Two-dimensional vector turbojet aircraft and control method thereof

Abstract

The invention discloses a two-dimensional vector turbojet aircraft and a control method thereof. Swing jet blades are arranged at nozzles of vector turbojet engines of the aircraft; according to the two-dimensional vector control assembly, a first driving unit drives the vector turbojet engine to integrally deflect, and the thrust direction of the engine is changed; the second driving unit drives each swing spraying blade to deflect independently, and the airflow direction of an engine nozzle is changed. The angle calculation unit calculates the overall deflection angle of each vector turbojet engine and the deflection angle of the corresponding swing spraying blade according to the actual attitude and the expected attitude of the aircraft, the overall deflection angles of the vector turbojet engines are controlled through the first driving unit, and meanwhile the deflection angles of the swing spraying blades are controlled through the second driving unit. And the lateral thrust and the yaw moment of the aircraft are adjusted. By introducing a two-dimensional vector control structure, the aircraft can directly and independently generate lateral thrust and yaw control torque, and flight stability, control accuracy and maneuvering efficiency are effectively improved.

Description

Technical Field

[0001] This invention relates to the field of aircraft control technology, and in particular to a two-dimensional vector turbojet aircraft and its control method. Background Technology

[0002] Vertical takeoff and landing (VTOL) light manned aircraft are ideal vehicle platforms for future urban air commuting and rapid cross-domain access scenarios. Their core value lies in their ability to adapt to takeoff and landing requirements in confined spaces and complex terrains. Among various technological approaches, vectoring turbojet aircraft combine the advantages of vector thrust technology and the high thrust-to-weight ratio of turbojet engines. By deflecting the nozzle, they achieve seamless switching between VTOL and high-speed cruise modes, combining the flexibility of helicopters with the efficiency of fixed-wing aircraft. They demonstrate great potential in complex scenario rescue and low-altitude manned and cargo transport.

[0003] A typical existing technical solution is a one-dimensional vector turbojet control structure. In this structure, the thrust direction is changed by driving the turbojet engine to rotate as a whole, thereby achieving a balance between the vertical and forward forces of the aircraft. Pitch and roll moments are generated through the coordinated control of the yaw angles of the four engines to complete the forward flight and roll maneuvers. This scheme uses a PID algorithm for closed-loop feedback control to maintain attitude stability during flight. A key reason for adopting this control strategy is the inherent characteristic of turbojet engines having a long thrust dynamic response period. Therefore, keeping the thrust of each engine consistent during flight and instead achieving attitude control by rapidly adjusting its yaw angles becomes a reasonable choice.

[0004] However, the aforementioned one-dimensional vector turbojet control structure has a fundamental technical flaw: because its thrust direction can only deflect in a single plane, it cannot provide lateral (Y-direction) thrust to the aircraft, and also loses the ability to control yaw (Z-direction) moment. This directly leads to uncontrollable yaw motion when the aircraft equipped with this structure hovers. To compensate for this deficiency, the aircraft has to indirectly adjust its heading through complex roll and pitch coupling maneuvers, which not only significantly increases the complexity of the control algorithm, but also severely reduces the stability and control efficiency of the flight process, and leads to a significant reduction in endurance performance due to redundant maneuvers. Therefore, solving the problem of independent and controllable lateral force and yaw moment of the aircraft has become the key to improving the performance and practicality of this type of aircraft. Summary of the Invention

[0005] The purpose of this invention is to provide a two-dimensional vector turbojet aircraft and its control method. By introducing a two-dimensional vector control structure, the aircraft can directly and independently generate lateral thrust and yaw control torque, thereby effectively improving flight stability, control accuracy, and maneuverability.

[0006] To solve the above-mentioned technical problems, the present invention provides a two-dimensional vector turbojet aircraft, including: a fuselage, a plurality of vector turbojet engines located on the fuselage, and a two-dimensional vector control assembly, wherein each of the vector turbojet engines is provided with a swaying spray blade at its nozzle. The two-dimensional vector control component includes: a first driving unit, a second driving unit, and an angle calculation unit; The first drive unit drives each of the vector turbojet engines to deflect as a whole, changing the thrust direction of the vector turbojet engine; The second drive unit drives each of the aforementioned oscillating spray blades to deflect independently, thereby changing the direction of the airflow at the nozzle of the vector turbojet engine; The angle calculation unit calculates the overall deflection angle of each vector turbojet engine and the deflection angle of the corresponding oscillating jet blade based on the actual attitude and desired attitude of the aircraft. The first drive unit controls the overall deflection angle of the vector turbojet engine, while the second drive unit controls the deflection angle of the oscillating jet blade, thereby adjusting the lateral thrust and yaw moment of the aircraft.

[0007] Furthermore, the angle calculation unit is configured as follows: Based on the error between the desired attitude angle and the current attitude angle, the desired attitude angular velocity is calculated by the outer loop PID controller; Based on the error between the desired attitude angular velocity and the current attitude angular velocity, the desired three-axis control torque is calculated by the inner loop PID controller; Based on the desired three-axis control torque and desired aircraft altitude, and using the aircraft's dynamic model, the desired turbojet engine deflection angle of each of the vector turbojet engines and the desired blade deflection angle of each of the oscillating jet blades are calculated.

[0008] Furthermore, the desired blade deflection angle The calculation formula is: ; ; ; ; ; in, To achieve the desired roll control torque, To achieve the desired pitch control torque, To achieve the desired yaw control torque, For the desired lift in the Z direction, The thrust of a single turbojet engine; The desired turbojet engine deflection angle The calculation formula is: Furthermore, the formula for calculating the desired three-axis control torque by the inner-loop PID controller is as follows: ; ; ; in, For roll rate error, This is the roll angle acceleration error. For pitch angular velocity error, For pitch angle acceleration error, For yaw rate error, This refers to the yaw angle acceleration error. , , These are the proportional coefficient, integral coefficient, and derivative coefficient of the inner loop controller, respectively. The desired thrust in the Z direction The calculation formula is: ; in, The error between the aircraft's desired altitude and its current altitude. This represents the integral coefficient of the altitude controller.

[0009] Furthermore, the formula for calculating the desired attitude angular velocity obtained by the outer loop PID controller is as follows: ; ; ; in, These represent the errors between the expected and current values ​​of the roll, pitch, and yaw angles, respectively. is the proportional coefficient of the outer loop PID controller.

