A method and system for attitude control of a micro helicopter with a single main rotor and a tail rotor without a flybar

Abstract

The present invention discloses a method and system for attitude control of a micro-sized helicopter with a single main rotor and a tail rotor without aileron, and relates to the technical field of helicopter attitude control, so as to solve the attitude stability control problem of a micro-sized helicopter with a single main rotor and a tail rotor without aileron. The technical key points of the present invention include: obtaining the real-time attitude angle information and the expected attitude angle of the helicopter, wherein the real-time attitude angle information includes the helicopter attitude angle and the attitude angular velocity; processing the helicopter attitude angle to obtain the helicopter attitude angular velocity instruction; calculating the attitude angular velocity deviation according to the attitude angular velocity instruction and the attitude angular velocity, and inputting the attitude angular velocity deviation into a PID controller to obtain the helicopter three-axis control torque; converting the helicopter three-axis control torque into control signals distributed to the main rotor swash plate servo and the tail rotor servo to realize attitude control. The present invention can realize the smart attitude control of a micro-sized helicopter with a single main rotor and a tail rotor without aileron.

Description

Technical Field

[0001] The invention relates to the technical field of helicopter attitude control, and in particular to an attitude control method and system for a micro helicopter with a single main rotor and a tail rotor without a flybar. Background Art

[0002] There have been a lot of research works on helicopter flight control. At present, many control methods based on proportional-integral-differential control, linear quadratic regulator control, linear quadratic Gaussian control, H∞ robust control, explicit model tracking control, gain scheduling control, feedback linearization, sliding mode control, backstepping and model predictive control have been developed to solve the problems of helicopters in channel coupling, strong model nonlinearity and uncertainty. Among them, proportional-integral-differential control, linear quadratic regulator control, explicit model tracking control and gain scheduling control are widely used in the actual engineering of helicopter flight control. Proportional-Integral-Derivative (PID)-based helicopter control uses the deviation and its integral and differential to calculate the control signal. It does not require an accurate model of the helicopter. It has the advantages of clear physical meaning and simple structure. It is currently the most used method in the engineering implementation of helicopter flight control laws; linear quadratic regulator (PID)-based helicopter control uses the deviation and its integral and differential to calculate the control signal. It does not require an accurate model of the helicopter. It has the advantages of clear physical meaning and simple structure. It is currently the most used method in the engineering implementation of helicopter flight control laws; linear quadratic regulator (PID)-based helicopter control uses the deviation and its integral and differential to calculate the control signal. It does not require an accurate model of the helicopter. It has the advantages of clear physical meaning and simple structure. It is currently the most used method in the engineering implementation of helicopter flight control laws; linear quadratic regulator (PID)-based helicopter control uses the deviation and its integral and differential to calculate the control signal. It does not require a precise ... has the advantages of clear physical meaning and The helicopter control based on explicit model tracking control is used for the control of helicopter linear systems in the form of state-space expressions. Given a quadratic objective function, the state feedback matrix is ​​obtained by solving the Raccati algebraic equation to minimize the helicopter control objective function. The helicopter control based on explicit model tracking control obtains the desired response characteristics of each channel through explicit model design, which can effectively reduce the coupling strength between the channels of the helicopter. However, the current research only focuses on the hovering state or a single flight state and has not been verified in a larger flight envelope. The basic idea of ​​helicopter control based on gain scheduling is to linearize the dynamic model of the helicopter under different flight modes according to the typical flight mode of the helicopter, and obtain a series of linear models under different flight modes. The controller under each flight mode is designed using the design method of the linear system to obtain a series of control parameters, and the logic of switching and transition of the control parameters is designed to automatically adjust the controller parameters according to the actual flight state during the flight of the helicopter.

[0003] However, the above-mentioned helicopter control methods are basically designed for heavy manned aircraft and small and medium-sized unmanned helicopters. For micro helicopters with a single main rotor and a tail rotor without flybars, they are generally combined with micro sensors to perform tasks that require high dexterity. Since micro helicopters are small in size and light in weight, they are easily affected by environmental interference such as wind, so there are special challenges in their control. Summary of the invention

[0004] To this end, the present invention proposes a micro-helicopter attitude control method and system with a single main rotor and a tail rotor without a flap, so as to solve the attitude stability control problem of the micro-helicopter with a single main rotor and a tail rotor without a flap.

