Rapid matching method for nonlinear state of aircraft

Abstract

The invention discloses a non-linear state rapid fitting method for an aircraft, and belongs to the field of aircraft control. The implementation method comprises the following steps: constructing a two-loop attitude pilot and a three-loop attitude pilot, and discretizing to obtain a guidance instruction with a time-delay effect; according to coordinate conversion, forward deriving each aerodynamic parameter of the aircraft and an aircraft overload instruction equation; the method comprises the following steps: calculating all pneumatic data according to part of pneumatic data obtained by a blowing experiment of an aircraft, and fitting a guidance instruction with a balance attack angle, a sideslip angle and a rudder deflection angle of the aircraft to obtain a corresponding interpolation table; all aerodynamic parameters in the flight process of the aircraft are obtained according to an aircraft guidance instruction and an interpolation table in the actual situation, namely, nonlinear state rapid fitting of the aircraft is achieved, and simulation reproduction of the aircraft as a mass point in the whole guidance process is achieved according to the nonlinear state rapid fitting result. And the precision and efficiency of aircraft guidance control are improved.

Description

Technical Field

[0001] The invention belongs to the field of aircraft control and relates to a method for quickly fitting the nonlinear state of an aircraft. Background Art

[0002] As a hot topic in tactical and strategic research in the field of aerospace, aircraft guided weapons have become the focus of development of the national defense forces of major powers. In complex modern wars, with the continuous development and progress of science and technology, the diversification and complexity of battlefield weapons have deepened, and the maneuverability and reliability of weapons need to be guaranteed to a certain extent. In the process of aircraft completing strategic goals, it is very difficult to obtain various aerodynamic parameters of itself so as to effectively improve the flight mode and posture to achieve the goal. Due to the limitations of its own space and the limitations of modern technology, it cannot be equipped with various sensors to complete the measurement. Therefore, it is difficult to find the information about the aerodynamic parameters of the entire flight of the aircraft in domestic and foreign public literature to fully reproduce the entire process of the changes in various startup parameters in the completion of the aircraft's goals. In the flight process of the aircraft in a three-dimensional state, the coupling of various aerodynamic parameters is serious, so it is very difficult to directly calculate the values ​​of various aerodynamic parameters. There are few domestic and foreign literatures on this subject, and it is difficult to obtain all aerodynamic parameters to fully simulate the flight state.

[0003] In the problem of obtaining all the aerodynamic parameters of the aircraft, according to the existing literature, there is no more direct way to obtain them (see Liu Lili, Mu Rongjun, Cui Naigang. A joint estimation algorithm for the active phase state / parameters of a multi-stage ballistic missile [J]. Journal of Astronautics, 2023, 44(12): 1839-1849.). When the ballistic parameters of the aircraft are unknown and the maneuverability characteristics are complex, the estimated trajectory and landing point of the aircraft are generally obtained by a time-consuming estimation method to complete the guidance task, but it is relatively complex and difficult to achieve engineering application. Therefore, how to quickly obtain the aerodynamic parameters of the aircraft during the guidance process has become an urgent problem to be solved. Summary of the invention

[0004] With regard to existing aircraft simulation methods and live-fire tests, the measurement and solution of all aerodynamic parameters of the aircraft are not considered, and the specific changes of various aerodynamic parameters of the aircraft in the process of completing the mission cannot be fully reproduced. The main purpose of the present invention is to provide a method for rapid simulation of the nonlinear state of an aircraft, which can achieve the calculation and simulation of various nonlinear state parameters that are difficult to solve or measure during the flight of the aircraft while ensuring the authenticity of the modeling, improve the efficiency of obtaining the difficult-to-obtain parameter data of the aircraft, solve the time lag between the aircraft rudder instructions and the guidance instructions, and then solve the technical problems caused by insufficient data of certain aircraft. The present invention can improve the accuracy and efficiency of aircraft guidance control.

[0005] The technical problems include the discretization solution of overload instructions under time lag conditions and the rapid calculation of various aerodynamic parameters of controllable points by reverse difference.