[0010] Furthermore, the dynamic model of the aircraft includes force balance equations and torque balance equations; The formula for calculating the force balance equation is as follows: ; The formula for calculating the torque balance equation is as follows: ; in, For the mass of the aircraft, Let be the position vector of the aircraft in the inertial frame. Let be the rotation matrix from the body coordinate system to the inertial coordinate system. The resultant force along the three axes in the body coordinate system. The gravity vector This is the drag vector. The moment of inertia of the aircraft. The angular acceleration vector in the body coordinate system. The three-axis control torque in the body coordinate system. This is the angular velocity vector in the body coordinate system. This is the damping torque vector.

[0011] Furthermore, the resultant force of the three axes in the body coordinate system The calculation formula is: ; in, The thrust of a single turbojet engine. Let i be the overall deflection angle of the i-th engine. Let be the deflection angle of the i-th engine's swivel spray blade.

[0012] Furthermore, the three-axis control torque in the body coordinate system The calculation formula is: ; in, Let Y be the lever arm of the engine relative to the aircraft's center of mass. The force arm of the engine relative to the center of mass of the aircraft is the X-axis direction.

[0013] Furthermore, the wind resistance vector The calculation model is as follows: ; in, Here is the damping coefficient matrix. ξ ˙ represents the velocity vector of the aircraft in the inertial coordinate system. These are the damping coefficients in the three axial directions, These represent the velocity components of the aircraft in the three axes.

[0014] Accordingly, a second aspect of the present invention provides a two-dimensional vector turbojet aircraft control method for controlling the aforementioned two-dimensional vector turbojet aircraft, the control method comprising the following steps: Obtain the current attitude angle, current attitude angular velocity, and desired attitude angle of the aircraft in the current control cycle, and calculate the attitude angle error between the current attitude angle and the desired attitude angle; Based on the attitude angle error, the desired attitude angular velocity is calculated; Calculate the attitude angular velocity error based on the current attitude angular velocity and the desired attitude angular velocity; Based on the attitude angular velocity error, the desired three-axis control torque is calculated; Based on the desired three-axis control torque and the desired aircraft altitude, the overall deflection angle of each of the vector turbojet engines and the corresponding deflection angle of the oscillating jet blades are calculated using the aircraft dynamics model. Based on the overall deflection angle of all the vector turbojet engines and the deflection angle of the corresponding swivel blades, the movement of the vector turbojet engines and their corresponding swivel blades is controlled to adjust the lateral thrust and yaw moment of the aircraft.

[0015] The above-described technical solutions of the embodiments of the present invention have the following beneficial technical effects: 1. By adding sway blades to each turbojet engine nozzle, a two-dimensional vector control structure was constructed, which is jointly generated by the overall deflection of the engine and the deflection of the nozzle blades. This fundamental structural innovation breaks through the limitation of traditional one-dimensional vector control, which can only provide pitch and roll control. For the first time, it introduces the ability to independently generate lateral thrust and yaw control torque for aircraft. This enables the aircraft to directly and accurately control yaw attitude and lateral translation when hovering and maneuvering at low speeds, without having to rely on inefficient and unstable pitch-roll coupling for indirect compensation. This eliminates uncontrollable yaw phenomena from a physical level and lays a solid foundation for stable flight under all operating conditions. 2. To address the newly added two-dimensional control degrees of freedom and the resulting strong coupling and nonlinear dynamic characteristics, this invention specifically designs a dual-closed-loop PID control architecture based on a precise dynamic model. This control method decouples the outer-loop attitude angle control from the inner-loop angular velocity control, and through the established mathematical model, accurately analyzes the attitude and torque commands from the higher levels into the overall deflection angle and the deflection angle of each engine's swivel nozzle blades at the lower level. This effectively overcomes the inherent defect of slow thrust response in turbojet engines, fully leverages the advantages of fast response in two-dimensional vector structures, and ensures that the aircraft can not only fly stably but also accurately and quickly track complex flight attitude commands, achieving synergistic optimization between the control system and the mechanical structure. 3. Based on the aforementioned structural and control innovations, the two-dimensional vector turbojet aircraft of this invention achieves a leap in overall performance. Since yaw and lateral motion become controllable, the aircraft avoids energy loss during maneuvers caused by compensating for excessive attitude, significantly improving aerodynamic efficiency and endurance. The direct control method also reduces attitude oscillations, greatly enhancing flight stability and safety. Ultimately, this solution enables simultaneous optimization of multiple key performance indicators, including takeoff and landing flexibility, maneuver accuracy, flight stability, and energy economy, providing comprehensive assurance for its reliable and efficient operation in complex application scenarios. Attached Figure Description

[0016] Figure 1 This is a schematic diagram of the structure of a two-dimensional vector turbojet aircraft provided in an embodiment of the present invention; Figure 2 This is a side view of a two-dimensional vector turbojet engine provided in an embodiment of the present invention; Figure 3 This is a block diagram of the control mode of a two-dimensional vector turbojet aircraft provided in an embodiment of the present invention; Figure 4 This is a flowchart of a two-dimensional vector turbojet aircraft control method provided in an embodiment of the present invention.

[0017] Attached image labels: 1. Airframe, 2. Vector turbojet engine, 3. Swinging jet blades, 4. Nozzle. Detailed Implementation

[0018] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to specific embodiments and the accompanying drawings. It should be understood that these descriptions are merely exemplary and not intended to limit the scope of the invention. Furthermore, descriptions of well-known structures and techniques are omitted in the following description to avoid unnecessarily obscuring the concept of the invention.

[0019] Please refer to Figure 1 , Figure 2 and Figure 3 The present invention provides a two-dimensional vector turbojet aircraft, including: a body 1, a plurality of vector turbojet engines 2 located on the body 1, and a two-dimensional vector control assembly, wherein each vector turbojet engine 2 is provided with a swaying spray blade 3 at its nozzle 4.

[0020] This invention, by introducing the key component 3, the sway nozzle blade, extends traditional one-dimensional thrust vector control to a two-dimensional plane, providing aircraft with a completely new aerodynamic control dimension. This structural layout forms the physical basis for achieving independent control of lateral force and yaw moment, enabling the aircraft to gain unprecedented attitude control freedom while maintaining the high thrust-to-weight ratio advantage of turbojet engines. From the perspective of principle and improvement, this structure innovatively integrates independently controllable aerodynamic control surfaces at the final stage of power output, namely the engine nozzle 4, while retaining the overall deflection function of the turbojet engine. This improved structure avoids the inherent defect of slow thrust response in turbojet engines, achieving rapid thrust reconfiguration by adjusting the microscopic direction of the airflow at the nozzle 4. Each engine thus becomes a propulsion unit with two degrees of freedom output capability, and the thrust vector it generates is no longer limited to a single plane but can be precisely decomposed in three-dimensional space. This upgrade of the basic structure provides a hardware-level implementation method for solving key problems such as uncontrollable yaw and lack of lateral maneuverability in the background technology.