[0005] According to one aspect of the present invention, a method for attitude control of a micro helicopter with a single main rotor and a tail rotor without a flybar is provided, wherein the structure of the micro helicopter comprises a single propeller main rotor and a tail rotor with a rotor surface parallel to the longitudinal plane of the fuselage; the attitude control method comprises the following steps:

[0006] Step 1: obtaining the real-time attitude angle information and the expected attitude angle of the helicopter; the real-time attitude angle information includes the helicopter attitude angle and attitude angular velocity;

[0007] Step 2: Process the helicopter attitude angle to obtain the helicopter attitude angular velocity instruction;

[0008] Step 3: Calculate the attitude angular velocity deviation according to the attitude angular velocity instruction and the attitude angular velocity, and input the attitude angular velocity deviation into a PID controller to obtain the three-axis control torque of the helicopter;

[0009] Step 4: Convert the helicopter's three-axis control torque into control signals for the main rotor swashplate servo and tail rotor servo to achieve attitude control.

[0010] Furthermore, the specific steps of step 2 include:

[0011] Step 21: Calculate the attitude angle deviation according to the helicopter attitude angle and the expected attitude angle;

[0012] Step 22: Convert the attitude angle deviation value to between [-π,π];

[0013] Step 2 and 3: Limit the attitude angle deviation between [-π,π], and the limiting formula is:

[0014]

[0015]

[0016]

[0017] In the formula, ψ lim ,θ lim and φ lim Respectively represent the upper limits of the absolute values ​​of the yaw angle, pitch angle, and roll angle; Δψ, Δθ, Δφ represent the attitude angle deviation converted to [-π, π] respectively;

[0018] Step 24: Convert the limited attitude angle deviation into attitude angular velocity command. The conversion formula is:

[0019] p d =k φ φ e

[0020] q d =k θ θ e

[0021] r d =k ψ ψ e

[0022] In the formula, p d ,q d 、r d They represent the roll angular velocity command, pitch angular velocity command, and yaw angular velocity command, respectively, and k θ and k ψ They are the roll angle proportional coefficient, pitch angle proportional coefficient and yaw angle proportional coefficient respectively.

[0023] Furthermore, the control signal of the main rotor swash plate servo in step 4 is calculated as follows:

[0024]

[0025] In the formula, represents the control signal of the ith main rotor swashplate servo; K represents the proportional coefficient of the control torque to the pulse width modulation signal output; α i Indicates the total distance control command; γ i represents the installation phase angle of the i-th servo on the swash plate; represents the trim control command of the i-th servo; T roll ,T pitch They represent the control moments around the x-axis and y-axis in the body axis coordinate system respectively.

[0026] Furthermore, the control signal of the tail rotor servo in step 4 is calculated as follows:

[0027] δ tail =KT yaw

[0028] In the formula, δ tail Indicates the control signal of the tail rotor servo; T yaw It represents the control torque around the z-axis in the body axis coordinate system.

[0029] According to another aspect of the present invention, there is provided an attitude control system for a micro helicopter with a single main rotor and a tail rotor without a flybar, wherein the structure of the micro helicopter comprises a single propeller main rotor and a tail rotor with a rotor surface parallel to the longitudinal plane of the fuselage; the attitude control system comprises:

[0030] An attitude information acquisition module, configured to acquire real-time attitude angle information and a desired attitude angle of the helicopter; the real-time attitude angle information includes the helicopter attitude angle and attitude angular velocity;

[0031] An angular velocity instruction processing module, configured to process the helicopter attitude angle and obtain a helicopter attitude angular velocity instruction;

[0032] a control torque calculation module, configured to calculate an attitude angular velocity deviation according to the attitude angular velocity instruction and the attitude angular velocity, and input the attitude angular velocity deviation into a PID controller to obtain a three-axis control torque of the helicopter;

[0033] The attitude control module is configured to convert the helicopter's three-axis control torque into control signals distributed to the main rotor swash plate servo and the tail rotor servo to achieve attitude control.