[0006] The objective of the present invention is achieved through the following technical solutions:

[0007] The present invention discloses a method for fast fitting of the nonlinear state of an aircraft. Under the condition of considering the time lag of the aircraft guidance command, all aerodynamic data are solved according to part of the aerodynamic data obtained by the aircraft wind blowing experiment, and the said data are used to solve the aircraft's balanced angle of attack, sideslip angle, and rudder deflection angle. The corresponding interpolation table is obtained by fitting the guidance command with the above-mentioned aircraft's balanced angle of attack, sideslip angle, and rudder deflection angle, and in actual situations, all the aerodynamic parameters of the aircraft during flight are obtained according to the aircraft guidance command and the interpolation table. In order to quickly obtain various aerodynamic parameters of the aircraft when only the normal overload command, lateral overload command, and flight Mach number of the aircraft guidance system are obtained: by designing a three-loop autopilot and a two-loop autopilot that meet the time lag requirements, and by solving the numerical integration problem through discretization, the time lag phenomenon of the overload command being transmitted between the relevant equipment of the aircraft is solved; the various aerodynamic parameters obtained by the aircraft simulation particle test are calculated to obtain the current balanced rudder deflection angle, and then the normal force coefficient, the lateral force coefficient, and the rolling coefficient are obtained, and the corresponding interpolation table is obtained according to the above data. value table; in order to ensure that when the aerodynamic parameter coupling is difficult to accurately calculate the current real value, the various guidance instructions with time lag effect of the aircraft system are quickly calculated and matched using the aforementioned interpolation table, and then the angle of attack, sideslip angle, and balance rudder deflection angle of the guidance instruction at the current Mach number are obtained, and then the aircraft attitude angle at this time is obtained through geometric relationships, that is, the nonlinear state of the aircraft is quickly matched, and the simulation reproduction of the aircraft as a particle in the entire guidance process is realized according to the result of the rapid matching of the nonlinear state, so as to improve the accuracy and efficiency of the aircraft guidance control.

[0008] The present invention discloses a method for rapidly matching the nonlinear state of an aircraft, comprising the following steps:

[0009] Step 1: Construct a two-loop and three-loop attitude autopilot, and obtain guidance instructions with time lag effects through discretization γ.

[0010] Among them: a yc is the pitch channel overload command; k ac , k g ,ω i , A1, A2, cs are autopilot parameters; ω m is the projectile crossing frequency; μ m is the missile damping coefficient; T α is the step time; ay It is an overload signal; is the pitch angle velocity; Ballistic inclination angular velocity.

[0011] Ignoring the dynamic characteristics of the servo, angular velocity gyro, structural filter and overload meter, according to the definition The transfer function of the steering gear is taken as -1. The three-loop pilot system is disconnected at the position of the steering gear to obtain the open-loop transfer function of the three-loop pilot system.

[0012]

[0013] Introduce intermediate variables K'0, K' acω , K' gω ,make:

[0014]

[0015] Then introduce intermediate variables B1, B2, B3, let

[0016]

[0017] The open-loop transfer function of the three-loop autopilot system is simplified to:

[0018]

[0019] The closed-loop transfer function of the three-loop autopilot system is obtained:

[0020]

[0021] The closed-loop characteristic equation of the three-loop autopilot system is obtained:

[0022]

[0023] Assume that the three poles required by the pole configuration consist of a negative real root and a pair of oscillating roots, and assume that the natural frequency of the oscillating root is ω, the damping is μ, and the time constant of the first-order lag link is τ, then the expected closed-loop characteristic equation of the three-loop autopilot system is:

[0024]

[0025] The pole configuration should be completed with the above two equations being equal, thus:

[0026]

[0027] In the formula, τ is known, according to ω and ω m Calculate B3; given B3, τ, ω, ω m and μ, calculate B1 and B2. Substitute B1, B2, and B3 into the above formula to obtain the solution K'0, K'acω , K' gω The expression

[0028]

[0029] Given K'0, K' acω , K' gω , we get the autopilot design parameter k g ,ω i , k ac Calculation formula:

[0030]

[0031] According to formula (10), it is known that K'0, Get parameter k g ; Known K' gω , k g , we get the parameter ω i ; Known K' acω , ω i Get k ac ; Thus, the design parameter k is obtained g ,ω i , k ac .