[0021] Specifically, the two-dimensional vector control component includes: a first drive unit, a second drive unit, and an angle calculation unit; the first drive unit drives each vector turbojet engine 2 to deflect as a whole, changing the thrust direction of the vector turbojet engine 2; the second drive unit drives each swaying nozzle 3 to deflect independently, changing the airflow direction of the nozzle 4 of the vector turbojet engine 2. The aforementioned two-dimensional vector control component translates control commands into specific mechanical actions. Through the coordinated operation of the two drive units, it achieves decoupled control of the engine's macroscopic thrust direction and the nozzle's microscopic airflow direction. This separate drive architecture ensures the independence of the pitch, roll, yaw, and lateral control channels, providing execution-level assurance for precise six-degree-of-freedom flight control.

[0022] The first drive unit primarily serves the main thrust vector adjustment during vertical-to-horizontal flight mode transitions and the macroscopic control of pitch and roll attitude during cruise, by changing the engine mounting angle or driving the engine to rotate around its axis. The second drive unit focuses on high-frequency, small-range thrust fine-tuning, deflecting the engine exhaust by changing the angle of attack of the swivel nozzles 3, thereby superimposing an adjustable lateral component on top of the main thrust. This division of labor allows the system to simultaneously meet the power requirements for large-range attitude adjustments and the response speed requirements for rapid attitude stabilization. The angle calculation unit, as the control center, is responsible for coordinating the timing and amplitude of the actions of the two drive units to ensure that the thrust vector is synthesized as expected.

[0023] Furthermore, the angle calculation unit calculates the overall deflection angle of each vector turbojet engine 2 and the deflection angle of the corresponding swivel blade 3 based on the actual attitude and desired attitude of the aircraft. The first drive unit controls the overall deflection angle of the vector turbojet engine 2, while the second drive unit controls the deflection angle of the swivel blade 3, thereby adjusting the lateral thrust and yaw moment of the aircraft.

[0024] The angle calculation unit translates high-level flight mission commands into low-level actuator control parameters. By calculating and allocating control quantities for two degrees of freedom in real time, the aircraft can simultaneously and independently control its spatial attitude and translational motion. This process achieves a deep integration of flight mechanics and control engineering. The angle calculation unit continuously receives attitude feedback from the navigation system and command information from mission planning. By solving the force balance relationship of the aircraft in the current state, it inversely derives the required thrust vector distribution. For conditions requiring yaw moment or lateral thrust, the calculation unit prioritizes allocating the deflection task of the yaw nozzle 3, as its response speed is much faster than the overall engine deflection. During large-angle pitch or roll maneuvers, the first drive unit is coordinated and called upon to participate. This control quantity allocation strategy based on a physical model not only ensures control efficiency but also minimizes mutual interference between actuators, achieving optimal utilization of the power system's output energy.

[0025] This invention, through the innovative design of a two-dimensional vector control structure and the synergistic optimization of corresponding control methods, enables the aircraft to independently control lateral forces and yaw moments, fundamentally solving the problems of yaw instability and difficult lateral maneuvers inherent in traditional one-dimensional vector turbojet aircraft. This scheme significantly improves the aircraft's attitude stability and control accuracy during hovering, transition, and low-speed flight phases. Simultaneously, by avoiding redundant maneuvers to compensate for yaw, it effectively reduces energy loss and extends flight time. The entire system achieves simultaneous improvements in handling quality and flight efficiency while maintaining the high thrust-to-weight ratio advantage of turbojet engines.

[0026] Furthermore, the angle calculation unit calculates the desired attitude angular velocity using an outer-loop PID controller based on the error between the desired attitude angle and the current attitude angle; the angle calculation unit calculates the desired three-axis control torque using an inner-loop PID controller based on the error between the desired attitude angular velocity and the current attitude angular velocity; and the angle calculation unit calculates the desired turbojet engine deflection angle of each vector turbojet engine 2 and the desired blade deflection angle of each swaying blade 3 based on the dynamic model of the aircraft, according to the desired three-axis control torque and the desired aircraft altitude.

[0027] The angle calculation unit, through a strategy combining cascaded PID control and dynamic feedforward compensation, decomposes the complex attitude control problem into multiple cascaded single-variable adjustment problems, thereby reducing the complexity of system design while ensuring control accuracy. From the perspective of principle and improvement, this dual-closed-loop control structure embodies the idea of ​​hierarchical control. The outer loop controller focuses on position-level accuracy adjustment, mapping attitude angle errors to reasonable angular velocity commands through a proportional element, consistent with the dynamic characteristics of an aircraft as an inertial system. The inner loop controller is responsible for dynamic performance optimization, rapidly tracking the angular velocity commands output by the outer loop through a complete PID algorithm and generating the three-axis control torques required to achieve the motion. Most importantly, the system does not stop at the torque command level but further integrates a precise dynamic model of the aircraft. By solving the force / torque balance equations, the abstract torque commands are converted into eight specific control quantities: four engine deflection angles and four jet blade deflection angles. This model-based control allocation method fully considers the unique control coupling characteristics of a two-dimensional vector thrust system, ensuring the physical realizability of the control commands.

[0028] Furthermore, the desired blade deflection angle The calculation formula is: ; ; ; ; ; in, To achieve the desired roll control torque, To achieve the desired pitch control torque, To achieve the desired yaw control torque, For the desired lift in the Z direction, The thrust of a single turbojet engine.

[0029] The above calculation formula is used to accurately translate high-level control commands into low-level actuator actions. Through a control allocation strategy based on thrust vector synthesis, the desired resultant force and resultant torque calculated by the flight control system are decomposed and mapped onto the blade deflection angles of the four turbojet engines, thereby achieving six-degree-of-freedom precise control of the spacecraft's spatial motion.

[0030] In the above formula, The parameters can be viewed as the component requirements of the combined thrust command allocated to each engine in the 1Z-axis direction of the airframe. Four The calculation formula exhibits a regular symmetry in its structure, a design intended to achieve differential control of roll and pitch moments. For example, when a positive roll moment is required, the system increases the power of the left engine (…). and The Z-axis thrust component of the engine is reduced, while the thrust of the right engine is decreased. and The corresponding components of ) thus form a torque about the X-axis. Subsequently, the inverse cosine function is used to solve for... Its physical meaning lies in the Z-axis thrust component required by the engine ( The ratio of thrust to total thrust (T) is used to deduce the angle between the thrust vector and the Z-axis of the aircraft to achieve this thrust decomposition. In one embodiment, if the aircraft needs to maintain an altitude ( While maintaining stability, accelerate pitch ( (increase), then the two engines located at the front... The value will increase, and the rear engine's As the value decreases, the calculated δ angle will cause the thrust of the forward engine to be more horizontal and the thrust of the rear engine to be more vertical, thus synthesizing the required pitching moment. At the same time, lift balance is maintained by the sum of the Z-axis components of the four engines. This calculation process ensures that control commands can be accurately and conflict-free distributed to each actuator in complex coupled dynamic environments.

