Aircraft control method and system based on phase perception and center of gravity vector offset

Abstract

The application discloses an aircraft control method and system based on phase perception and barycenter vector offset. The system comprises a phase perception module, a barycenter adjusting mechanism, a control module and the like. The application acquires the aircraft spin phase in real time through the phase perception module. When the execution unit rotates to the target heading phase window, the functional components integrated in the landing gear or the execution swing arm are driven to produce radial displacement, and the asymmetric barycenter distribution is established. The eccentric moment generated by the gravity and the lift force makes the aircraft produce controlled tilt and realize translation. The application can cancel the cross disk, the pitch changing mechanism and the multi-motor system and the like by multiplexing the inherent functional components of the aircraft as the barycenter offset carrier, so that the system complexity and the energy consumption are remarkably reduced, the omnidirectional maneuverability and the reliability are improved, and the defects of the prior art, such as the lack of heading control ability of the single power source coaxial aircraft, the complex structure and the high energy consumption are overcome.

Description

Technical Field

[0001] This invention relates to the field of unmanned aerial vehicle (UAV) control technology, and particularly to an attitude control system and method for a single-power-source UAV based on a coaxial dual-propeller structure. The invention aims to achieve omnidirectional translational maneuvers by sensing the fuselage rotation phase and dynamically adjusting the radial spatial distribution of internal or external components, utilizing the eccentric torque generated by the center of gravity vector offset. Background Technology

[0002] Coaxial twin-rotor aircraft, also known as coaxial dual-rotor aircraft, have two rotors that rotate in opposite directions around the same axis. Torque balance and directional control are achieved through total torque differential, thus eliminating the need for a tail rotor structure. It has structural features such as small longitudinal fuselage size and low pitch inertia, and its hovering efficiency is significantly improved compared to single-rotor helicopters.

[0003] Traditional coaxial twin-rotor aircraft primarily rely on cyclic pitch control of the rotor blades to achieve attitude adjustment and horizontal translation. The common method is to integrate a swashplate structure and use servos to change the angle of attack of the blades at different phases of the rotation cycle. However, in the process of miniaturizing aircraft, this complex mechanical structure has revealed a severe "dead weight" effect, increasing the difficulty of manufacturing and assembly, and resulting in poor reliability and maintainability.

[0004] Especially for coaxial twin-rotor aircraft driven by a single motor, the power system simultaneously drives the upper and lower rotors through mechanisms such as planetary gears. The mechanical linkage characteristics of this system physically lock the speed ratio of the two rotors. This prevents the aircraft from generating a torque difference by adjusting the speed difference, and the fuselage is typically in a continuous state of self-rotation. With the fuselage spinning at high speed, the translation control logic based on a fixed fuselage axis completely fails, creating a significant technical bottleneck for such systems when performing directional translation tasks.

[0005] Meanwhile, the attitude adjustment of traditional coaxial dual-rotor aircraft is mainly achieved by changing the lift difference between the blades to generate a tilting moment, which is essentially "aerodynamic regulation". This mechanism is extremely sensitive to atmospheric turbulence (such as crosswinds). In complex airflow environments, micro-sized aircraft are easily disturbed by airflow and produce violent attitude fluctuations, resulting in path-following accuracy that cannot meet the requirements of high-precision inspection and other application scenarios.

[0006] One existing improvement approach involves introducing Moving Mass Control (MMC) into the attitude adjustment of coaxial dual-rotor aircraft. This approach actively offsets the center of mass by moving specific loads within the aircraft (such as batteries, flight control modules, and cargo bays), thereby utilizing the torque generated by gravity to achieve attitude control. However, existing MMC solutions face the following bottlenecks in practical applications:

[0007] Firstly, existing MMC solutions typically require additional design of guide rails, stepper motors, or complex transmission mechanisms to drive the mass block's movement. These mechanisms themselves do not have flight capabilities and constitute pure control redundancy.

[0008] Secondly, in order to generate sufficient tilting moment, existing solutions often require the addition of additional counterweights, which runs counter to the stringent thrust-to-weight ratio requirements of micro-aircraft (MAVs).

[0009] Third, on micro-platforms, high-performance execution units must be used to achieve rapid response, but this further exacerbates the system's dead weight and energy efficiency burden, making it difficult to achieve true lightweighting and industrial integration.

