed-point hovering under windless full-load conditions.

3. The fully electronic attitude control system for a mechanical-free vector disc-shaped low-altitude manned aircraft according to claim 1, characterized in that, The three-layer composite electromagnetic shielding interlayer has a total weight of 15kg; the rated power of a single oil-cooled motor is 200kW, and the total peak power of the two motors is 480kW, each equipped with an independent oil-cooling heat dissipation circuit; the flight control software has a built-in ±45° heading rotation hard limit.

4. The fully electronic attitude control system for a mechanical-free vector disc-shaped low-altitude manned aircraft according to claim 1, characterized in that, The airborne real-time safety verification linkage submodule can establish data communication with the cloud-vehicle-airborne universal AI command verification system, share the whole machine power and load safety boundary parameters, and realize two-layer cross-verification of flight control commands.

5. A fully electronic attitude control method for a mechanically vector-disc-shaped low-altitude manned aircraft, characterized in that, Running on a 400Hz high-frequency redundant flight control processor, the following steps are included: S1. The high-speed differential bus of the aircraft continuously collects dynamic load, IMU attitude, ambient wind speed, motor operating conditions, and battery SOC sensor data at a period of 2.5ms. S2. Retrieve the aerodynamic matching parameters of the whole machine, the 350kg full load weight boundary, and the minimum safe thrust-to-weight ratio constraint threshold from the onboard local parameter memory cache. S3. Based on the dual-motor electronic differential logic, the original attitude torque calculation for pitch, yaw, and emergency braking is completed; S4. Input the calculated original attitude torque and power distribution parameters into the airborne real-time safety verification linkage submodule, compare them with the pre-stored safety boundary parameters, and identify overload, insufficient thrust-to-weight ratio, torque mismatch control errors. S5. If a control error is detected during verification, the motor output power is gradually limited, and a corrected motor power difference control command is issued; if there is no control error, the original drive signal is issued directly; in the hovering condition, the micro electronic servo guide plate is synchronously linked to fine-tune the airflow and stabilize the body. S6. Store samples of gust and heavy-load maneuver flight conditions, and adaptively iteratively update the overall power matching verification threshold. The S7 maintains a constant 18kW range-extending power generation + battery 70%~85% SOC floating charge energy dispatch mode to achieve long-term fixed-point hovering energy balance.

6. A computer-readable storage medium, characterized in that, The storage medium stores a computer program, which, when executed by a 400Hz high-frequency redundant flight control processor, implements the fully electronic attitude control method for a mechanical vector disc low-altitude manned aircraft as described in claim 5.

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