[0031] Desired turbojet engine deflection angle The calculation formula is: The above formula is used to generate and control the yaw moment of the aircraft, and is one of the core components of realizing two-dimensional vector control. Its core function is to drive the deflection of the yaw blades 3 in a specific group, so that the engine jet produces a controllable lateral force, thereby forming a pure torque around the vertical axis of the fuselage 1, realizing independent and precise control of the yaw channel.

[0032] When engines in symmetrical positions generate lateral forces in opposite directions, they primarily create a yaw moment around the Z-axis, with minimal impact on the translational motion of the aircraft's center of mass. The calculation formula is based on the desired yaw moment ( The sign (positive or negative) of the yaw torque divides the four engines into two logical groups. For example, when a positive yaw torque is required ( When the angle is >0, the system only deflects the yaw blades 3 of engines 1 and 3, causing their jets to generate a lateral force, while keeping the blade angles of engines 2 and 4 at zero. Since engines 1 and 3 are diagonally symmetrical in the fuselage layout, the lateral forces they generate are in the same direction, but due to the lever arm, they together form a pure yaw moment around the Z-axis. The inverse cosine function acos( The function of / T is to determine the magnitude of the required torque ( The thrust (T) of a single engine is used to precisely calculate the angle of deflection required by the swivel nozzle 3 to achieve that torque. In one specific embodiment, if the aircraft needs to make a left turn in place (corresponding to a negative...) If the control system activates the deflection function of the 3rd deflection blades of the 2nd and 4th engines according to the above formula, it will deflect the jet stream in a specific direction to generate lateral force, thereby efficiently achieving yaw maneuvering without changing the overall thrust or main thrust direction of the engine, thus minimizing coupling interference to the other channels.

[0033] Furthermore, the formula for calculating the desired three-axis control torque by the inner-loop PID controller is as follows: ; ; ; in, For roll rate error, This is the roll angle acceleration error. For pitch angular velocity error, For pitch angle acceleration error, For yaw rate error, This refers to the yaw angle acceleration error. , , These are the proportional coefficient, integral coefficient, and derivative coefficient of the inner loop controller, respectively.

[0034] The inner-loop control torque calculation formula constitutes the core decision-making link for flight attitude stability and control in the overall technical solution. Its function is to convert the desired angular velocity command output by the outer-loop controller into the three-axis control torque required to achieve that motion state on the aircraft body 1. This conversion process is directly related to the dynamic response characteristics of the aircraft and is a key step to ensure that the aircraft can quickly, accurately, and smoothly track attitude commands.

[0035] The inner-loop PID controller design embodies deep compensation and optimization of the aircraft's rotational dynamics. The proportional term generates control torque based on the current angular velocity error, and its function is to rapidly reduce angular velocity deviation. The integral term, through the continuous accumulation of angular velocity error, is specifically used to eliminate steady-state angular velocity error caused by constant external disturbances or system asymmetry, ensuring that the aircraft can maintain a constant rotational speed when encountering disturbances such as crosswinds. The derivative term innovatively introduces angular acceleration error as a feedback signal to sense the inertial trend of rotational motion, thereby providing advanced damping torque before drastic changes in angular velocity. This effectively suppresses the angular velocity oscillation phenomenon that easily occurs when the aircraft performs rapid maneuvers or is disturbed, significantly improving the smoothness and stability of attitude control.

[0036] In one specific embodiment, when the aircraft performs a roll maneuver, if the sensor detects a rapid increase in roll angular acceleration, the differential term immediately generates a counteracting torque to prevent the roll angular velocity from exceeding the target value. This avoids repeated oscillations around the desired attitude, achieving a rapid and overshoot-free stabilization process. This direct sensing and compensation of angular acceleration reflects the dynamic changes of the system more accurately than traditional angular velocity differentiation, thus providing more effective damping control.

[0037] This inner-loop control torque calculation strategy integrates proportional, integral, and derivative (based on angular acceleration) feedback of angular velocity error to construct a fast-responding, zero-steady-state-error, and excellent-damping closed-loop angular velocity control system. This enables the aircraft to exhibit superior attitude tracking capability and dynamic stability under various flight conditions and external disturbances, providing a reliable dynamic basis for the precise execution of upper-level flight control tasks.

[0038] Z-direction desired thrust The calculation formula is: ; in, The error between the aircraft's desired altitude and its current altitude. This represents the integral coefficient of the altitude controller.

[0039] The above calculation formula generates the total lift command required to maintain or change the aircraft's vertical position by continuously integrating the altitude error. This pure integral control strategy aims to eliminate steady-state errors during altitude holding, ensuring that the aircraft can accurately stabilize at the desired reference altitude.

[0040] This altitude controller employs a dedicated design for the characteristics of the vertical channel. Any persistent altitude error is continuously accumulated in the integrator, thereby continuously adjusting the Z-axis thrust command. For example, when the actual altitude of the aircraft is lower than the desired value, a positive altitude error will be integrated to... The command is continuously increased, instructing the propulsion system to increase total lift, causing the aircraft to ascend until the error is eliminated; conversely, the command decreases. The static equilibrium point of the aircraft in the vertical direction is determined by the condition that total lift equals gravity. Integral control fundamentally eliminates steady-state error by automatically finding the thrust value that can accurately balance gravity. In a specific embodiment, when the aircraft is hovering and its mass decreases slightly due to fuel consumption, or its altitude decreases slowly due to downdraft disturbances, traditional proportional control may not be able to fully compensate for this slowly changing deviation, resulting in a persistently low altitude. The integral controller in this design can automatically and progressively increase the thrust command until the aircraft returns to and stabilizes at the preset altitude, achieving zero steady-state error altitude control. This design decouples altitude control from the traditional method of indirect adjustment based on pitch attitude, realizing independent and direct control of the vertical channel.

[0041] Furthermore, the formula for calculating the desired attitude angular velocity obtained by the outer-loop PID controller is as follows: ; ; ; in, These represent the errors between the expected and current values ​​of the roll, pitch, and yaw angles, respectively. This is the proportional coefficient of the outer loop PID controller.