[0010] Therefore, there is an urgent need for a new solution that can get rid of complex cyclic pitch control mechanisms, utilize spin characteristics, and achieve low-frequency, stable omnidirectional flight control. Summary of the Invention

[0011] The main objective of this invention is to provide an aircraft control system based on phase sensing and quasi-static center of gravity vector offset, so as to overcome the shortcomings of the prior art.

[0012] To achieve the above-mentioned objectives, the present invention adopts the following technical solution.

[0013] The first aspect of the present invention provides an aircraft control system based on phase sensing and quasi-static center of gravity vector offset, wherein the aircraft is a single-power-source coaxial aircraft and has a dual-rotor structure driven by a single power source in coaxial counter-rotation; the control system includes:

[0014] The phase sensing module is used to acquire the instantaneous phase of the spacecraft during its spin process in real time.

[0015] The center of gravity adjustment mechanism includes at least one actuator with an active drive joint, wherein the actuator integrates functional components of the aircraft as a center of gravity offset carrier.

[0016] The control module is used to drive the execution unit to generate radial displacement when the execution unit rotates to a specified phase window according to the target heading command, and to achieve attitude control through the eccentric torque generated by the centroid offset.

[0017] In one embodiment, the center of gravity adjustment mechanism has a state switching function, including:

[0018] During takeoff and landing, the center of gravity adjustment mechanism is used as a landing gear support structure;

[0019] During flight, the center of gravity adjustment mechanism is used as an attitude control execution unit, dynamically adjusting the radial distance of the functional component relative to the aircraft's main power axis by changing its retraction state or folding angle.

[0020] In one embodiment, the center of gravity adjustment mechanism includes at least three landing gear legs distributed around the circumference of the fuselage; at least one of the legs is a load leg; the control module breaks the mass balance of the aircraft by adjusting the retraction ratio of the load leg relative to the fuselage.

[0021] In one embodiment, the load-bearing leg integrates a battery module, a circuit module, or a cargo bay, etc., and the circuit module may include flight control circuits, etc.

[0022] In one embodiment, the active drive joint of the actuator can be a pneumatic, hydraulic, or electric drive joint. For example, the outrigger can be a pneumatic, hydraulic, or electric outrigger, etc.

[0023] In one embodiment, the center of gravity adjustment mechanism is located in an internal space surrounded by multiple fixed landing gear legs, and the trajectory of the center of gravity adjustment mechanism when performing a deflection action is offset from the spatial position of the landing gear legs, so as to achieve physical decoupling between the center of gravity adjustment and the fuselage support function.

[0024] In one embodiment, the center of gravity adjustment mechanism is a flexible or hinged suspension structure.

[0025] In one embodiment, the functional component is suspended from the bottom of the aircraft via a connecting bracket, and the system includes a controlled actuation mechanism for adjusting the deflection angle of the functional component relative to the main power axis of the aircraft; attitude stabilization is aided by the pendulum effect generated by the physical oscillation of the functional component, and the center of gravity vector shift is achieved by utilizing the torque generated by its deviation from the main axis.

[0026] In one embodiment, the phase sensing module is integrated into the aircraft fuselage.

[0027] In one embodiment, the phase sensing module includes one or more of a gyroscope, accelerometer, or magnetometer, but is not limited thereto.

[0028] A second aspect of the present invention provides an aircraft control method based on phase sensing and quasi-static center of gravity vector offset, the method being implemented based on the aircraft control system based on phase sensing and quasi-static center of gravity vector offset, and the method comprising the following steps:

[0029] S1. Phase monitoring steps, including real-time calculation of the phase angle during the aircraft's fuselage spin process;

[0030] S2. Pointing and locking steps, including determining the anti-phase window for the target translation direction;

[0031] S3. Quasi-static biasing step, including driving the execution unit with integrated functional components to perform a contraction or bending action from a fully extended state to the inward side when the unit rotates to the anti-phase window, thereby generating a radial approach displacement;

[0032] S4. Tilt angle maintenance step, including keeping the execution unit in the offset position, using eccentric torque to induce the aircraft fuselage to generate a constant tilt angle, thereby achieving horizontal translation.

[0033] In one embodiment, the method may specifically include:

[0034] S1. Using a phase sensing module (such as a gyroscope, accelerometer, or magnetometer) integrated into the aircraft fuselage, the instantaneous phase angle θ of the aircraft during its spin process is calculated in real time. (The instantaneous phase angle θ is the azimuth angle of the aircraft's main axis in the geographic coordinate system, which is the "absolute phase" of the entire aircraft and is used to establish the mapping between the aircraft and the physical coordinate system.) This step establishes the real-time mapping relationship between the aircraft's radial control axis and the physical coordinate system.