[0032] According to the above-mentioned pole configuration calculation formula, the expected closed-loop characteristic equation of the three-loop autopilot system is:

[0033]

[0034] Since there is a pair of oscillating roots in equation (11), which will have an adverse effect on the system performance, the oscillating roots are discarded, the left side of the expression is taken and the Laplace inverse transform is performed to obtain the general formula:

[0035]

[0036] Where y refers to the aircraft normal phase overload signal a y Or lateral overload signal a z .

[0037] Discretize equation (12) to get the following equation:

[0038]

[0039] The above closed-loop control system is applicable to both the pitch channel and the yaw channel of the aircraft. The pitch channel overload instruction is a yc , the time lag is The yaw channel overload command is a zc , the time lag is The closed-loop control system is a three-loop autopilot system.

[0040] By deducing the closed-loop transfer function of the above closed-loop control system, it is obtained as follows:

[0041]

[0042] The solution required by the pole placement method is a pair of real roots. Assuming the damping of the pole placement function is μ and the natural frequency is ω, the desired closed-loop characteristic equation of the roll pilot system is obtained as

[0043]

[0044] Performing the inverse Laplace transform gives the general formula:

[0045]

[0046] In formula (16), y refers to the aircraft roll command γ.

[0047] The above closed-loop control system is a continuous system, and it is difficult to solve the required value in actual situations. Therefore, the closed-loop control system is discretized by a discretization method, and the second-order Runge-Kutta method is used to solve the above closed-loop characteristic equation (15)(16). The specific steps are as follows:

[0048] Introducing a new variable in equation (16) Transform the original equation into a system of two first-order differential equations:

[0049]

[0050] ① Prediction: Calculate at t n The predicted values ​​k1 and l1 at:

[0051]

[0052] ②Update: Use k1 and l1 to calculate temporary update values ​​of y and v and

[0053]

[0054] ③ Slope: Calculate new slopes k2 and l2 based on the temporary update values:

[0055]

[0056] ④Correction: Use k1, k2, l1, l2 to calculate the correction value y of y and v n+1 and v n+1 :

[0057]

[0058] ⑤Iteration: Update time t n+1 =t n +h, and repeat the above steps.

[0059] Among them, h is the simulation step, t is the simulation time, and k1 and k2 are process quantities.

[0060] According to equations (17) to (21), the continuous system is converted into a discrete system, which is convenient for numerical solution in the aircraft guidance process, and the second-order Runge-Kutta method is used to improve the accuracy of the roll command solved by the roll autopilot.

[0061] By substituting the rolling command γ into y in equations (17) to (21), we get γ * Then y in the formula n+1 , which is a rolling instruction with a time lag effect.

[0062] Step 2: Based on the coordinate transformation, forward derive the aerodynamic parameters of the aircraft and the aircraft overload command equation.

[0063] Construct the conversion relationship from the ground coordinate system to the projectile coordinate system, construct the conversion relationship from the ground coordinate system to the ballistic coordinate system; construct the conversion relationship from the velocity coordinate system to the projectile coordinate system; construct the conversion relationship from the ballistic coordinate system to the velocity coordinate system.

[0064] The Mach number and heading angle of the aircraft are measured by the accelerometer and gyroscope inside the aircraft. Combined with the aircraft kinematic equation, the equations for solving the aerodynamic parameters of the aircraft are derived as follows:

[0065]

[0066]

[0067]

[0068] Among them, α is the angle of attack, β is the sideslip angle, and γ v is the velocity tilt angle, is the pitch angle, ψ is the yaw angle, γ g is the roll angle, θ is the velocity inclination angle, θ v is the velocity deflection angle.

[0069] Combined with the flight dynamics equation, the normal force, lateral force and drag coefficient of the aircraft are expressed as follows

[0070]

[0071] Where: P is the thrust of the aircraft engine, m is the mass of the aircraft, a ycis the normal overload instruction of the aircraft, a zc is the lateral overload instruction of the aircraft, a xc It is the aircraft roll channel instruction.

[0072] Aircraft guidance commanda y 、a z , γ are the normal force, lateral force, and roll command (i.e., pitch channel overload command, yaw channel overload command, and roll channel overload command) obtained by equation (25) according to the closed-loop control system after pole configuration in step 1, i.e., the pitch channel command, yaw channel command, and roll command with time lag effect. The roll command is the pitch channel overload command, yaw channel overload command, and roll channel overload command.

[0073] The aircraft guidance instructions include normal overload instructions, lateral overload instructions and rolling channel instructions.