[0042] The aforementioned outer-loop PID controller constitutes the first-level decision-making link of attitude control, converting the aircraft's attitude angle error into the desired angular velocity command required to achieve attitude correction. This conversion process establishes a bridge between static attitude deviation and dynamic rotational motion, providing a clear tracking target for inner-loop angular velocity control.

[0043] This outer-loop controller, based on the dynamic characteristics of the aircraft as the controlled object in the angular velocity control system, sets a linear relationship between attitude angle error and desired angular velocity. When an attitude angle deviation is detected, the controller immediately generates an angular velocity command proportional to it, directed to reduce the deviation, with its magnitude determined by a proportionality coefficient. Decision. When a change in flight attitude is needed, a corresponding rotational angular velocity must first be applied. For example, when the aircraft needs to transition from a horizontal attitude to a turn with a certain bank angle, the outer loop controller calculates a desired roll angular velocity based on the deviation between the current roll angle and the target roll angle. Proportional coefficient. The value of determines the system's sensitivity to attitude deviations and its response speed; a higher value indicates a higher sensitivity. A high value will make the system respond more quickly to attitude deviations, but an excessively high value may increase the burden on the inner loop control or even cause oscillations. This pure proportional outer loop design, combined with the complete PID control of the inner loop, forms a dual closed-loop structure with clearly defined responsibilities: the outer loop focuses on determining "how fast the angular velocity should be" to correct the attitude, while the inner loop is responsible for precisely realizing "how to generate a specific torque" to achieve that angular velocity.

[0044] Furthermore, the dynamic model of the aircraft includes force balance equations and moment balance equations; The formula for calculating the force balance equation is: ; The formula for calculating the moment balance equation is: ; in, For the mass of the aircraft, Let be the position vector of the aircraft in the inertial frame. Let be the rotation matrix from the body coordinate system to the inertial coordinate system. The resultant force along the three axes in the body coordinate system. The gravity vector This is the drag vector. The moment of inertia of the aircraft. The angular acceleration vector in the body coordinate system. The three-axis control torque in the body coordinate system. This is the angular velocity vector in the body coordinate system. This is the damping torque vector.

[0045] This dynamic model forms the theoretical foundation and mathematical basis for the control algorithm design in the overall technical solution, accurately describing the six-degree-of-freedom motion of the aircraft under the coupled effects of multiple physical fields such as thrust, gravity, and aerodynamic forces. By establishing a complete set of dynamic equations, the control system can accurately predict the motion evolution process based on the current state of the aircraft and the actuator inputs, thereby achieving a precise mapping from control commands to actuator actions.

[0046] The dynamic model comprehensively considers the translational and rotational dynamics of the aircraft in three-dimensional space. The force balance equations establish the mathematical relationship between the thrust output in the body coordinate system and the translational acceleration of the aircraft in the inertial frame. The rotation matrix realizes the vector transformation from the body coordinate system to the inertial coordinate system, ensuring that the thrust generated by the engine can be correctly projected into the geographic coordinate system, which, together with gravity and wind drag, determines the trajectory change of the aircraft. The moment balance equations describe the rotational dynamics of the aircraft in a more complex way. The Coriolis force and centripetal force terms accurately characterize the gyroscopic effect and inertial coupling phenomenon generated during the rotation of the aircraft, which is particularly crucial for the stability analysis of vector thrust aircraft during rapid maneuvers.

[0047] In one specific embodiment, when the aircraft performs a roll maneuver, the dynamic model can accurately calculate the cross-inertial torque generated by the coupling of pitch and yaw angular velocities, enabling the control system to compensate for this coupling interference in advance and maintain the purity and stability of the maneuver. This model not only considers the balance between control torque and inertial torque but also introduces an aerodynamic damping torque term, which allows the control system to better adapt to changes in dynamic characteristics at different flight speeds.

[0048] The establishment of a complete dynamic model provides the control system with high-precision feedforward compensation and state prediction capabilities, enabling the control system to fully understand and compensate for the nonlinear coupling characteristics and dynamic inertial effects of the aircraft during complex maneuvers. This significantly improves the tracking accuracy and anti-interference capability of the control system, providing a solid theoretical guarantee for the stable control of the aircraft throughout its flight envelope.

[0049] Furthermore, the resultant force of the three axes in the body coordinate system The calculation formula is: ; in, The thrust of a single turbojet engine. Let i be the overall deflection angle of the i-th engine. Let be the deflection angle of the i-th engine's swivel spray blade 3.

[0050] The resultant force calculation formula is used in the overall technical scheme to synthesize the mechanical effects of eight independent control variables (four engine deflection angles and four swivel nozzle deflection angles) into net thrust in three directions in the body coordinate system. It establishes a quantitative relationship between actuator action and the total external force on the aircraft, and is the basis for analyzing the translational motion and control distribution of the aircraft.

[0051] This formula precisely describes the unique mechanical characteristics of a two-dimensional vector thrust system. The three components in the formula correspond to the forward force, lateral force, and vertical force in the airframe coordinate system, respectively: the forward force component is mainly determined by the engine deflection angle. The sine function and the deflection angle of the oscillating spray blades The cosine function of the thrust is jointly determined, reflecting the projection of the thrust in the X-axis direction; the lateral force component is entirely determined by the deflection angle of the swivel nozzle blades. The sine function determines the lateral control mechanism, which is the core mechanism for achieving lateral control in this scheme. Its sign distribution ensures that the symmetrical engines can produce a differential effect. The vertical force component is formed by the combined action of the cosine functions of the two deflection angles, constituting the main source of lift for the aircraft. In a specific embodiment, when the aircraft needs to translate to the right, the control system adjusts the deflection angle of the thrust vanes of each engine. While maintaining a relatively constant total lift, the right engine generates a lateral force to the left, and the left engine generates a lateral force to the right. The net effect of these lateral forces in the Y-axis direction is the lateral thrust that propels the aircraft to the right. The forward and vertical forces are then... and The coordinated configuration maintains balance. This precise mathematical description allows the control system to be solved in reverse, that is, to deduce the deflection state of each actuator from the desired resultant force of the three axes.

[0052] Furthermore, the three-axis control torque in the body coordinate system The calculation formula is: ; in, Let Y be the lever arm of the engine relative to the aircraft's center of mass. The force arm of the engine relative to the center of mass of the aircraft is the X-axis direction.

