A Stability Design Method for Rotating Aircraft with a New Controller

Abstract

The present invention belongs to the field of control system stability design, and specifically relates to a method for designing the stability of a rotating aircraft for a new controller, the method comprising: step 1: designing a new structural controller; step 2: designing a new controller under the constraint of a disturbance isolation index. The method can quickly and stably control a rotating aircraft, and for the first time analyzes and derives the stability conditions of the new controller. Based on this, a method for designing a new controller for a rotating aircraft based on the constraint of an isolation index is proposed. By reasonably designing the control parameters of the new controller, the rotating aircraft can ensure the stable convergence of the control system under random disturbances, thereby improving the dynamic stability of the rotating aircraft during flight.

Description

Technical Field

[0001] The present invention belongs to the field of control system stability design, and specifically relates to a method for designing the stability of a rotating aircraft for a new controller, which ensures the stable convergence of the control system of the rotating aircraft under random disturbances by reasonably designing the control parameters of the new controller, thereby improving the dynamic stability of the rotating aircraft during flight. Background Art

[0002] For guided aircraft that use a rotating system, rotation reduces the number of sensor devices and thus simplifies the composition of the control system. Rotation is also conducive to weakening the attack of laser interception weapons. However, the combined effect of the rotating motion of the rotating aircraft and the lateral velocity component of the incoming flow during flight will produce the Magnus effect. In addition, there is a time delay in the response of the actuator to the control command, resulting in the actual control force lagging the control command by a certain angle, so that the control action of the aircraft in the lateral channel is coupled to the longitudinal channel, and simultaneously generates an out-of-plane torque perpendicular to the sideslip angle plane with the Magnus effect, which ultimately causes the aircraft to fly in a spiral forward (i.e., conical motion), reducing the dynamic stability of the rotating aircraft and the guidance accuracy. There are two main measures to deal with the conical motion of the rotating aircraft. The following introduces their methods and disadvantages.

[0003] 1. Active suppression of cone motion, typically by adding a pair of independent retractable wings to the front of the aircraft. Due to the high rotation speed of the rotating aircraft, the key issue of the active suppression method is whether the actuator can respond to the command in a timely and accurate manner, otherwise the cone motion will be aggravated. This solution will add an additional front-end control mechanism, which puts higher requirements on the overall design of the aircraft; in addition, the front-end control mechanism is difficult to achieve an ideal high-quality dynamic response to the control command, and the cost-effectiveness is relatively low.

[0004] 2. Based on the classic controller, the stability scheme is derived. The typical approach is to introduce common controllers such as damping loop controller / classical three-loop controller, analyze the aircraft cone motion problem caused by radar seeker radome error / strapping seeker inertial device scale factor deviation, and derive the stability conditions of the system. However, this method is no longer applicable to new controllers and the aircraft cone motion problem caused by the disturbance torque of the platform image seeker is considered. Summary of the invention

[0005] 1. Technical issues to be resolved

[0006] The technical problem to be solved by the present invention is: how to provide a rotating aircraft stability design method for a new controller.

[0007] (II) Technical solution

[0008] In order to solve the above technical problems, the present invention provides a method for designing the stability of a rotating aircraft for a new controller, the method comprising:

[0009] Step 1: Design a new structural controller;

[0010] Step 2: Design of a new controller under the disturbance isolation index constraint.

[0011] In step 1, a new structure controller is designed; the details are as follows:

[0012] The disturbance torque model of the platform image seeker is shown in formula (1);

[0013]

[0014] Among them, T G is the disturbance torque of the platform image seeker, K a is the damping torque coefficient, K b is the spring torque coefficient, is the seeker frame angle;

[0015] The novel structural controller, which is for the case of the longitudinal channel of a rotating aircraft, is different from the classical three-loop controller in that the middle loop uses the signal measured by the accelerometer;

[0016] In step 2, a new controller is designed under the constraint of disturbance isolation index;

[0017] The equation of motion of the rotating aircraft is expressed as:

[0018]