[0034] Furthermore, the specific steps of processing the helicopter attitude angle in the angular velocity instruction processing module include:

[0035] Step 21: Calculate the attitude angle deviation according to the helicopter attitude angle and the expected attitude angle;

[0036] Step 22: Convert the attitude angle deviation value to between [-π,π];

[0037] Step 2 and 3: Limit the attitude angle deviation between [-π,π], and the limiting formula is:

[0038]

[0039]

[0040]

[0041] In the formula, ψ lim ,θ lim and φ lim Respectively represent the upper limits of the absolute values ​​of the yaw angle, pitch angle, and roll angle; Δψ, Δθ, Δφ represent the attitude angle deviation converted to [-π, π] respectively;

[0042] Step 24: Convert the limited attitude angle deviation into attitude angular velocity command. The conversion formula is:

[0043] p d =k φ φ e

[0044] q d =k θ θ e

[0045] r d =k ψ ψ e

[0046] In the formula, p d ,q d 、r d They represent the roll angular velocity command, pitch angular velocity command, and yaw angular velocity command, respectively, and k φ , k θ and k ψ They are the roll angle proportional coefficient, pitch angle proportional coefficient and yaw angle proportional coefficient respectively.

[0047] Furthermore, the control signal of the main rotor swash plate servo in the attitude control module is calculated as follows:

[0048]

[0049] In the formula, represents the control signal of the ith main rotor swashplate servo; K represents the proportional coefficient of the control torque to the pulse width modulation signal output; α i Indicates the total distance control command; γ i represents the installation phase angle of the i-th servo on the swash plate; represents the trim control command of the i-th servo; T roll ,T pitch They represent the control moments around the x-axis and y-axis in the body axis coordinate system respectively.

[0050] Furthermore, the control signal of the tail rotor servo in the attitude control module is calculated as follows:

[0051] δ tail =KT yaw

[0052] In the formula, δ tail Indicates the control signal of the tail rotor servo; T yaw It represents the control torque around the z-axis in the body axis coordinate system.

[0053] The beneficial technical effects of the present invention are:

[0054] The present invention can realize the smart attitude control of a micro helicopter with a single main propeller and a tail rotor without a flybar. The attitude angular velocity command and the control torque are calculated step by step through a hierarchical control architecture, and finally the PWM control command of the servo is obtained through torque distribution, so as to realize the attitude control of the micro helicopter with a single main propeller and a tail rotor without a flybar. Compared with the traditional method, the present invention has the following advantages: 1) a hierarchical control structure is adopted, the attitude angular velocity command is first calculated according to the attitude angle command, and then the control torque is calculated, and finally the control torque is distributed as the servo control command, and the meaning of each layer is clear, which is convenient for debugging and determining parameters separately; 2) the attitude angle control is extracted, and the dexterity of the attitude control can be adjusted by adjusting the attitude deviation to the control coefficient of the angular velocity command. BRIEF DESCRIPTION OF THE DRAWINGS

[0055] The present invention can be better understood by referring to the description given below in conjunction with the accompanying drawings, which together with the following detailed description are included in this specification and form a part of this specification, and are used to further illustrate the preferred embodiments of the present invention and explain the principles and advantages of the present invention.

[0056] Figure 1 The present invention is a flowchart of an attitude control method for a micro helicopter with a single main rotor and a tail rotor without a flap.

[0057] Figure 2 It is a three-dimensional visual image of the simulation environment in an embodiment of the present invention.

[0058] Figure 3 2 is an example diagram of a roll angle response curve in an embodiment of the present invention.

[0059] Figure 4 1 is an example diagram of a pitch angle response curve in an embodiment of the present invention.

[0060] Figure 5 1 is an example diagram of a yaw angle response curve in an embodiment of the present invention. DETAILED DESCRIPTION

[0061] In order to enable those skilled in the art to better understand the scheme of the present invention, exemplary implementations or embodiments of the present invention will be described below in conjunction with the accompanying drawings. Obviously, the described implementations or embodiments are only implementations or embodiments of a part of the present invention, not all of them. Based on the implementations or embodiments of the present invention, all other implementations or embodiments obtained by ordinary technicians in the field without creative work should fall within the scope of protection of the present invention.