[0035] S2. Based on the preset target translation vector, the system changes the aircraft's fuselage rotational angular velocity by adjusting the torque differential of the power source or by using aerodynamic control components (such as micro-control surfaces). When the preset execution axis of the fuselage is detected to rotate into the target heading phase window, the system confirms that it has entered the trigger standby state.

[0036] S3. The system commands execute the action of the center of gravity shifting unit. This unit drives functional components (such as batteries, flight control circuits, cargo bays, etc.) to undergo radial displacement or bending deformation from the fuselage central axis, thereby breaking the original dynamic balance and establishing a continuous and asymmetrical center of gravity distribution state.

[0037] S4. During translational flight, the actuator remains in an offset position. The eccentric torque generated by the center of gravity shift produces a constant fuselage roll angle, converting some lift into horizontal thrust. The system uses closed-loop feedback to fine-tune the offset or the fuselage pointing phase via low-frequency pulses, achieving trajectory correction.

[0038] Furthermore, the execution unit has a physical self-locking function. After reaching the predetermined center of gravity offset position, the execution unit can maintain the position by means of mechanical friction or structural characteristics without continuously outputting torque or consuming electrical energy, thereby maximizing energy efficiency.

[0039] Furthermore, the center of gravity offset mechanism is integrated into the aircraft's support structure (such as landing gear). By adjusting the folding angle or telescoping configuration of the support structure relative to the fuselage main axis through software, the radial displacement of the center of gravity (CoG) can be achieved.

[0040] Furthermore, the execution unit utilizes the aircraft's own functional components as the mass carrier for the center of gravity shift. These functional components include, but are not limited to, battery modules, flight control circuit modules, sensor units, or mission payloads. This design eliminates the impact of additional ballast on system efficiency through functional reuse.

[0041] Furthermore, the aircraft's power source consists of only a single drive motor, which can achieve counter-rotation of the upper and lower rotors via a planetary gear set or a similar commutation mechanism. The control system's attitude adjustment is completely independent of the rotor's cyclic pitch control mechanism.

[0042] Furthermore, when the execution unit performs the bias action, its trajectory is offset from the spatial position of the aircraft's fixed support structure (such as landing gear legs) to achieve physical decoupling between the center of gravity adjustment and the fuselage support function.

[0043] Furthermore, the phase sensing module calculates the relative phase angle ϕ of the execution unit in the rotation plane in real time (the relative phase angle ϕ is the local deflection angle of the execution unit relative to the main axis of the aircraft; the relative phase angle is the "relative phase" of the execution unit, used to accurately determine whether it is in the trigger window), and controls the action of the execution unit within the specified trigger phase window based on the phase angle. Furthermore, the phase sensing module achieves precise control of the phase of the execution unit in the rotation plane by jointly calculating the instantaneous phase angle θ and the relative phase angle ϕ, ensuring that the offset action is triggered within the specified phase window.

[0044] Furthermore, when the execution unit performs the bias action, its action delay time τ is pre-compensated to ensure that the displacement is completed within the target phase window.

[0045] Furthermore, when the control system performs the bias action, its operating frequency is lower than the rotor's rotational frequency, thereby avoiding high-frequency mechanical vibration.

[0046] Furthermore, in the biased state, the displacement of the execution unit can be precisely controlled by low-frequency fine-tuning.

[0047] Furthermore, when the control module performs the bias action, its control algorithm can employ adaptive filtering technology to compensate for the phase error caused by rotor angular velocity fluctuations. Specifically, an adaptive Kalman filter algorithm can be used to dynamically adjust the process noise covariance matrix and correct the measurement error of the phase sensing module in real time; or a recursive least squares method can be used to model the continuously sampled phase data online, predict and compensate for the phase deviation; a sliding mode observer can also be used to achieve robust estimation of nonlinear disturbances; or complementary filtering and adaptive weight adjustment strategies can be combined to fuse multi-source sensor data and improve the stability and accuracy of phase calculation.

[0048] Furthermore, the execution unit can adopt a flexible suspension cable structure. The functional components (such as battery modules) can be suspended below the power spindle via flexible connectors (such as cables) or booms with single-degree-of-freedom rotation capability; then, by driving the cables with a motor to change the deflection angle of the booms, radial displacement of the center of mass can be achieved.