[0019] Among them, θ represents the angle between the aircraft velocity vector and the horizontal plane, ψ V Represents the angle between the projection of the aircraft velocity vector on the horizontal plane and the ground coordinate ox axis. represents the angle between the longitudinal axis of the aircraft and the horizontal plane, ψ represents the angle between the projection of the longitudinal axis of the aircraft on the horizontal plane and the ground coordinate ox axis, α represents the aircraft's angle of attack, β represents the aircraft's sideslip angle, and δ z represents the deflection angle of the aircraft's longitudinal channel control mechanism, δ y Represents the deflection angle of the aircraft's lateral channel control mechanism; the remaining variables are defined as follows:

[0020]

[0021] a ω =m′ ω qs ref l 2 / (J lat Vm ),a δ = -m′ δ qs ref l / J lat ,

[0022] ; The meanings of the symbols are as follows: m s ′ is the derivative of the static stability moment coefficient, m′ m is the derivative of the Magnus torque coefficient, m′ ω is the derivative of the damping torque coefficient, m′ δ is the derivative of the control torque coefficient, c l ′ is the lift coefficient derivative, c′ m is the derivative of the Magnus force coefficient, s ref is the reference area, l is the characteristic length, d is the lateral diameter of the aircraft, P is the thrust of the aircraft engine, q is the dynamic pressure, J lat is the lateral moment of inertia, J lon is the longitudinal moment of inertia, The angular velocity of the aircraft, V m is the aircraft speed;

[0023] By defining the complex angle α C =β+jα and δ=δ y +jδ z , we get the linearized complex differential equation of the rotating vehicle:

[0024]

[0025] The guidance instructions for the new controller rotating the aircraft considering the isolation factor can be expressed as:

[0026]

[0027] Where N is the aircraft navigation ratio, V c is the relative approach speed between the aircraft and the target, and are the angular rates of rotation of the aircraft-target line in the longitudinal channel and the lateral channel, R d is the isolation degree, and are the pitch motion angular rate of the longitudinal channel and the yaw motion angular rate of the lateral channel, respectively, and C is the distance between the accelerometer position and the center of mass of the aircraft;

[0028] The acceleration response of the longitudinal channel and the lateral channel of the rotating aircraft is expressed as and The error signal between the longitudinal channel and the transverse channel is expressed as e z and e y It is expressed as:

[0029]

[0030] For a rotating vehicle with the actuator in front of the center of mass, in order to produce a positive pitch motion, the actuator deflection must be positive when the vehicle pitches; conversely, in a yaw motion of the vehicle, a positive actuator deflection is required to produce a negative yaw motion. For a linear time-invariant system, stability is not affected by the input signal, so the input signal can be used in stability analysis. and Zero; the input command of the actuator can be described as:

[0031]

[0032] Define the complex angle δ c =δ yc +jδ zc and φ=ψ V +jθ; where δ yc Represents the input command of the lateral channel control mechanism, δ zc represents the input command of the longitudinal channel control mechanism, ψ V represents the angle between the projection of the aircraft velocity vector on the horizontal plane and the ground coordinate ox axis, θ represents the angle between the aircraft velocity vector and the horizontal plane, and j represents the imaginary part of the complex number; the input command of the rotating aircraft actuator can be expressed as:

[0033]

[0034] The relationship between the output command and the input command of the actuator is expressed as:

[0035]

[0036] Among them, k s is the actuator gain, k r is the aerodynamic gain of the actuator system, γ d It is the rotation error angle caused by the aircraft rotation and the actuator dynamics delay;

[0037] Based on the obtained complex model, the stability criterion of the complex coefficient system is used to derive the stability domain of the design parameters of the new controller, so as to ensure the final convergence of the conical motion of the rotating aircraft and achieve higher guidance accuracy.

[0038] Substituting formula (7) into formula (8), we can get:

[0039]

[0040] Substituting formula (9) into formula (3), we can get:

[0041]

[0042] Where:

[0043]

[0044]

[0045] Defining state variables Then the state equation of the rotating aircraft guidance control system considering the isolation degree is expressed as:

[0046]

[0047] From formula (13), the characteristic equation corresponding to the guidance system can be obtained as:

[0048] λ 3 +(M1+jN1)λ 2 +(M2+jN2)λ+(M3+jN3)=0 (14)

[0049] Where:

[0050]

[0051] According to the stability condition of the third-order complex coefficient differential equation, the necessary and sufficient condition for the stability of the new controller rotating aircraft guidance system is:

[0052]

[0053] Among them, M i 、N i The expression of (i=1,2,3) is given by equations (11), (12) and (15).