[0062] Aiming at the requirements of a micro-sized helicopter with a single main rotor and a tail rotor without aileron for flight sensitivity and resistance to environmental interference when performing tasks, the present invention proposes an attitude control method for a micro-sized helicopter with a single main rotor and a tail rotor without aileron, calculates the angular velocity deviation through an estimated attitude angle value and an instruction value, and then calculates the control torque in a body axis coordinate system, and finally distributes the control torque to PWM control instructions of a main rotor swash plate servo and a tail rotor servo, so as to realize the attitude control of the micro-sized helicopter with a single main rotor and a tail rotor without aileron.

[0063] The configuration of a micro helicopter with a single main rotor and a tail rotor without ailerons includes: the main rotor is a single propeller, there is no stabilizing mechanism such as a balance bar, the root of the main rotor is connected to the cross plate through a ball joint, the cross plate is connected to at least three servo servos, the cross plate inclination is controlled by the servo servos, and then the total pitch and periodic pitch of the main rotor are controlled; the tail rotor is installed at the root of the tail rod, the rotor surface is parallel to the longitudinal plane of the fuselage, and provides a force perpendicular to the longitudinal plane of the fuselage to control the yaw angle of the fuselage. In the micro configuration, the tail rotor has a small inertia and is directly driven by a brushless motor or a hollow cup motor, and the control force is adjusted by the speed.

[0064] The attitude control method of a micro helicopter with a single main rotor and a tail rotor without a flybar is described in an embodiment of the present invention. Figure 1 As shown, the following steps are included:

[0065] Step 1: Calculate and obtain the helicopter attitude angular velocity command based on the attitude information output by the helicopter attitude estimation module and the expected attitude information input externally.

[0066] According to an embodiment of the present invention, firstly, the attitude angles ψ, θ, φ of the helicopter and the attitude angular velocities p, q, r in the body axis coordinate system output by the attitude estimation module are obtained. The attitude estimation module reads the angular velocity and acceleration information from the MPU6050 inertial measurement unit, and then uses the Kalman filter algorithm to fuse and estimate the helicopter attitude angle and attitude angular velocity; wherein ψ, θ, φ are respectively the yaw angle, pitch angle and roll angle of the body axis coordinate system relative to the local north-east coordinate system, corresponding to the rotation angles around the Z axis, Y axis and X axis, and the coordinate system conversion order is zyx; the desired attitude information input externally is the desired attitude angle ψ d ,θ d ,φ d , corresponding to the expected values ​​of attitude angles ψ, θ, φ respectively.

[0067] Then, calculate the attitude angle deviation:

[0068] Δψ=ψ d -ψ

[0069] Δθ=θ d -θ

[0070] Δφ=φ d -φ

[0071] Then, convert the attitude angle deviation value to [-π,π]:

[0072] Δψ=mod(Δψ+π,2π)-π

[0073] Δθ=mod(Δθ+π,2π)-π

[0074] Δφ=mod(Δφ+π,2π)-π

[0075] In the formula, y = mod(x,x 0 ) is the remainder function, which means x divided by x 0 The remainder is y.

[0076] Then, limit the attitude angle deviation to a certain range:

[0077] ψ e =f sat (Δψ,-ψ lim ,ψ lim )

[0078] θ e =f sat (Δθ,-θ lim ,θ lim )

[0079] φ e =f sat (Δφ,-φ lim ,φ lim )

[0080] In the formula, ψ lim ,θ lim and φ lim are the upper limits of the absolute values ​​of the yaw angle, pitch angle and roll angle, respectively, and f sat (x,x lowerlim ,x upperlim ) is the clipping function, which is in the form of:

[0081]

[0082] That is, the attitude angle deviation after limiting is expressed as:

[0083]

[0084]

[0085]

[0086] Finally, the attitude angle deviation is converted into attitude angular velocity instruction:

[0087] pd =k φ φ e

[0088] q d =k θ θ e

[0089] r d =k ψ ψ e

[0090] In the formula, k φ , k θ and k ψ They are the roll angle proportional coefficient, pitch angle proportional coefficient and yaw angle proportional coefficient respectively.

[0091] Step 2: Calculate the three-axis control torque in the helicopter body axis coordinate system according to the attitude angular velocity command and the attitude angular velocity output by the attitude estimation module.