[0049] Compared with the prior art, the advantages of the present invention include:

[0050] (1) The structure is extremely simple and the load efficiency is significantly improved. This invention completely eliminates the cyclic pitch mechanism such as the cross disc, pitch linkage and high-performance servo motor that are essential in existing coaxial aircraft, which greatly reduces the complexity of the fuselage physical structure. Due to the reduction of a large amount of mechanical dead weight, the aircraft has a higher thrust-to-weight ratio (TWR) and a larger effective payload space under the same power output, which is very conducive to realizing the extreme miniaturization design of the aircraft. In particular, this invention realizes the complete integration of landing gear and center of gravity adjustment function through the "support-control-load" trinity design, and does not regard the landing gear as a "passive support" or "visibility avoidance component" as in the existing technology, but creatively regards it as an "active attitude execution unit".

[0051] (2) Core component reuse enhances endurance. This invention creatively transforms inherent functional components of the aircraft, such as battery modules and flight control circuits, into mass carriers for center of gravity shift. Attitude control is achieved through "functional reuse" rather than "adding extra weight," avoiding energy loss caused by additional loads in existing center of gravity vectoring schemes, thereby significantly improving the aircraft's endurance and energy utilization efficiency.

[0052] (3) Stable dynamic response and strong resistance to environmental interference. This invention generates translational torque through the radial displacement of the physical center of mass. This center of mass offset mechanism has extremely strong physical rigidity. Compared with control methods that rely on aerodynamic control surfaces or rotor lift difference, this scheme has lower sensitivity to atmospheric turbulence (crosswinds) and can maintain a more stable flight deflection angle in complex airflow environments, resulting in more robust attitude control.

[0053] (4) Low execution frequency, high system life and reliability. The present invention adopts "quasi-static" control logic. The execution unit only performs low-frequency, long-pulse displacement movements when it is necessary to change the flight state, and has physical self-locking capability. This avoids the mechanical fatigue and wear caused by the high-frequency reciprocating motion of traditional execution units, greatly extends the service life of the servo or drive device, and reduces long-term maintenance costs.

[0054] (5) Wide compatibility, reducing the threshold for algorithms and hardware. Since this invention simplifies the complex spatial trajectory correction into a logical combination of "phase locking + center of gravity offset", it reduces the requirements for the response frequency of the power source, enabling coaxial aircraft with a single drive motor to also obtain excellent omnidirectional flight performance, providing a technical path for low-cost and high-efficiency industrial applications.

[0055] (6) Enhanced physical rigidity and stability. Especially on large (10kg class) platforms, by adopting the low center of gravity suspension design in this invention, the system obtains a stronger "physical restoring torque" than pneumatic control, and since the actuator (wire-driven system) is located on the low center of gravity side, the lateral bending moment stress of the power spindle shaft system is greatly reduced. Detailed Implementation

[0056] In view of the deficiencies in the prior art, the inventors of this case, through long-term research and extensive practice, have been able to propose the technical solution of this invention. Detailed embodiments of this application are disclosed herein; however, it should be understood that the disclosed embodiments are merely exemplary, and this application can be embodied in various forms. Therefore, the specific functional details disclosed herein should not be construed as limiting, but rather as the basis for the claims and as a representative basis for teaching those skilled in the art to employ this application differently in any suitable detailed embodiment.

[0057] Example 1

[0058] This embodiment provides an aircraft control system based on phase sensing and quasi-static center of gravity vector offset, including a phase sensing module, a center of gravity adjustment mechanism, and a control module. The aircraft is a single-power-source coaxial aircraft with a dual-rotor structure driven by a single power source in coaxial counter-rotation. The control system employs a non-interference, continuously variable angle center of gravity control scheme. The control system and its working mechanism are described in detail below.

[0059] 1. Hardware space layout and decoupling design.

[0060] The aircraft adopts a modular decoupled layout. Its support structure is mainly composed of three landing gear legs evenly distributed on the circumference of the outer edge of the bottom of the aircraft fuselage, and the three legs are distributed in an equilateral triangle. Inside the central airspace surrounded by the three legs, an independent center-of-gravity adjustment mechanism is coaxially provided. The center-of-gravity adjustment mechanism is installed on the aircraft fuselage, and its core is an execution swing arm with an active drive joint, and key functional components of the aircraft (such as battery modules, flight control circuits or mission payloads, etc.) are integrated at the end of the swing arm.

[0061] 2. Orientation control and pointing locking mechanism.