[0054] The method proposes a new structural controller for the problem of conical motion of a rotating aircraft caused by the interference torque of the platform image seeker. The system response overshoot of the new controller is zero, and the rotating aircraft can be controlled quickly and stably. Based on the new controller, a new controller design method for a rotating aircraft based on isolation index constraints is proposed, and the stability conditions of the new controller are analyzed and derived for the first time, thereby obtaining the design parameter stability domain of the new controller. Based on the parameter stability domain, the main design parameters of the controller can be selected quickly and accurately to ensure that the conical motion of the rotating aircraft will eventually converge, and ultimately achieve higher guidance accuracy.

[0055] (III) Beneficial effects

[0056] Compared with the prior art, the present invention improves on the shortcomings of the existing methods. For the problem of conical motion of a rotating aircraft caused by the interference torque of the platform image seeker, a new controller is used to control the rotating aircraft. The stability conditions of the new controller are analyzed and derived for the first time, and a new controller design method for a rotating aircraft based on isolation index constraints is proposed to improve the stability of the conical motion. BRIEF DESCRIPTION OF THE DRAWINGS

[0057] Figure 1 Schematic diagram of the new structure controller.

[0058] Figure 2 This is a schematic diagram of the acceleration response curve of the new controller.

[0059] Figure 3 Schematic diagram of the control surface deflection angle response curve of the new controller.

[0060] Figure 4(a) is a schematic diagram of the guidance principle of the new controller rotating aircraft.

[0061] Figure 4(b) is a schematic diagram of the new controller's rotating aircraft guidance model.

[0062] Figure 5 Schematic diagram of the parameter stability region for the new controller design (forward path gain K A for example).

[0063] Figure 6(a) and Figure 6(b) show K A1 = 0.0227. Schematic diagram of the stable state of the rotating aircraft. Among them, Figure 6 (a) is the cone motion curve, and Figure 6 (b) is the acceleration response.

[0064] Figure 7(a) and Figure 7(b) show K A2 Schematic diagram of the critical stable state of the rotating aircraft when =0.0427.

[0065] Among them, Figure 7(a) is the cone motion curve, and Figure 7(b) is the acceleration response.

[0066] Figure 8(a) and Figure 8(b) show K A3 Schematic diagram of the divergent instability state of the rotating aircraft when =0.0627.

[0067] Among them, Figure 8(a) is the cone motion curve, and Figure 8(b) is the acceleration response. DETAILED DESCRIPTION

[0068] In order to make the purpose, content, and advantages of the present invention more clear, the specific implementation methods of the present invention are further described in detail below in conjunction with the accompanying drawings and examples.

[0069] In order to solve the above technical problems, the present invention provides a method for designing the stability of a rotating aircraft for a new controller, the method comprising:

[0070] Step 1: Design a new structural controller;

[0071] Step 2: Design of a new controller under the disturbance isolation index constraint.

[0072] In step 1, a new structure controller is designed; the details are as follows:

[0073] During the flight of a rotating aircraft, the aircraft attitude disturbance will be coupled into the platform seeker, resulting in errors in the output target line of sight angular velocity information; the seeker isolation represents the seeker's ability to isolate the aircraft disturbance; for the platform image seeker, the relative rotation between the base and the seeker frame causes interference torque, which is the main factor causing the isolation problem of the platform seeker; the platform image seeker interference torque model is reasonably simplified, as shown in formula (1);

[0074]

[0075] Among them, T G is the disturbance torque of the platform image seeker, K a is the damping torque coefficient, K b is the spring torque coefficient, is the seeker frame angle;

[0076] The novel structure controller is designed for the longitudinal channel of a rotating aircraft. Figure 1 As shown; compared with the classic three-loop controller, the difference is that the middle loop uses the signal measured by the accelerometer;