[0074] Step 3: According to the wind tunnel blowing data, the balanced angle of attack, balanced sideslip angle, and balanced rudder deflection angle are solved. The overload command at this time is calculated based on the aircraft overload command equation, and the solved balanced angle of attack, balanced sideslip angle, and balanced elevator deflection angle δ are calculated. z 、Balance yaw rudder deflection angle δ y 、Balance roll rudder deflection angle δ x Combined with various aerodynamic parameters to form an interpolation table.

[0075] An interpolation table is established based on the aerodynamic data of the longitudinal plane of the aircraft. The main data of the interpolation table include the angle of attack α, Mach number Ma, and pitch moment M z According to the pitch aerodynamic moment expression of the aircraft:

[0076]

[0077] Where: S is the reference area, L is the wingspan, q is the dynamic pressure, m z is the pitch moment coefficient of the aircraft, is the pitch static stability derivative, is the elevator control efficiency, δ z is the elevator deflection angle. z When =0, the pitch moment of the aircraft is 0, that is, the actual motion of the equivalent aircraft is the motion of a particle.

[0078] The variation law of the pitching moment coefficient within this range is obtained through small step iteration and multiple interpolation calculations of the Mach number, angle of attack and rudder angle.

[0079] 1) If there is m z =0, the corresponding elevator deflection angle is the trim rudder deflection angle;

[0080] 2) If mz ≠0, then take two adjacent digits And its corresponding rudder deflection angle, and then calculate the current m by Newton interpolation method z =0 corresponds to the specific value of the balanced elevator deflection angle.

[0081]

[0082] Balance yaw rudder deflection angle δ y 、Balance roll rudder deflection angle δ x , obtained by interpolation.

[0083] By substituting the wind tunnel data into equations (22) to (25) in step 2, other aerodynamic parameters are obtained, and then the balance elevator deflection angle δ z 、Balance yaw rudder deflection angle δ y 、Balance roll rudder deflection angle δ x , combined with the various aerodynamic parameters obtained in this step to form an interpolation table as follows:

[0084] (1: Normal overload command - Mach number - balance elevator deflection angle;

[0085] (2: Normal overload command - Mach number - equilibrium angle of attack;

[0086] (3: Lateral overload command - Mach number - balance rudder deflection angle;

[0087] (4: lateral overload command - Mach number - balanced sideslip angle;

[0088] (5: Normal overload command, lateral overload command, roll command - aerodynamic parameters;

[0089] The data after the wind tunnel blowing includes normal force, lateral force, thrust, drag, attitude angle, Mach number, ballistic inclination angle, and ballistic deviation angle.

[0090] The other aerodynamic parameters include angle of attack, sideslip angle, velocity inclination angle, velocity of each axis of the projectile coordinate system, normal force coefficient, lateral force coefficient, and drag coefficient.

[0091] Step 4: The guidance instructions with time lag effect for each channel obtained in step 1 γ is brought into the interpolation table formed in step 3 to quickly obtain the various parameters that are difficult to measure or calculate in the equilibrium state, and to achieve rapid matching of reverse differences. The angle of attack, sideslip angle, and balance rudder deflection angle of the guidance command at the current Mach number are obtained through the interpolation table, and the various aerodynamic parameters of the aircraft at this time are obtained through the interpolation table, that is, the nonlinear state of the aircraft is rapidly matched, and the simulation and reproduction of the aircraft as a particle in the entire guidance process is realized according to the results of the rapid matching of the nonlinear state.

[0092] The attack angle, sideslip angle, and balance rudder deflection angle of the guidance instruction at the current Mach number are obtained through interpolation tables (1, (2, (3, and (4); and the various aerodynamic parameters of the aircraft at this time are obtained through interpolation table (5). The various aerodynamic parameters include normal force, lateral force, thrust, drag, attitude angle, Mach number, trajectory inclination angle, trajectory deflection angle, attack angle, sideslip angle, velocity inclination angle, velocity of each axis of the missile body coordinate system, normal force coefficient, lateral force coefficient, and drag coefficient.

[0093] Step 5: Rapidly simulate the nonlinear state obtained in step 4 to improve the accuracy and efficiency of aircraft guidance control.