[0092] According to an embodiment of the present invention, the specific steps of step 2 include: firstly calculating the attitude angular velocity deviation:

[0093] Δp=p d -p

[0094] Δq=q d -q

[0095] Δr=r d -r

[0096] Then, the attitude angular velocity deviation is limited to a certain range:

[0097] p e =f sat (Δp,-p lim ,p lim )

[0098] q e =f sat (Δq,-q lim ,q lim )

[0099] r e =f sat (Δr,-r lim ,r lim )

[0100] In the formula, p lim ,q lim and r lim They represent the upper limits of the absolute values ​​of the roll angular velocity, pitch angular velocity, and yaw angular velocity, respectively.

[0101] Finally, the attitude angular velocity deviation is input into the PID controller to obtain the control torque T around the x, y and z axes in the body axis coordinate system. roll ,T pitch ,T yaw .

[0102] Step 3: Substitute the control torque T calculated in step 2 roll ,T pitch ,T yaw The conversion is distributed into a pulse width modulation signal (PWM) for the main rotor swashplate servo and a PWM control signal for the tail rotor servo.

[0103] According to an embodiment of the present invention, step three specifically includes:

[0104] First, all main rotor swashplate servo control commands are calculated:

[0105]

[0106] In the formula, is the PWM control signal of the i-th main rotor swashplate servo, K is the proportional coefficient of the control torque to PWM output, α i is the collective control command, γ i is the installation phase angle of the i-th servo on the swash plate, is the trim control command for the i-th servo.

[0107] Then, all tail rotor servo control commands are calculated:

[0108] δ tail =KT yaw

[0109] In the formula, δ tail It is the tail rotor servo PWM control signal, thus completing the calculation of attitude control instructions.

[0110] The technical effect of the present invention is further verified through experiments.

[0111] The correctness and rationality of the present invention are verified by numerical simulation. First, the simulation environment and visual environment of a micro helicopter with a single main rotor and a tail rotor without flybars are constructed using the C++ programming language. Figure 2As shown in the figure, the main rotor and tail rotor in the model are combined models of finite element and blade element momentum theory to calculate the lift, power and torque of the rotor; the vertical and horizontal stabilizer models are included; the second-order servo steering model is adopted; the gravity model is WGS-84; the standard atmosphere model and wind field model are included; the fourth-order five-step Runge-Kutta integration is adopted. Initially, the helicopter is on the ground, and the main rotor and tail rotor have reached a constant speed. The main rotor swash plate is controlled by three evenly distributed swash plate servo steering gears, and the phase angles are and γ 3 =0rad, in the simulation environment, the trim control command values ​​of the three servos are all 0, that is, The simulation test software runs in Windows 10+Visual Studio 2020, and the hardware environment is Intel 9300H CPU+GTX 1660TiGPU+16.0GB RAM. Set the upper limit of the absolute value of the attitude angle deviation Attitude angle deviation proportional coefficient k φ =5, k θ =5 and k ψ =4, the upper limit of the absolute value of the attitude angular velocity deviation p lim =140rad / s,q lim =140rad / s, p lim =120rad / s.

[0112] After the simulation, the helicopter takes off, and the initial attitude angle command is 0°. After 5 seconds, the desired attitude angle command of 15° is input into the pitch channel, roll channel and yaw channel respectively to test the attitude angle control effect. The attitude angle response curves are as follows: Figure 3 , Figure 4 and Figure 5 The roll angle channel has a small oscillation in the initial response stage due to the serious channel coupling. The roll and yaw channels respond relatively quickly and can eventually stabilize the attitude at the expected value. This shows that the designed attitude control method for a micro helicopter with a single main rotor and a tail rotor without flybars has the ability to stabilize and track the attitude, in order to optimize the control effect.

[0113] Another embodiment of the present invention further provides a micro helicopter attitude control system with a single main rotor and a tail rotor without a flybar, wherein the micro helicopter includes a single propeller main rotor and a tail rotor with a rotor surface parallel to the longitudinal plane of the fuselage, and the system includes:

[0114] An attitude information acquisition module, configured to acquire real-time attitude angle information and a desired attitude angle of the helicopter; the real-time attitude angle information includes the helicopter attitude angle and attitude angular velocity;

[0115] An angular velocity instruction processing module, configured to process the helicopter attitude angle and obtain a helicopter attitude angular velocity instruction;

[0116] a control torque calculation module, configured to calculate an attitude angular velocity deviation according to the attitude angular velocity instruction and the attitude angular velocity, and input the attitude angular velocity deviation into a PID controller to obtain a three-axis control torque of the helicopter;

[0117] The attitude control module is configured to convert the helicopter's three-axis control torque into control signals distributed to the main rotor swash plate servo and the tail rotor servo to achieve attitude control.