[0077] Figure 1 In, a yc and a y are the controller input and output instructions respectively, and the controller design parameter K A , k g ,ω I and K DC They represent the forward path gain, inner loop feedback gain, intermediate loop feedback gain and regulation gain respectively, k ac is the feedback gain of the outer loop;

[0078] Taking a statically unstable aircraft as an example, the difference between the new controller and the classic three-loop controller is verified; Table 1 gives the dynamic coefficients of the aircraft;

[0079] Table 1 Dynamic coefficient

[0080]

[0081] The dynamic transfer function of the control mechanism is: Taking the controller open-loop cutoff frequency as a constraint, the design results are shown in Table 2;

[0082] Table 2 Design results with open-loop cutoff frequency as constraint

[0083]

[0084] Based on the design results shown in Table 2, under unit step input, Figure 2 and Figure 3 The acceleration response curves and control surface deflection angle response curves of two different structure controllers are given respectively;

[0085] Depend on Figure 2-3 It can be seen that after the introduction of the control mechanism dynamics, under the constraint of the same open-loop cut-off frequency, the overshoot of the new controller is zero, and the transition time of the two controllers is the same, so the new controller can quickly and stably control the rotating aircraft.

[0086] In step 2, a new controller is designed under the constraint of disturbance isolation index;

[0087] The equation of motion of the rotating aircraft is expressed as:

[0088]

[0089] Among them, θ represents the angle between the aircraft velocity vector and the horizontal plane, ψ V Represents the angle between the projection of the aircraft velocity vector on the horizontal plane and the ground coordinate ox axis. represents the angle between the longitudinal axis of the aircraft and the horizontal plane, ψ represents the angle between the projection of the longitudinal axis of the aircraft on the horizontal plane and the ground coordinate ox axis, α represents the aircraft's angle of attack, β represents the aircraft's sideslip angle, and δ z represents the deflection angle of the aircraft's longitudinal channel control mechanism, δ y Represents the deflection angle of the aircraft's lateral channel control mechanism; the remaining variables are defined as follows:

[0090]

[0091] a ω =m′ ω qs ref l 2 / (J lat V m ),a δ = -m′ δ qs ref l / J lat ,

[0092] ; The meanings of the symbols are as follows: m s ′ is the derivative of the static stability moment coefficient, m′ m is the derivative of the Magnus torque coefficient, m′ ω is the derivative of the damping torque coefficient, m′ δ is the derivative of the control torque coefficient, c l ′ is the lift coefficient derivative, c′ m is the derivative of the Magnus force coefficient, s ref is the reference area, l is the characteristic length, d is the lateral diameter of the aircraft, P is the thrust of the aircraft engine, q is the dynamic pressure, J lat is the lateral moment of inertia, J lon is the longitudinal moment of inertia, The angular velocity of the aircraft, V m is the aircraft speed;

[0093] By defining the complex angle α C =β+jα and δ=δ y +jδ z , we get the linearized complex differential equation of the rotating vehicle:

[0094]

[0095]

[0096] Figure 4(a) shows the principle diagram of the new controller for rotating aircraft guidance. Figure 1 The new controller model shown is related to the isolation R d After the model, Figure 4(b) shows the guidance model of the rotating aircraft with the new controller;

[0097] The guidance instructions for the new controller rotating the aircraft considering the isolation factor can be expressed as:

[0098]

[0099] Where N is the aircraft navigation ratio, V c is the relative approach speed between the aircraft and the target, and are the angular rates of rotation of the aircraft-target line in the longitudinal channel and the lateral channel, R d is the isolation degree, and are the pitch motion angular rate of the longitudinal channel and the yaw motion angular rate of the lateral channel, respectively, and C is the distance between the accelerometer position and the center of mass of the aircraft;

[0100] The acceleration response of the longitudinal channel and the lateral channel of the rotating aircraft is expressed as and The error signal between the longitudinal channel and the transverse channel is expressed as e z and e y It is expressed as:

[0101]