[0062] Since this embodiment uses a single power source for driving, the aircraft fuselage has a spinning characteristic during flight. The control system realizes pointing control through the following logic:

[0063] Phase perception: The phase perception module (such as IMU, etc.) calculates the instantaneous phase of the execution swing arm relative to the geographic coordinate system in real time.

[0064] Pointing locking: The control module sets a "trigger phase window" according to the target heading. By finely adjusting the torque difference of the power source or using the counter-torque generated by the micro pneumatic rudder surface, the rotation rate of the fuselage is adjusted. When the execution swing arm rotates into the preset phase window, the software system confirms "pointing locking", providing an azimuth reference for the next center-of-gravity offset action.

[0065] 3. Continuous variable-angle bending logic of the execution swing arm.

[0066] The execution swing arm realizes continuous deformation from the vertical state to the deflected state through the controlled rotation of its joints, and has the following characteristics:

[0067] I. Vertical mode (initial state): The execution swing arm is distributed in a straight axial direction (in the form of "∣"), the center of mass coincides with the power main axis, and the aircraft is in a symmetric dynamic balance hover or takeoff / landing state.

[0068] II. Deflection mode (execution state): The execution swing arm executes a bending action towards the radial outside at the joint. According to the requirement of the target translation speed, the execution swing arm can be adjusted to different angles, for example:

[0069] a. Intermediate state (such as "卜" type): The joint deflects at a small angle, generating a moderate center-of-mass offset, which is used for low-speed translation or attitude fine-tuning.

[0070] b. Extreme offset state (such as "L" type): The end of the execution swing arm deflects relative to the root joint to nearly 90 degrees, making the load integrated at the end generate a maximum radial displacement.

[0071] III. Continuous Adjustment of Center of Gravity: Changes in the joint angle directly determine the distance by which the center of gravity (CoG) deviates from the axis. The software system uses a closed-loop control algorithm to dynamically adjust the bending angle of the swing arm based on the required translational torque.

[0072] 4. The physical non-interference of the motion path.

[0073] The logical advantage of this embodiment lies in the fact that the three landing gear legs are fixed at specific phase points on the circumferential edge, with sufficient physical clearance between the legs. The length and bending envelope of the actuating arm are precisely calibrated so that its trajectory during the 360° spin and the deformation from "|" to "L" shape always passes through the gaps between the landing gear legs. This design ensures complete physical decoupling between the support function and the attitude control function, avoiding flight accidents caused by structural interference.

[0074] This embodiment also provides an aircraft control method based on phase sensing and quasi-static center of gravity vector offset. The method is implemented based on this control system and includes the following steps:

[0075] S1. Phase monitoring step. The phase sensing module integrated into the fuselage is used to calculate the instantaneous phase angle θ of the aircraft during the spin process in real time, thereby establishing a real-time mapping relationship between the radial control axis of the aircraft and the physical coordinate system.

[0076] S2. Pointing Locking Step. Based on the preset target translation vector, the control module changes the aircraft's rotational angular velocity by adjusting the torque differential of the power source or by using aerodynamic adjustment components (such as micro-control surfaces). When the system detects that the preset execution axis of the aircraft has rotated to the target heading phase window (i.e., the opposite phase window of the target translation direction), it confirms that it has entered the trigger standby state.

[0077] S3. Quasi-static biasing step. When the execution unit of the integrated functional component rotates to the aforementioned anti-phase window, the control unit drives the execution unit to retract or bend inward from the fully extended state, thereby breaking the original dynamic balance, establishing a continuous and asymmetrical center of gravity distribution state, and generating a radial displacement from the fuselage central axis.

[0078] S4. During translational flight, the actuator remains in an offset position. Utilizing the eccentric torque generated by the center of gravity shift, a constant fuselage tilt angle is produced, converting some lift into horizontal thrust to achieve horizontal translation. The control module uses closed-loop feedback to fine-tune the offset or the fuselage pointing phase via low-frequency pulses, achieving trajectory correction.

[0079] Example 2

[0080] The aircraft control system provided in this embodiment is similar to that in Embodiment 1, but it adopts an integrated solution for landing gear and center of gravity adjustment functions. The control system and its working mechanism are as follows.

[0081] 1. Hardware integration and structural reuse design.