[0118] In this embodiment, preferably, the specific steps of processing the helicopter attitude angle in the angular velocity instruction processing module include:

[0119] Step 21: Calculate the attitude angle deviation according to the helicopter attitude angle and the expected attitude angle;

[0120] Step 22: Convert the attitude angle deviation value to between [-π,π];

[0121] Step 2 and 3: Limit the attitude angle deviation between [-π,π], and the limiting formula is:

[0122]

[0123]

[0124]

[0125] In the formula, ψ lim ,θ lim and φ lim Respectively represent the upper limits of the absolute values ​​of the yaw angle, pitch angle, and roll angle; Δψ, Δθ, Δφ represent the attitude angle deviation converted to [-π, π] respectively;

[0126] Step 24: Convert the limited attitude angle deviation into attitude angular velocity command. The conversion formula is:

[0127] p d =k φ φ e

[0128] q d =k θ θ e

[0129] r d =k ψ ψ e

[0130] In the formula, p d,q d 、r d They represent the roll angular velocity command, pitch angular velocity command, and yaw angular velocity command respectively; k φ , k θ and k ψ They are the roll angle proportional coefficient, pitch angle proportional coefficient and yaw angle proportional coefficient respectively.

[0131] In this embodiment, preferably, the control signal of the main rotor swash plate servo in the attitude control module is calculated as follows:

[0132]

[0133] In the formula, represents the control signal of the i-th main rotor swashplate servo; K represents the proportional coefficient of the control torque to the pulse width modulation signal output; α i Indicates the total distance control command; γ i represents the installation phase angle of the i-th servo on the swash plate; represents the trim control command of the i-th servo; T roll ,T pitch They represent the control moments around the x-axis and y-axis in the body axis coordinate system respectively.

[0134] In this embodiment, preferably, the control signal of the tail rotor servo in the attitude control module is calculated as follows:

[0135] δ tail =KT yaw

[0136] In the formula, δ tail Indicates the control signal of the tail rotor servo; T yaw It represents the control torque around the z-axis in the body axis coordinate system.

[0137] The function of the attitude control system of a micro-helicopter with a single main rotor and a tail rotor without failrons described in this embodiment can be explained by the aforementioned attitude control method of a micro-helicopter with a single main rotor and a tail rotor without failrons. Therefore, for the parts not described in detail in this embodiment, please refer to the above method embodiments and will not be repeated here.

[0138] Although the present invention has been described according to a limited number of embodiments, it will be apparent to those skilled in the art, with the benefit of the above description, that other embodiments are contemplated within the scope of the invention thus described. The disclosure of the present invention is intended to be illustrative rather than restrictive of the scope of the invention, which is defined by the appended claims.