[0053] The formula for calculating control torque establishes a precise mathematical relationship between actuator state and aircraft rotational dynamics, transforming the spatial force effects of eight control variables into rotational torques around the three axes of the aircraft body. This provides a key mechanical model foundation for achieving three-axis attitude stability and maneuver control of the aircraft, accurately describes the spatial torque distribution characteristics generated by the two-dimensional vector thrust system, and is the core basis for the inverse solution from desired torque to actuator command in the control allocation algorithm.

[0054] The torque calculation formula fully considers the lever arm effect generated by the engine spatial layout and the complex coupling characteristics of the two-dimensional thrust vector. The three components in the formula correspond to roll, pitch, and yaw moments, respectively: the roll moment component mainly reflects the differential effect of the thrust components of the left and right symmetrical engines in the X-axis direction, and its sign distribution law ensures that the asymmetry of the thrust of the left and right engines can generate effective roll control; the pitch moment component reflects the differential effect of the thrust components of the front and rear symmetrical engines in the Z-axis direction, and pitch control is achieved by coordinating the thrust vector directions of the front and rear engines; the most complex yaw moment component consists of two parts. The first part comes from the special differential combination of the thrust components of the front and rear engines in the Y-axis direction, and the second part is directly formed by the lateral thrust differential generated by the swivel nozzle 3. This composite structure is the key physical mechanism for realizing independent yaw control. In one specific embodiment, when the aircraft needs to achieve pure yaw motion, the control system adjusts the deflection angle of the diagonally positioned engine blades 3 to generate specific lateral forces. These lateral forces, under the action of the lever arm, form a pure yaw moment about the Z-axis without generating additional roll or pitch moment interference. This precise moment modeling allows the control system to accurately predict the full impact of each actuator's actions on the aircraft's rotational motion.

[0055] Furthermore, the wind resistance vector The calculation model is as follows: ; in, Here is the damping coefficient matrix. ξ ˙ represents the velocity vector of the aircraft in the inertial coordinate system. These are the damping coefficients in the three axial directions, These represent the velocity components of the aircraft in the three axes.

[0056] The drag model characterizes the aerodynamic drag effect experienced by an aircraft when it moves in the air. Its core function is to provide the control system with key dynamic factors of the interaction between the aircraft and the environment, enabling the control system to accurately predict and compensate for the impact of aerodynamic drag on the aircraft's motion state, thereby maintaining control accuracy under various flight conditions.

[0057] This linear drag model employs a velocity-dependent aerodynamic modeling method, where the diagonal structure of the damping coefficient matrix reflects the decoupling characteristics of aerodynamic drag in the three main axes. The model describes the aerodynamic drag characteristics of the aircraft in the three principal axes using three independent velocity-damping coefficients. The drag magnitude in each direction is proportional to the flight velocity in that direction, and the direction is opposite to the velocity direction. In a specific embodiment, as the forward velocity of the aircraft increases, the forward aerodynamic drag calculated by this model will increase linearly. At this time, the control system needs to correspondingly increase the forward thrust component to maintain the flight velocity. When encountering crosswinds, the aircraft will generate relative airflow velocity laterally. This model can accurately calculate the resulting lateral aerodynamic drag, allowing the control system to adjust the angle of the thruster blades in a timely manner to generate sufficient lateral thrust to counteract the wind disturbance. This velocity-based linear drag modeling method has sufficient accuracy within the normal flight speed range of the aircraft while maintaining computational efficiency, making it suitable for the requirements of real-time control systems. Furthermore, the damping coefficients in different directions can be determined separately based on the actual aerodynamic shape characteristics of the aircraft, which allows the model to more accurately reflect the differences in aerodynamic characteristics of the aircraft in different directions.

[0058] In summary, this invention constructs a two-dimensional vector control hardware architecture by adding a swivel blade 3 at each turbojet engine nozzle 4, which is based on the coordinated action of the overall engine deflection and the airflow deflection at the nozzle 4. On this basis, a complete flight control model system is established. This system includes an outer-loop control law that generates the desired angular velocity command based on attitude angle error, an inner-loop PID control law that calculates the desired three-axis control torque based on angular velocity and angular acceleration errors, an inverse algorithm that maps the desired control torque and altitude hold command to the deflection angles of each engine and the swivel blade 3, and a dynamic model that accurately describes the space motion of the aircraft under the coupling of thrust, gravity and aerodynamics. The resultant force and torque model quantifies the contribution of eight control variables to the three-axis force / torque, while the linear drag model provides adaptive compensation capability for the external aerodynamic environment. The entire scheme achieves precise decoupling and stable control of the six-degree-of-freedom motion of the aircraft through the deep synergy of hardware innovation and control algorithms.

[0059] Accordingly, please refer to Figure 4 The second aspect of this invention provides a control method for a two-dimensional vector turbojet aircraft, used to control a two-dimensional vector turbojet aircraft, the control method comprising the following steps: Step S100: Obtain the current attitude angle, current attitude angular velocity and desired attitude angle of the aircraft in the current control cycle, and calculate the attitude angle error between the current attitude angle and the desired attitude angle.

[0060] At the beginning of the flight control cycle, the control system integrates data from multiple sensors, including the inertial measurement unit, global positioning system, and barometers, to acquire real-time measured values ​​of the aircraft's three-axis attitude angles, including roll, pitch, and yaw, within the current control cycle. This step involves vector subtraction between the current attitude angles calculated by the navigation system and the desired attitude angles issued by the flight management computer to obtain the attitude angle errors in the three axes. For example, when the aircraft needs to maintain a specific heading during inter-island transport missions, the system continuously calculates the deviation between the current yaw angle and the target heading angle; when navigating between urban buildings, it simultaneously calculates the attitude errors in the roll and pitch channels to ensure the aircraft smoothly passes through narrow spaces. This error calculation process provides the most basic input signal for subsequent attitude correction control, and its accuracy directly determines the steady-state performance of the entire control system.

[0061] Step S200: Calculate the desired attitude angular velocity based on the attitude angle error.

[0062] Based on the attitude angle error obtained in the previous step, the outer-loop controller uses a proportional control law to convert it into a corresponding desired angular velocity command. This control loop simulates the pilot's operational logic—larger attitude deviations require faster rotation rates for correction. During the transition from hovering to forward flight, when a pitch angle deviation from the target value is detected, the control system generates a pitch angular velocity command proportional to it, the magnitude of which is precisely adjusted by the outer-loop proportional coefficient. This design allows the aircraft to adaptively adjust its rotation speed in the face of attitude deviations of varying magnitudes, ensuring smooth adjustment under small deviations and rapid response under large deviations. This desired angular velocity command serves as the setpoint for the inner-loop control, laying the foundation for the aircraft's dynamic response characteristics.