[0102] For a rotating vehicle with the actuator in front of the center of mass, in order to produce a positive pitch motion, the actuator deflection must be positive when the vehicle pitches; conversely, in a yaw motion of the vehicle, a positive actuator deflection is required to produce a negative yaw motion. For a linear time-invariant system, stability is not affected by the input signal, so the input signal can be used in stability analysis. and Zero; the input command of the actuator can be described as:

[0103]

[0104] Define the complex angle δ c =δ yc +jδ zc and φ=ψ V +jθ; where δ yc Represents the input command of the lateral channel control mechanism, δ zc represents the input command of the longitudinal channel control mechanism, ψ V represents the angle between the projection of the aircraft velocity vector on the horizontal plane and the ground coordinate ox axis, θ represents the angle between the aircraft velocity vector and the horizontal plane, and j represents the imaginary part of the complex number; the input command of the rotating aircraft actuator can be expressed as:

[0105]

[0106] The relationship between the output command and the input command of the actuator is expressed as:

[0107]

[0108] Among them, k s is the actuator gain, k r is the aerodynamic gain of the actuator system, γ d It is the rotation error angle caused by the aircraft rotation and the actuator dynamics delay;

[0109] Based on the obtained complex model, the stability criterion of the complex coefficient system is used to derive the stability domain of the design parameters of the new controller, so as to ensure the final convergence of the conical motion of the rotating aircraft and achieve higher guidance accuracy.

[0110] Substituting formula (7) into formula (8), we can get:

[0111]

[0112] Substituting formula (9) into formula (3), we can get:

[0113]

[0114] Where:

[0115]

[0116]

[0117] Defining state variables Then the state equation of the rotating aircraft guidance control system considering the isolation degree is expressed as:

[0118]

[0119] From formula (13), the characteristic equation corresponding to the guidance system can be obtained as:

[0120] λ 3 +(M1+jN1)λ 2 +(M2+jN2)λ+(M3+jN3)=0 (14)

[0121] Where:

[0122]

[0123] According to the stability condition of the third-order complex coefficient differential equation, the necessary and sufficient condition for the stability of the new controller rotating aircraft guidance system is:

[0124]

[0125] Among them, M i 、N i The expression of (i=1,2,3) is given by equations (11), (12) and (15).

[0126] The method proposes a new structural controller for the problem of conical motion of a rotating aircraft caused by the interference torque of the platform image seeker. The system response overshoot of the new controller is zero, and the rotating aircraft can be controlled quickly and stably. Based on the new controller, a new controller design method for a rotating aircraft based on isolation index constraints is proposed, and the stability conditions of the new controller are analyzed and derived for the first time, thereby obtaining the design parameter stability domain of the new controller. Based on the parameter stability domain, the main design parameters of the controller can be selected quickly and accurately to ensure that the conical motion of the rotating aircraft will eventually converge, and ultimately achieve higher guidance accuracy.

[0127] Example 1

[0128] The present embodiment is as follows:

[0129] Step 1: Substitute the basic parameters of the rotating aircraft into equation (16). Table 3 gives a set of rotating aircraft parameters.

[0130] Table 3 Rotating vehicle parameter settings

[0131]

[0132]

[0133] Step 2: Solve the stability conditions of the new controller and obtain the stability domain (value range) of the design parameters of the new controller.

[0134] New controller design parameters: forward path gain K A , inner loop feedback gain k g , intermediate loop feedback gain ω I and adjust the gain K DC .

[0135] For example: When the intermediate loop feedback gain of the new controller is fixed at ω I =3.05, the inner loop feedback gain is k g =0.12, 0.17, 0.22, forward path gain K A The value range of the isolation combination parameter NV c R d Changes such as Figure 5 The corresponding value range is shown in Table 4.

[0136] Table 4 New controller forward path gain K A Value range

[0137]

[0138]

[0139] Step 3: Based on the seeker isolation index requirements issued by the rotating aircraft as a whole, the forward path gain K of the new controller is developed in combination with Table 4. A The derivation process of the value ranges of other parameters of the new controller is similar.