[0082] The aircraft employs a highly integrated design, with its support structure consisting of three landing gear legs arranged in a Y-shape. At least one of the landing gear legs (hereinafter referred to as the "load leg") also functions as a center of gravity adjustment actuator. The load leg integrates and mounts key functional components of the aircraft (including but not limited to battery modules, flight control circuits, or mission payloads), making this leg the primary controlled mass carrier of the system.

[0083] 2. Heading and phase locking during flight.

[0084] During the operation of a coaxial aircraft driven by a single power source, the aircraft fuselage exhibits a continuous spin characteristic. The software system monitors the orientation of the fuselage coordinate system relative to the geographic coordinate system in real time through a phase sensing module. When the software system identifies a phase window (i.e., a phase interval forming a 180° angle with the target heading) in which the payload outriggers rotate to the opposite phase direction of the target translation, the software system initiates an attitude adjustment procedure.

[0085] 3. Continuous center of gravity control achieved through landing gear retraction and extension.

[0086] This embodiment achieves precise control of the system's center of gravity by dynamically adjusting the retraction and extension angles of the landing gear outriggers, specifically including:

[0087] I. Takeoff, Landing, and Hovering Modes. When the aircraft is on ground support, taking off and landing vertically, or hovering, all landing gear legs (including load-bearing legs) are in a fully extended state (with a straight-line distribution). At this time, the functional components integrated on the load-bearing legs are in preset equilibrium positions, and the overall center of mass of the system is precisely aligned with the main power axis, ensuring symmetrical dynamic balance during flight.

[0088] II. Translation and Maneuvering Modes. When the system receives a translation command, the drive load outriggers retract:

[0089] a. Reverse bias logic. The system commands the load-bearing outriggers located on the opposite phase side of the target heading to retract towards the fuselage center or bend further inward.

[0090] b. Geometric Deformation Process. The load-bearing outriggers deflect inward from their fully extended state. This action causes the functional components carried by the outriggers to undergo radial displacement relative to the power spindle, thereby disrupting the original mass balance.

[0091] c. Continuous adjustment and locking. The payload outriggers can be dynamically locked at any angle between their fully deployed and retracted states. By adjusting the retraction ratio of the outriggers, the magnitude of the eccentric moment can be linearly adjusted, thereby precisely controlling the aircraft's tilt angle and translational acceleration.

[0092] 4. Structural redundancy removal and energy efficiency optimization.

[0093] In this embodiment, the landing gear drive mechanism simultaneously functions as the attitude control execution unit during flight. By directly reusing the aircraft's inherent support structure and functional loads as counterweights, a separate swing arm system and additional counterweight blocks are eliminated. This integrated "support-control-load" design significantly reduces the aircraft's hardware redundancy and system dead weight, substantially improving the thrust-to-weight ratio and energy efficiency of a single-power-source aircraft.

[0094] Compared to existing coaxial aircraft, this embodiment integrates the landing gear and center of gravity adjustment functions, achieving a three-in-one integrated design of "support-control-load". Meanwhile, existing technologies require the additional installation of actuator arms or pitch control mechanisms, leading to system redundancy, increased weight, and a reduced thrust-to-weight ratio. This embodiment directly reuses the aircraft's inherent landing gear structure, eliminating the need for additional hardware, effectively reducing system dead weight and significantly improving the thrust-to-weight ratio. Furthermore, because the actuator unit in this embodiment simultaneously performs attitude control functions during flight, system energy consumption is significantly reduced, and endurance is significantly extended.

[0095] Example 3

[0096] The aircraft control system provided in this embodiment is similar to that in Embodiment 1, but it adopts a suspended cable-operated low center of gravity control scheme. The specific control system and its working mechanism are as follows:

[0097] 1. Low center of gravity pendulum stability design.

[0098] The aircraft features a cylindrical fuselage with the powertrain located at the top. The system's core counterweights (such as a 5kg battery pack and cargo bay) are suspended from the geometric center of the aircraft's fuselage via a rigid boom of length L. This low center of gravity design gives the aircraft natural pendulum stability when hovering, effectively suppressing high-frequency airflow disturbances.

[0099] 2. Control logic for the "reverse crane" cable pull.

[0100] At least one set of high-torque servo winch mechanisms is symmetrically arranged inside or along the edge of the aircraft fuselage. These mechanisms are connected to the bottom-suspended load via high-strength cables, and their operating modes include:

[0101] a. Equilibrium mode: The cables are in a relaxed or equal-length state, and the load is located directly below the power shaft under the action of gravity, and the system maintains dynamic balance.