[0063] Step S300: Calculate the attitude angular velocity error based on the current attitude angular velocity and the desired attitude angular velocity.

[0064] The three-axis angular velocity error is obtained by comparing the measured rotational rate of the aircraft 1 from the angular velocity sensor with the desired angular velocity generated in the previous step. When the aircraft is flying in a turbulent environment, this detection mechanism can sensitively capture non-command angular velocity changes caused by external disturbances; during coordinated turning maneuvers, it ensures that the roll angular velocity accurately tracks the command value. Angular velocity error reflects the instantaneous dynamic characteristics of the aircraft better than attitude angular error, providing crucial control information for the inner-loop controller, enabling the system to implement closed-loop control of the aircraft's rotational motion.

[0065] Step S400: Calculate the desired three-axis control torque based on the attitude angular velocity error.

[0066] The inner-loop PID controller, based on the angular velocity error signal, generates the three-axis control torque required to achieve the desired rotational motion through proportional-integral-derivative operations. The proportional term generates the basic control torque based on the current error magnitude; the integral term accumulates historical errors to eliminate steady-state deviations; and the derivative term provides damping through angular acceleration feedback. When the aircraft maintains its heading under crosswind conditions, the system continuously adjusts the yaw control torque to counteract wind disturbances; during pitch maneuvers, it precisely controls the magnitude and timing of the pitch torque. This control torque calculation process transforms the rotational motion control problem into a torque distribution problem, providing explicit physical quantity commands for subsequent actuator control.

[0067] In step S500, based on the desired three-axis control torque and the desired aircraft altitude, the overall deflection angle of each vector turbojet engine 2 and the deflection angle of the corresponding swivel jet blade 3 are calculated using the aircraft dynamics model.

[0068] By solving the inverse dynamics model of the aircraft, the three-axis control torque and altitude-holding commands at higher levels are mapped to specific deflection angles of the eight actuators at the lower levels. The control allocation algorithm is based on the thrust vector synthesis principle, while also considering the engine's spatial layout and two-dimensional deflection characteristics. The optimal combination of deflection angles is obtained by solving the overdetermined equations. When pure yaw adjustment is needed in hovering, the algorithm prioritizes using the 3-way deflection of the yaw jet blades to generate yaw torque; when pitch control is needed during high-speed forward flight, it coordinates the combination of overall engine deflection and blade deflection. This step achieves a precise conversion from abstract control commands to specific actuator actions, which is a key manifestation of two-dimensional vector control capabilities.

[0069] In step S600, based on the overall deflection angle of all vector turbojet engines 2 and the deflection angle of the corresponding swivel blades 3, the movement of the vector turbojet engines 2 and their corresponding swivel blades 3 is controlled to adjust the lateral thrust and yaw moment of the aircraft.

[0070] After receiving deflection angle commands from all engines and the thruster blades 3, the control system simultaneously adjusts the position states of eight actuators via a high-bandwidth servo drive unit. Each turbojet engine deflects as a whole under hydraulic or electric servo drive, while the thruster blades 3 at each engine nozzle 4 independently adjust their deflection angle under micro-servo control. When the aircraft performs lateral translational maneuvers, the control system coordinates the deflection states of the four propulsion units to generate net lateral thrust without causing unnecessary rotational torque; during altitude adjustments, it synchronously changes the thrust vector direction of all engines to adjust the total lift. This collaborative work of multiple actuators ensures maximized control efficiency of the two-dimensional vector thrust system, enabling the aircraft to precisely achieve six-degree-of-freedom spatial maneuvers.

[0071] This control method, through a hierarchical control architecture and precise model calculation, achieves stable control and precise maneuvering of a two-dimensional vector turbojet aircraft within its entire flight envelope. It significantly improves the aircraft's adaptability and mission performance in complex environments, providing a reliable technical approach for the practical development of vertical takeoff and landing aircraft.

[0072] The embodiments of the present invention aim to protect a two-dimensional vector turbojet aircraft and its control method, and have the following effects: 1. By adding sway blades to each turbojet engine nozzle, a two-dimensional vector control structure was constructed, which is jointly generated by the overall deflection of the engine and the deflection of the nozzle blades. This fundamental structural innovation breaks through the limitation of traditional one-dimensional vector control, which can only provide pitch and roll control. For the first time, it introduces the ability to independently generate lateral thrust and yaw control torque for aircraft. This enables the aircraft to directly and accurately control yaw attitude and lateral translation when hovering and maneuvering at low speeds, without having to rely on inefficient and unstable pitch-roll coupling for indirect compensation. This eliminates uncontrollable yaw phenomena from a physical level and lays a solid foundation for stable flight under all operating conditions. 2. To address the newly added two-dimensional control degrees of freedom and the resulting strong coupling and nonlinear dynamic characteristics, this invention specifically designs a dual-closed-loop PID control architecture based on a precise dynamic model. This control method decouples the outer-loop attitude angle control from the inner-loop angular velocity control, and through the established mathematical model, accurately analyzes the attitude and torque commands from the higher levels into the overall deflection angle and the deflection angle of each engine's swivel nozzle blades at the lower level. This effectively overcomes the inherent defect of slow thrust response in turbojet engines, fully leverages the advantages of fast response in two-dimensional vector structures, and ensures that the aircraft can not only fly stably but also accurately and quickly track complex flight attitude commands, achieving synergistic optimization between the control system and the mechanical structure. 3. Based on the aforementioned structural and control innovations, the two-dimensional vector turbojet aircraft of this invention achieves a leap in overall performance. Since yaw and lateral motion become controllable, the aircraft avoids energy loss during maneuvers to compensate for excessive attitude, significantly improving aerodynamic efficiency and endurance. The direct control method also reduces attitude oscillations, greatly enhancing flight stability and safety. Ultimately, this solution enables simultaneous optimization of multiple key performance indicators, including takeoff and landing flexibility, maneuver accuracy, flight stability, and energy economy, providing comprehensive assurance for its reliable and efficient operation in complex application scenarios.

[0073] Those skilled in the art will understand that embodiments of this application can be provided as methods, systems, or computer program products. Therefore, this application can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, this application can take the form of a computer program product embodied on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code.

[0074] This application is described with reference to flowchart illustrations and / or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of this application. It will be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, special-purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, generate instructions for implementing the flowchart... Figure 1 One or more processes and / or boxes Figure 1 A device that provides the functions specified in one or more boxes.

[0075] These computer program instructions may also be stored in a computer-readable storage medium that can direct a computer or other programmable data processing device to function in a particular manner, such that the instructions stored in the computer-readable storage medium produce an article of manufacture including instruction means, which are implemented in a process Figure 1 One or more processes and / or boxes Figure 1 The function specified in one or more boxes.