[0140] Assume that the isolation degree of a rotating aircraft meets the index R d ≤2%, to ensure system stability, select isolation R d =2% as the design constraint for the stability of the new controller guidance system. From Table 4, we can see that when the intermediate loop feedback gain ω I =3.05, inner loop feedback gain k g =0.22, the forward path gain K of the new controller AThe value range is 0<K A <0.0427.

[0141] In order to verify the correctness of the parameter design range, the forward path gain K is taken A1 =0.0227, K A2 =0.0427, K A3 =0.0627, and the other parameters are set according to the above example. Figure 6(a) to Figure 8(b) The effects of three different forward path gains on the conical motion of the rotating aircraft are given. It can be seen that: under the influence of the initial disturbance α0 = 4°, only when K A When it is in the interval [0,0.0427], the guidance system converges to zero quickly, and the conical motion of the rotating vehicle finally converges. A Outside the interval [0,0.0427], the guidance system quickly diverges and the conical motion of the rotating vehicle eventually becomes unstable.

[0142] In summary, the derived stability constraint conditions are consistent with the simulation results, thus verifying the correctness of the stability constraint conditions.

[0143] The above is only a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the technical principles of the present invention. These improvements and modifications should also be regarded as the scope of protection of the present invention.

Claims

1. A method for designing the stability of a rotating aircraft for a new controller, characterized in that: The method comprises: Step 1: Design a new structural controller; Step 2: Design of a new controller under the disturbance isolation index constraint.

2. The method for designing the stability of a rotating aircraft for a new controller according to claim 1, characterized in that: In step 1, a new structure controller is designed; the details are as follows: The disturbance torque model of the platform image seeker is shown in formula (1); Among them, T G is the disturbance torque of the platform image seeker, K a is the damping torque coefficient, K b is the spring torque coefficient, is the seeker frame angle; The novel structural controller, which is aimed at the longitudinal channel of a rotating aircraft, is different from a classical three-loop controller in that the middle loop uses a signal measured by an accelerometer.

3. The method for designing the stability of a rotating aircraft for a new controller according to claim 1, characterized in that: In step 2, a new controller is designed under the constraint of disturbance isolation index; The equation of motion of the rotating aircraft is expressed as: Among them, θ represents the angle between the aircraft velocity vector and the horizontal plane, ψ V Represents the angle between the projection of the aircraft velocity vector on the horizontal plane and the ground coordinate ox axis. represents the angle between the longitudinal axis of the aircraft and the horizontal plane, ψ represents the angle between the projection of the longitudinal axis of the aircraft on the horizontal plane and the ground coordinate ox axis, α represents the aircraft's angle of attack, β represents the aircraft's sideslip angle, and δ z represents the deflection angle of the aircraft's longitudinal channel control mechanism, δ y Represents the deflection angle of the aircraft's lateral channel control mechanism; the remaining variables are defined as follows: a ω =m′ ω qs ref l 2 / (J lat V m ),a δ =-m′ δ qs ref l / J lat , b L =c l ′qs ref / (mV m )+P / (mV m ), The symbols have the following meanings: m s ′ is the derivative of the static stability moment coefficient, m′ m is the derivative of the Magnus torque coefficient, m′ ω is the derivative of the damping torque coefficient, m′ δ is the derivative of the control torque coefficient, c l ′ is the lift coefficient derivative, c′ m is the derivative of the Magnus force coefficient, s ref is the reference area, l is the characteristic length, d is the lateral diameter of the aircraft, P is the thrust of the aircraft engine, q is the dynamic pressure, J lat is the lateral moment of inertia, J lon is the longitudinal moment of inertia, The angular velocity of the aircraft, V m is the aircraft speed; By defining the complex angle α C =β+jα and δ=δ y +jδ z , we get the linearized complex differential equation of the rotating vehicle: The guidance instructions for the new controller rotating the aircraft considering the isolation factor can be expressed as: Where N is the aircraft navigation ratio, V c is the relative approach speed between the aircraft and the target, and are the angular rates of rotation of the aircraft-target line in the longitudinal channel and the lateral channel, R d is the isolation degree, and are the pitch motion angular rate of the longitudinal channel and the yaw motion angular rate of the lateral channel, respectively, and C is the distance between the accelerometer position and the center of mass of the aircraft; The acceleration response of the longitudinal channel and the lateral channel of the rotating aircraft is expressed as and The error signal between the longitudinal channel and the transverse channel is expressed as e z and e y It is expressed as: For a rotating vehicle with the actuator in front of the center of mass, in order to produce a positive pitch motion, the actuator deflection must be positive when the vehicle pitches; conversely, in a yaw motion of the vehicle, a positive actuator deflection is required to produce a negative yaw motion. For a linear time-invariant system, stability is not affected by the input signal, so the input signal can be used in stability analysis. and Zero; the input command of the actuator can be described as: Define the complex angle δ c =δ yc +jδ zc and φ=ψ V +jθ; where δ yc Represents the input command of the lateral channel control mechanism, δ zc represents the input command of the longitudinal channel control mechanism, ψ V represents the angle between the projection of the aircraft velocity vector on the horizontal plane and the ground coordinate ox axis, θ represents the angle between the aircraft velocity vector and the horizontal plane, and j represents the imaginary part of the complex number; the input command of the rotating aircraft actuator can be expressed as: The relationship between the output command and the input command of the actuator is expressed as: Among them, k s is the actuator gain, k r is the aerodynamic gain of the actuator system, γ d It is the rotation error angle caused by the aircraft rotation and the actuator dynamics delay; Based on the obtained complex model, the stability criterion of the complex coefficient system is used to derive the stability domain of the design parameters of the new controller, so as to ensure the final convergence of the conical motion of the rotating aircraft and achieve higher guidance accuracy. Substituting formula (7) into formula (8), we can get: Substituting formula (9) into formula (3), we can get: Where: Defining state variables Then the state equation of the rotating aircraft guidance control system considering the isolation degree is expressed as: From formula (13), the characteristic equation corresponding to the guidance system can be obtained as: l 3 +(M1+jN1)λ 2 +(M2+jN2)λ+(M3+jN3)=0 (14) Where: According to the stability condition of the third-order complex coefficient differential equation, the necessary and sufficient condition for the stability of the new controller rotating aircraft guidance system is: Among them, M i 、N i The expression of (i=1,2,3) is given by equations (11), (12) and (15).