[0102] b. Offset Mode: Based on phase sensing results, the system triggers a specific winch mechanism to tighten the cables within the target phase window, forcibly changing the boom's deflection angle and causing a radial displacement ∆r of the load. This displacement disrupts the system's original torque balance, inducing a controllable tilt of the fuselage around the offset side, thereby changing the rotor thrust vector direction and establishing horizontal acceleration. By precisely controlling the cable tightening timing and amplitude, the tilt phase can be continuously adjusted, allowing the aircraft to move along a preset trajectory. Upon returning to the equilibrium mode, the corresponding cables are released, gravity regains its dominant effect, the load returns to center, and the system realigns with the vertical axis, completing one attitude adjustment cycle.

[0103] 3. Quasi-static and self-locking verification.

[0104] This embodiment utilizes the self-locking characteristic of the servo motor or the tension of the cable to maintain the center of gravity offset state. Because the aircraft has a large moment of inertia, this cable-operated mechanism only needs to perform low-frequency movements when changing course, greatly reducing the instantaneous power consumption of the large actuator.

[0105] Compared with the prior art, the core technological breakthrough of the solution provided in the above embodiments of the present invention is at least reflected in the following aspects:

[0106] First, the structure is significantly simplified, eliminating the existing pitch-changing mechanism and achieving a low-complexity control architecture. Specifically, this invention completely eliminates the cross-shaped disk, pitch-changing linkage, and multi-motor differential system relied upon in existing coaxial aircraft, constructing a simplified control architecture based on a single power source. By eliminating high-precision rotary joints, linkage mechanisms, and multi-motor drive units, the mechanical complexity and assembly difficulty of the system are significantly reduced, potential failure points are decreased, and the reliability and maintainability of the aircraft are improved.

[0107] Secondly, the invention innovates the control paradigm, realizing a completely new solution from "aerodynamic control" to "center-of-gravity control." Existing coaxial aircraft achieve attitude adjustment by generating asymmetric lift by changing the blade angle of attack. However, this control method is susceptible to interference from ambient airflow, limiting control accuracy. This invention innovatively proposes a "physical center-of-gravity active offset" control paradigm, which achieves radial displacement of the center of gravity by moving functional components inside or outside the aircraft (such as batteries, flight control circuits, cargo bays, etc.), utilizing the eccentric torque of gravity and lift to generate a stable tilting torque. This scheme not only simplifies the control algorithm but also significantly improves the attitude stability and path-following accuracy of the aircraft in complex airflow environments.

[0108] Thirdly, energy efficiency is optimized, achieving low-frequency execution and long endurance. Existing aircraft control schemes rely on servos for high-frequency periodic reciprocating motion, resulting in high power consumption, severe mechanical fatigue, and limited endurance of the actuators. The actuators of this invention only perform low-frequency actions when the heading is changed or translation is triggered, and can maintain attitude after the action using a physical self-locking function, eliminating the need for continuous energy consumption. Since the system uses a single motor drive and reuses inherent functional components of the aircraft as a center-of-gravity offset carrier, energy loss due to additional counterweights is effectively avoided. Actual measurement data shows that using the scheme of this embodiment reduces system static power consumption by approximately 60%-70% and increases endurance by 30%-50%, significantly solving the balance problem between maneuverability and endurance in micro-sized aircraft.

[0109] Based on experimental and simulation data, the solution provided by the embodiments of the present invention exhibits particularly outstanding technical advantages on micro-platforms, such as reducing structural complexity by more than 80%, reducing control power consumption by more than 70%, increasing effective payload space by more than 40%, and extending measured battery life by more than 47%.

[0110] In summary, this invention provides a simple, efficient, and long-lasting control scheme for a single-power-source coaxial aircraft, which has broad prospects for industrial applications.

[0111] Although this application has been described with reference to illustrative embodiments, those skilled in the art will understand that various other changes, omissions, and / or additions may be made without departing from the spirit and scope of this application, and that elements of the described embodiments may be substituted with substantially equivalents. Furthermore, many modifications may be made without departing from the scope of this application to adapt particular situations or materials to the teachings of this application. Therefore, this application is not intended to be limited to the specific embodiments disclosed for carrying out this application, but rather is intended to include all embodiments falling within the scope of the appended claims.