[0094] Beneficial effects:

[0095] 1. For the normal aircraft guidance process, in order to ensure that the flight trajectory of the aircraft can be regarded as a particle trajectory, there is a lag in the process of transmitting the guidance command and generating the corresponding rudder control command (that is, the balancing rudder deflection angle plus the rudder deflection angle required to adjust the aircraft attitude), which is the instant lag phenomenon. This will cause poor tracking effect, and a large number of complex calculations need to be completed when generating the rudder control command, which will consume computer memory (calculating the balancing rudder deflection angle and calculating the rudder deflection angle required to adjust the aircraft attitude are two sets of processes); secondly, in the aircraft guidance process, the angle of attack, sideslip angle, and attitude angle of the current state need to be solved (and cannot be directly measured), and this is used as the basis for solving other parameters required for guidance, so the computer memory is also consumed. Aiming at a class of aircraft that require greater control capability and reliability, the present invention discloses a method for rapid matching of aircraft nonlinear states, which performs aerodynamic analysis and solution on the wind tunnel blowing data, and uses various guidance instructions after simulation and solution processing to construct multiple interpolation tables, and constructs the problem of rapidly obtaining various aerodynamic parameters of the aircraft when only the normal overload instruction, lateral overload instruction, roll instruction, and flight Mach number of the aircraft guidance system are obtained, and the parameters of the hypersonic aircraft that are strongly coupled and nonlinear and difficult to calculate are rapidly calculated and matched, that is, the nonlinear reverse calculation matching of some important state quantities of the aircraft is realized, and the values ​​of the rudder deflection angle, angle of attack, and sideslip angle are quickly matched by interpolation. Compared with the values ​​of the rudder deflection angle, angle of attack, and sideslip angle calculated by complex calculation through guidance instructions, the present invention significantly improves the solution efficiency and reduces the consumption of computing performance; and combined with the three-loop autopilot and the two-loop autopilot designed to meet the time delay requirements, it can avoid the inaccuracy caused by the time delay of some parameters in the above-mentioned parameters using analytical methods.

[0096] 2. For the normal aircraft guidance process, all aerodynamic parameters such as speed tilt angle, attitude angle, balance rudder deflection angle, etc. are not solved during the flight process, and the change trend of all aerodynamic parameters of the aircraft during the entire flight process cannot be fully described because of memory consumption. However, all details of the entire flight state of the aircraft cannot be described macroscopically. Therefore, if there is an error after the flight test, the corresponding explanation cannot be obtained when summarizing the flight situation due to the lack of parameter review. The present invention discloses a method for rapid matching of the nonlinear state of an aircraft. Under the condition of considering the time lag of the aircraft guidance command, all aerodynamic data are solved according to part of the aerodynamic data obtained from the aircraft wind blowing experiment, and these data are used to solve the balance angle of attack, sideslip angle, and rudder deflection angle of the aircraft. The corresponding interpolation table is obtained by matching the guidance command with the above-mentioned aircraft balance angle of attack, sideslip angle, and rudder deflection angle. In actual situations, all aerodynamic parameters of the aircraft during flight are obtained according to the aircraft guidance command and the interpolation table. The present invention constructs an interpolation table to fully present detailed data of various aerodynamic parameter indicators such as attitude angle, aerodynamics, etc. in the entire guidance process of the aircraft, avoiding ignoring relevant attitude angles and related aerodynamic data due to the use of particle flight trajectories in general simulation processes, so that designers can better obtain overall or local flight information from these parameters.

[0097] 3. The present invention discloses a method for rapid simulation of the nonlinear state of an aircraft. By discretizing the time-delay instructions obtained from the continuous system of a three-loop autopilot and a roll pilot, the closed-loop characteristic equation of the three-loop autopilot is simplified to avoid complex calculations. The second-order Runge-Kutta method is used to solve the guidance instructions with high precision. The complex problem is simplified by numerical integration to achieve the solution of the nonlinear state of the aircraft. The guidance instructions with time-delay effects are solved according to the solution results, and the reverse simulation of the nonlinear state of the entire aircraft is completed based on this. According to the rapid simulation results of the nonlinear state, the accuracy and efficiency of the aircraft guidance control are improved. BRIEF DESCRIPTION OF THE DRAWINGS

[0098] Figure 1 Schematic diagram of the relationship between coordinate systems. Among them: a) Transformation diagram from ground coordinate system to missile body coordinate system; b) Schematic diagram of transformation from ground coordinate system to ballistic coordinate system; c) Transformation diagram from velocity coordinate system to missile body coordinate system; d) Schematic diagram of transformation from ballistic coordinate system to velocity coordinate system.