4. The method for designing the stability of a rotating aircraft for a new controller according to claim 3, characterized in that: In formula (2), θ represents the angle between the aircraft velocity vector and the horizontal plane.

5. The method for designing the stability of a rotating aircraft for a new controller according to claim 4, characterized in that: In formula (2), ψ V Represents the angle between the projection of the aircraft velocity vector on the horizontal plane and the ground coordinate ox axis.

6. The method for designing the stability of a rotating aircraft for a new controller according to claim 5, characterized in that: In formula (2), Represents the angle between the longitudinal axis of the aircraft and the horizontal plane.

7. The method for designing the stability of a rotating aircraft for a new controller according to claim 6, characterized in that: In formula (2), ψ represents the angle between the projection of the aircraft's longitudinal axis on the horizontal plane and the ground coordinate ox axis.

8. The method for designing the stability of a rotating aircraft for a new controller according to claim 7, characterized in that: In formula (2), α represents the aircraft's angle of attack, and β represents the aircraft's sideslip angle.

9. The method for designing the stability of a rotating aircraft for a new controller according to claim 8, characterized in that: In formula (2), δ z represents the deflection angle of the aircraft's longitudinal channel control mechanism, δ y Represents the deflection angle of the aircraft's lateral channel control mechanism.

10. The method for designing the stability of a rotating aircraft for a new controller according to claim 8, characterized in that: The method proposes a new structural controller for the conical motion problem of the rotating aircraft caused by the disturbance torque of the platform image seeker. The system response overshoot of the new controller is zero, and the rotating aircraft can be controlled quickly and stably. Based on the new controller, a new controller design method for rotating aircraft based on isolation index constraint is proposed. The stability condition of the new controller is analyzed and derived for the first time, thus obtaining the stability domain of the design parameters of the new controller. Based on the parameter stability domain, the main design parameters of the controller can be selected quickly and accurately to ensure the final convergence of the conical motion of the rotating vehicle and ultimately achieve higher guidance accuracy.

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