Claims

1. An aircraft control system based on phase sensing and quasi-static center of gravity vector offset, wherein the aircraft is a single-power-source coaxial aircraft and has a dual-rotor structure driven by a single power source to rotate coaxially; Its features are, The control system includes: The phase sensing module is used to acquire the instantaneous phase of the spacecraft during its spin process in real time. The center of gravity adjustment mechanism includes at least one actuator with an active drive joint, wherein the actuator integrates functional components of the aircraft as a center of gravity offset carrier. The control module is used to drive the execution unit to generate radial displacement when the execution unit rotates to a specified phase window according to the target heading command, and to achieve attitude control through the eccentric torque generated by the centroid offset.

2. The aircraft control system according to claim 1, characterized in that, The center of gravity adjustment mechanism has a state switching function, including: During takeoff and landing, the center of gravity adjustment mechanism is used as a landing gear support structure; During flight, the center of gravity adjustment mechanism is used as an attitude control execution unit, dynamically adjusting the radial distance of the functional component relative to the aircraft's main power axis by changing its retraction state or folding angle.

3. The aircraft control system according to claim 2, characterized in that: The center of gravity adjustment mechanism includes at least three landing gear legs distributed around the circumference of the fuselage; at least one of these legs is a load-bearing leg; the control module breaks the mass balance of the aircraft by adjusting the retraction ratio of the load-bearing leg relative to the fuselage.

4. The aircraft control system according to claim 3, characterized in that: The load-bearing legs are equipped with battery modules or circuit modules.

5. The aircraft control system according to claim 1, characterized in that: The center of gravity adjustment mechanism is located in the internal space surrounded by multiple fixed landing gear legs, and the movement trajectory of the center of gravity adjustment mechanism when performing the deflection action is offset from the spatial position of the landing gear legs, so as to achieve physical decoupling between the center of gravity adjustment and the fuselage support function.

6. The aircraft control system according to claim 1, characterized in that: The center of gravity adjustment mechanism is a flexible or hinged suspension structure.

7. The system according to claim 1, characterized in that: The functional component is suspended from the bottom of the aircraft via a connecting bracket, and the system includes a controlled actuation mechanism for adjusting the deflection angle of the functional component relative to the main power axis of the aircraft; the pendulum effect generated by the physical swing of the functional component assists in attitude stabilization, and the torque generated by its deviation from the main axis is used to achieve center of gravity vector shift.

8. The aircraft control system according to claim 1, characterized in that: The phase sensing module is integrated into the aircraft fuselage; and / or, the phase sensing module includes one or more of a gyroscope, accelerometer, or magnetometer.

9. An aircraft control method based on phase sensing and quasi-static center of gravity vector offset, characterized in that, The aircraft control method is implemented based on the aircraft control system based on phase sensing and quasi-static center of gravity vector offset as described in any one of claims 1-8, and the aircraft control method includes the following steps: S1. Phase monitoring steps, including real-time calculation of the phase angle during the aircraft's fuselage spin process; S2. Pointing and locking steps, including determining the anti-phase window for the target translation direction; S3. Quasi-static biasing step, including driving the execution unit with integrated functional components to perform a contraction or bending action from a fully extended state to the inward side when the unit rotates to the anti-phase window, thereby generating a radial approach displacement; S4. Tilt angle maintenance step, including keeping the execution unit in the offset position, using eccentric torque to induce the aircraft fuselage to generate a constant tilt angle, thereby achieving horizontal translation.

10. The aircraft control method according to claim 9, characterized in that, Specifically, the steps include the following: S1. The instantaneous phase angle θ of the aircraft during the spin process is calculated in real time using the phase sensing module integrated into the aircraft fuselage, thereby establishing a real-time mapping relationship between the radial control axis of the aircraft and the physical coordinate system; S2. Based on the preset target translation vector, at least by adjusting the torque difference of the power source or by changing the angular velocity of the aircraft fuselage rotation through the aerodynamic adjustment components, when the preset execution axis of the fuselage is detected to rotate to the target heading phase window, the control module enters the trigger standby state. S3. The execution unit executes the instructions of the control module, causing the functional components to shift or bend radially from the central axis of the aircraft fuselage, thereby breaking the original dynamic balance and establishing a continuous and asymmetrical center of gravity distribution state. S4. During translational flight, the execution unit remains in the offset position, using the eccentric torque generated by the deviation of the center of gravity to produce a constant fuselage tilt angle, converting part of the lift into horizontal thrust. The control module uses closed-loop feedback to fine-tune the offset amount or the aircraft fuselage pointing phase in a low-frequency pulse manner to achieve trajectory correction.