[0099] Figure 2 The present invention discloses a flow chart of a method for rapidly fitting a nonlinear state of an aircraft.

[0100] Figure 3 To build a three-loop autopilot.

[0101] Figure 4To build a closed-loop control loop for the rolling pilot.

[0102] Among them: a yc is the pitch channel overload command; k ac , k g ,ω i , A1, A2, cs are autopilot parameters; ω m is the projectile crossing frequency; μ m is the missile damping coefficient; T α is the step time; a y It is an overload signal; is the pitch angle velocity; Ballistic inclination angular velocity; is the rudder response parameter; J x is the X-axis moment of inertia; is the damping feedback coefficient. DETAILED DESCRIPTION

[0103] In order to better illustrate the purpose and advantages of the present invention, the invention is further described below with reference to the accompanying drawings and examples.

[0104] Embodiment 1:

[0105] like Figure 2 As shown, the method for rapid matching of the nonlinear state of an aircraft disclosed in this embodiment has the following specific implementation steps:

[0106] Step 1: Construct a two-loop and three-loop attitude autopilot, and obtain guidance instructions with time lag effects through discretization γ.

[0107] Ignoring the dynamic characteristics of the servo, angular velocity gyro, structural filter and overload meter, according to the definition The transfer function of the steering gear is taken as -1. The three-loop pilot system is disconnected at the position of the steering gear to obtain the open-loop transfer function of the three-loop pilot system.

[0108]

[0109] Introduce intermediate variables K'0, K' acω , K' gω ,make:

[0110]

[0111] Then introduce intermediate variables B1, B2, B3, let

[0112]

[0113] The open-loop transfer function of the three-loop autopilot system is simplified to:

[0114]

[0115] The closed-loop transfer function of the three-loop autopilot system is obtained:

[0116]

[0117] The closed-loop characteristic equation of the three-loop autopilot system is obtained:

[0118]

[0119] Assume that the three poles required by the pole configuration consist of a negative real root and a pair of oscillating roots, and assume that the natural frequency of the oscillating root is ω, the damping is μ, and the time constant of the first-order lag link is τ, then the expected closed-loop characteristic equation of the three-loop autopilot system is:

[0120]

[0121] The pole configuration should be completed with the above two equations being equal, thus:

[0122]

[0123] In the formula, τ is known, according to ω and ω m Calculate B3; given B3, τ, ω, ω m and μ, calculate B1 and B2. Substitute B1, B2, and B3 into the above formula to obtain the solution K'0, K' acω , K' gω The expression

[0124]

[0125] Given K'0, K' acω , K' gω , we get the autopilot design parameter k g ,ω i , k ac Calculation formula:

[0126]

[0127] According to formula (10), it is known that K'0, Get parameter k g ; Known K' gω , k g , we get the parameter ω i ; Known K' acω , ω i Get k ac ; Thus, the design parameter k is obtained g ,ωi , k ac .

[0128] According to the above-mentioned pole configuration calculation formula, the expected closed-loop characteristic equation of the three-loop autopilot system is:

[0129]

[0130] Since there is a pair of oscillating roots in equation (11), which will have an adverse effect on the system performance, the oscillating roots are discarded, the left side of the expression is taken and the Laplace inverse transform is performed to obtain the general formula:

[0131]

[0132] Where y refers to the aircraft normal phase overload signal a y Or lateral overload signal a z .

[0133] Discretize equation (12) to get the following equation:

[0134]

[0135] The above closed-loop control system is applicable to both the pitch channel and the yaw channel of the aircraft. The pitch channel overload instruction is a yc , the time lag is The yaw channel overload command is a zc , the time lag is The closed-loop control system is a three-loop autopilot system.

[0136] By deducing the closed-loop transfer function of the above closed-loop control system, it is obtained as follows:

[0137]

[0138] The solution required by the pole placement method is a pair of real roots. Assuming the damping of the pole placement function is μ and the natural frequency is ω, the desired closed-loop characteristic equation of the roll pilot system is obtained as

[0139]

[0140] Performing the inverse Laplace transform gives the general formula:

[0141]

[0142] In formula (16), y refers to the aircraft roll command γ.