Claims

1. A method for attitude control of a micro helicopter with a single main rotor and a tail rotor without a flybar, wherein the micro helicopter comprises a single propeller main rotor and a tail rotor with a rotor surface parallel to the longitudinal plane of the fuselage. It is characterized in that The posture control method comprises the following steps: Step 1: obtaining the real-time attitude angle information and the expected attitude angle of the helicopter; the real-time attitude angle information includes the helicopter attitude angle and attitude angular velocity; Step 2: Processing the helicopter attitude angle to obtain the helicopter attitude angular velocity instruction; including: Step 21: Calculate the attitude angle deviation according to the helicopter attitude angle and the expected attitude angle; Step 22: Convert the attitude angle deviation value to between [-π,π]]; Step 2 and 3: Limit the attitude angle deviation between [-π,π]], and the limiting formula is: In the formula, ψ lim ,θ lim and φ lim They represent the upper limits of the absolute values ​​of the yaw angle, pitch angle, and roll angle, respectively; Δψ, Δθ, and Δφ represent the attitude angle deviations converted to [-π, π], respectively; Step 24: Convert the limited attitude angle deviation into attitude angular velocity command. The conversion formula is: p d =k φ f e q d = k θ θ e r d =k ψ ψ e In the formula, p d ,q d 、r d They represent the roll angular velocity command, pitch angular velocity command, and yaw angular velocity command, respectively, and k φ , k θ and k ψ They represent the roll angle proportional coefficient, pitch angle proportional coefficient and yaw angle proportional coefficient respectively; Step 3: Calculate the attitude angular velocity deviation according to the attitude angular velocity instruction and the attitude angular velocity, and input the attitude angular velocity deviation into a PID controller to obtain the three-axis control torque of the helicopter; Step 4: Convert the helicopter's three-axis control torque into control signals for the main rotor swash plate servo and the tail rotor servo to achieve attitude control; the control signal of the main rotor swash plate servo is calculated as follows: In the formula, represents the control signal of the ith main rotor swashplate servo; K represents the proportional coefficient of the control torque to the pulse width modulation signal output; α i Indicates the total distance control command; γ i represents the installation phase angle of the i-th servo on the swash plate; represents the trim control command of the i-th servo; T roll ,T pitch They represent the control moments around the x-axis and y-axis in the body axis coordinate system respectively.

2. The attitude control method of a micro helicopter with a single main rotor and a tail rotor without a flybar according to claim 1, It is characterized in that The control signal of the tail rotor servo in step 4 is calculated as follows: δ tail =KT yaw In the formula, δ tail Indicates the control signal of the tail rotor servo; T yaw It represents the control torque around the z-axis in the body axis coordinate system.

3. A micro helicopter attitude control system with a single main rotor and a tail rotor without a flybar, wherein the micro helicopter comprises a single propeller main rotor and a tail rotor with a rotor surface parallel to the longitudinal plane of the fuselage, It is characterized in that The attitude control system comprises: An attitude information acquisition module, configured to acquire real-time attitude angle information and a desired attitude angle of the helicopter; the real-time attitude angle information includes the helicopter attitude angle and attitude angular velocity; An angular velocity instruction processing module is configured to process the helicopter attitude angle and obtain the helicopter attitude angular velocity instruction; comprising: Step 21: Calculate the attitude angle deviation according to the helicopter attitude angle and the expected attitude angle; Step 22: Convert the attitude angle deviation value to between [-π,π]]; Step 2 and 3: Limit the attitude angle deviation between [-π,π]], and the limiting formula is: In the formula, ψ lim ,θ lim and φ lim They represent the upper limits of the absolute values ​​of the yaw angle, pitch angle, and roll angle, respectively; Δψ, Δθ, and Δφ represent the attitude angle deviations converted to [-π, π], respectively; Step 24: Convert the limited attitude angle deviation into attitude angular velocity command. The conversion formula is: p d =k φ f e q d =k θ θ e r d = k ψ ψ e In the formula, p d ,q d 、r d They represent the roll angular velocity command, pitch angular velocity command, and yaw angular velocity command respectively; k φ , k θ and k ψ They are roll angle proportional coefficient, pitch angle proportional coefficient and yaw angle proportional coefficient respectively; a control torque calculation module, configured to calculate an attitude angular velocity deviation according to the attitude angular velocity instruction and the attitude angular velocity, and input the attitude angular velocity deviation into a PID controller to obtain a three-axis control torque of the helicopter; The attitude control module is configured to convert the helicopter's three-axis control torque into control signals distributed to the main rotor swash plate servo and the tail rotor servo to achieve attitude control; the control signal of the main rotor swash plate servo is calculated as follows: In the formula, represents the control signal of the ith main rotor swashplate servo; K represents the proportional coefficient of the control torque to the pulse width modulation signal output; α i Indicates the total distance control command; γ i represents the installation phase angle of the i-th servo on the swash plate; represents the trim control command of the i-th servo; T roll ,T pitch They represent the control moments around the x-axis and y-axis in the body axis coordinate system respectively.

4. The attitude control system of a micro helicopter with a single main rotor and a tail rotor without a flybar according to claim 3, It is characterized in that The control signal of the tail rotor servo in the attitude control module is calculated as follows: δ tail =KT yaw where δ tail represents the control signal of the tail rotor servo; T yaw represents the control moment about the z-axis in the body-axis coordinate system.

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