[0076] These computer program instructions may also be loaded onto a computer or other programmable data processing equipment to cause a series of operational steps to be performed on the computer or other programmable equipment to produce a computer-implemented process, thereby providing instructions that execute on the computer or other programmable equipment for implementing the process. Figure 1 One or more processes and / or boxes Figure 1 The steps of the function specified in one or more boxes.

[0077] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and not to limit it. Although the present invention has been described in detail with reference to the above embodiments, those skilled in the art should understand that modifications or equivalent substitutions can still be made to the specific implementation of the present invention. Any modifications or equivalent substitutions that do not depart from the spirit and scope of the present invention should be covered within the scope of protection of the claims of the present invention.

Claims

1. A two-dimensional vectored turbofan aircraft, characterized in that, Comprise: A fuselage (1), a plurality of vector turbojet engines (2) located on the fuselage (1), and a two-dimensional vector control assembly, each of the vector turbojet engines (2) is provided with a swing nozzle blade (3) at the nozzle (4) of the vector turbojet engine (2); The two-dimensional vector control assembly comprises a first driving unit, a second driving unit and an angle calculation unit; The first driving unit drives each of the vector turbojet engines (2) to deflect as a whole, so as to change the thrust direction of the vector turbojet engine (2); The second driving unit drives each of the swing nozzle blades (3) to deflect independently, so as to change the airflow direction of the nozzle (4) of the vector turbojet engine (2); The angle calculation unit calculates the overall deflection angle of each of the vector turbojet engines (2) and the deflection angle of the corresponding swing nozzle blade (3) according to the actual attitude and the expected attitude of the aircraft, controls the overall deflection angle of the vector turbojet engine (2) through the first driving unit and controls the deflection angle of the swing nozzle blade (3) through the second driving unit, and adjusts the lateral thrust and the yawing moment of the aircraft.

2. The two-dimensional vector turbojet aircraft according to claim 1, wherein The angle calculation unit is configured to: Based on the error between the expected attitude angle and the current attitude angle, an expected attitude angular velocity is calculated through an outer loop PID controller; Based on the error between the expected attitude angular velocity and the current attitude angular velocity, an expected three-axis control moment is calculated through an inner loop PID controller; According to the expected three-axis control moment and the expected aircraft height value, the expected turbojet engine deflection angle of each of the vector turbojet engines (2) as a whole and the expected blade deflection angle of each of the swing nozzle blades (3) are calculated based on the dynamics model of the aircraft.

3. The two-dimensional vector turbojet aircraft according to claim 2, wherein the desired blade deflection angle The formula for calculating the desired blade deflection angle is: ； ； ； ； ； wherein, is a desired roll control moment, is a desired pitch control moment, is a desired yaw control moment, is a desired lift in the Z direction, is a thrust of a single said turbojet engine; The desired turbojet engine deflection angle The formula for calculating the desired turbojet engine deflection angle is: ； 。 4. The two-dimensional vector turbojet aircraft according to claim 3, wherein The calculation formula of the inner loop PID controller for calculating the expected three-axis control moment is: ； ； ； wherein, is a roll angular velocity error, is a roll angular acceleration error, is a pitch angular velocity error, is a pitch angular acceleration error, is a yaw angular velocity error, is a yaw angular acceleration error, , , are a proportional coefficient, an integral coefficient and a differential coefficient of the inner loop controller, respectively; The Z-direction desired thrust The calculation formula is: ； wherein, is the error between the desired altitude of the aircraft and the current altitude, is the integral coefficient of the altitude controller.

5. The two-dimensional vector turbojet aircraft according to claim 4, wherein The calculation formula of the outer loop PID controller for calculating the expected attitude angular velocity is: ； ； ； wherein, respectively, are errors between desired and current values of roll angle, pitch angle, yaw angle, is a proportional coefficient of the outer loop PID controller.

6. The two-dimensional vectored turbofan aircraft of claim 1, wherein, The dynamics model of the aircraft comprises a force balance equation and a moment balance equation; The calculation formula of the force balance equation is: ； The calculation formula of the moment balance equation is: ； wherein is the mass of the aircraft, is the position vector of the aircraft in the inertial frame, is the rotation matrix from the body frame to the inertial frame, is the three-axis resultant force in the body frame, is the gravity vector, is the wind drag vector, is the moment of inertia of the aircraft, is the angular acceleration vector in the body frame, is the three-axis control moment in the body frame, is the angular velocity vector in the body frame, is the damping moment vector.

7. The two-dimensional vector turbojet aircraft according to claim 6, wherein The three-axis resultant force under the body coordinate system The calculation formula is: ； wherein, is the thrust of the single turbojet engine, is the overall deflection angle of the i-th engine, is the deflection angle of the i-th engine swiveling vane (3).

8. The two-dimensional vector turbojet aircraft according to claim 7, wherein The three-axis control moment in the body coordinate system The calculation formula is: ； wherein, is the Y-axis direction lever arm of the engine relative to the aircraft center of mass, is the X-axis direction lever arm of the engine relative to the aircraft center of mass.

9. The two-dimensional vector turbojet aircraft according to claim 6, wherein The wind resistance vector The computational model is: ； wherein is a matrix of damping coefficients, A control method for the two-dimensional vector turbojet aircraft according to any one of claims 1-9, the control method comprising the following steps: is the velocity vector of the aircraft in the inertial coordinate system, are the damping coefficients in the three axial directions, respectively, are the velocity components of the aircraft in the three axial directions, respectively.

10. A method of controlling a two-dimensional vectored turbojet aircraft, characterized in that, In a current control period, the current attitude angle, the current attitude angular velocity and the expected attitude angle of the aircraft are obtained, and the attitude angle error between the current attitude angle and the expected attitude angle is calculated; Based on the attitude angle error, an expected attitude angular velocity is calculated; Based on the current attitude angular velocity and the expected attitude angular velocity, an attitude angular velocity error is calculated; ​ Based on the attitude angular velocity error, a desired three-axis control moment is calculated; Based on the desired three-axis control moment and the desired aircraft altitude, the overall deflection angle of each of the vectoring turbojet engines (2) and the corresponding deflection angle of the swing jet blades (3) are calculated by means of an aircraft dynamics model; Depending on the overall deflection angle of all the vectoring turbojet engines (2) and the corresponding deflection angle of the swing jet blades (3), the vectoring turbojet engines (2) and the corresponding swing jet blades (3) are controlled to adjust the lateral thrust and the yawing moment of the aircraft.

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