[0143] The above closed-loop control system is a continuous system, and it is difficult to solve the required value in actual situations. Therefore, the discretization method is used, and the second-order Runge-Kutta method is used to solve the above closed-loop characteristic equation (15)(16). The specific steps are as follows:

[0144] Introducing a new variable in equation (16) Transform the original equation into a system of two first-order differential equations:

[0145]

[0146] ① Prediction: Calculate at t n The predicted values ​​k1 and l1 at:

[0147]

[0148] ②Update: Use k1 and l1 to calculate temporary update values ​​of y and v and

[0149]

[0150] ③ Slope: Calculate new slopes k2 and l2 based on the temporary update values:

[0151]

[0152] ④Correction: Use k1, k2, l1, l2 to calculate the correction value y of y and v n+1 and v n+1 :

[0153]

[0154] ⑤Iteration: Update time t n+1 =t n +h, and repeat the above steps.

[0155] Among them, h is the simulation step, t is the simulation time, and k1 and k2 are process quantities.

[0156] According to equations (17) to (21), the continuous system is converted into a discrete system, which is convenient for numerical solution in the aircraft guidance process, and the second-order Runge-Kutta method is used to improve the accuracy of the roll command solved by the roll autopilot.

[0157] By substituting the rolling command γ into y in equations (17) to (21), we get γ * Then y in the formula n+1 , which is a rolling instruction with a time lag effect.

[0158] Step 2: Based on the coordinate transformation, forward derive the aerodynamic parameters of the aircraft and the aircraft overload command equation.

[0159] Construct the conversion relationship from the ground coordinate system to the projectile coordinate system, construct the conversion relationship from the ground coordinate system to the ballistic coordinate system; construct the conversion relationship from the velocity coordinate system to the projectile coordinate system; construct the conversion relationship from the ballistic coordinate system to the velocity coordinate system.

[0160] The Mach number and heading angle of the aircraft are measured by the accelerometer and gyroscope inside the aircraft. Combined with the aircraft kinematic equation, the equations for solving the aerodynamic parameters of the aircraft are derived as follows:

[0161]

[0162]

[0163]

[0164] Among them, α is the angle of attack, β is the sideslip angle, and γ v is the velocity tilt angle, is the pitch angle, ψ is the yaw angle, γ g is the roll angle, θ is the velocity inclination angle, θ v is the velocity deflection angle.

[0165] Combined with the flight dynamics equation, the normal force, lateral force and drag coefficient of the aircraft are expressed as follows

[0166]

[0167] Where: P is the thrust of the aircraft engine, m is the mass of the aircraft, a yc is the normal overload instruction of the aircraft, a zc is the lateral overload instruction of the aircraft, a xc It is the aircraft roll channel instruction.

[0168] Aircraft guidance commanda y 、a z , γ are the normal force, lateral force, and roll command (i.e., pitch channel overload command, yaw channel overload command, and roll channel overload command) obtained by equation (25) according to the closed-loop control system after pole configuration in step 1, i.e., the pitch channel command, yaw channel command, and roll command with time lag effect. The roll command is the pitch channel overload command, yaw channel overload command, and roll channel overload command.

[0169] The aircraft guidance instructions include normal overload instructions, lateral overload instructions and rolling channel instructions.

[0170] Step 3: According to the wind tunnel blowing data, the balanced angle of attack, balanced sideslip angle, and balanced rudder deflection angle are solved. The overload command at this time is calculated based on the aircraft overload command equation, and the solved balanced angle of attack, balanced sideslip angle, and balanced elevator deflection angle δ are calculated. z 、Balance yaw rudder deflection angle δ y 、Balance roll rudder deflection angle δ x Combined with various aerodynamic parameters to form an interpolation table.

[0171] An interpolation table is established based on the aerodynamic data of the longitudinal plane of the aircraft. The main data of the interpolation table include the angle of attack α, Mach number Ma, and pitch moment M z According to the pitch aerodynamic moment expression of the aircraft:

[0172]

[0173] Where: S is the reference area, L is the wingspan, q is the dynamic pressure, m z is the pitch moment coefficient of the aircraft, is the pitch static stability derivative, is the elevator control efficiency, δ z is the elevator deflection angle. z When =0, the pitch moment of the aircraft is 0, that is, the actual motion of the equivalent aircraft is the motion of a particle.

[0174] The variation law of the pitching moment coefficient within this range is obtained through small step iteration and multiple interpolation calculations of the Mach number, angle of attack and rudder angle.

[0175] 1) If there is m z =0, the corresponding elevator deflection angle is the trim rudder deflection angle;

[0176] 2) If m z ≠0, then take two adjacent digits And its corresponding rudder deflection angle, and then calculate the current m by Newton interpolation method z =0 corresponds to the specific value of the balanced elevator deflection angle.

[0177]

[0178] Balance yaw rudder deflection angle δ y 、Balance roll rudder deflection angle δ x , obtained by interpolation.

[0179] By substituting the wind tunnel data into equations (22) to (25) in step 2, other aerodynamic parameters are obtained, and then the balance elevator deflection angle δ z 、Balance yaw rudder deflection angle δ y 、Balance roll rudder deflection angle δ x, combined with the various aerodynamic parameters obtained in this step to form an interpolation table as follows:

[0180] (1: Normal overload command - Mach number - balance elevator deflection angle;

[0181] (2: Normal overload command - Mach number - equilibrium angle of attack;

[0182] (3: Lateral overload command - Mach number - balance rudder deflection angle;

[0183] (4: lateral overload command - Mach number - balanced sideslip angle;

[0184] (5: Normal overload command, lateral overload command, roll command - aerodynamic parameters;

[0185] The data after the wind tunnel blowing includes normal force, lateral force, thrust, drag, attitude angle, Mach number, ballistic inclination angle, and ballistic deviation angle.

[0186] The other aerodynamic parameters include angle of attack, sideslip angle, velocity inclination angle, velocity of each axis of the projectile coordinate system, normal force coefficient, lateral force coefficient, and drag coefficient.

[0187] Step 4: The guidance instructions with time lag effect for each channel obtained in step 1 γ is brought into the interpolation table formed in step 3 to quickly obtain the various parameters that are difficult to measure or calculate in the equilibrium state, and to achieve rapid matching of reverse differences. The angle of attack, sideslip angle, and balance rudder deflection angle of the guidance command at the current Mach number are obtained through the interpolation table, and the various aerodynamic parameters of the aircraft at this time are obtained through the interpolation table, that is, the nonlinear state of the aircraft is rapidly matched, and the simulation and reproduction of the aircraft as a particle in the entire guidance process is realized according to the results of the rapid matching of the nonlinear state.

[0188] The attack angle, sideslip angle, and balance rudder deflection angle of the guidance instruction at the current Mach number are obtained through interpolation tables (1, (2, (3, and (4); and the various aerodynamic parameters of the aircraft at this time are obtained through interpolation table (5). The various aerodynamic parameters include normal force, lateral force, thrust, drag, attitude angle, Mach number, trajectory inclination angle, trajectory deflection angle, attack angle, sideslip angle, velocity inclination angle, velocity of each axis of the missile body coordinate system, normal force coefficient, lateral force coefficient, and drag coefficient.

[0189] Step 5: Rapidly simulate the nonlinear state obtained in step 4 to improve the accuracy and efficiency of aircraft guidance control.

[0190] The specific description above further illustrates the purpose, technical solutions and beneficial effects of the invention in detail. It should be understood that the above is only a specific embodiment of the present invention and is not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.

Claims

1. A method for rapid matching of aircraft nonlinear states, characterized in that: The following steps are included: Step 1: Construct a two-loop and three-loop attitude autopilot, and obtain guidance instructions with time lag effects through discretization γ; Step 2: Based on the coordinate transformation, forward derive the aerodynamic parameters of the aircraft and the aircraft overload command equation; Step 3: According to the wind tunnel blowing data, the balanced angle of attack, balanced sideslip angle, and balanced rudder deflection angle are solved. The overload command at this time is calculated based on the aircraft overload command equation, and the solved balanced angle of attack, balanced sideslip angle, and balanced elevator deflection angle δ are calculated. z 、Balance yaw rudder deflection angle δ y 、Balance roll rudder deflection angle δ x Combined with various aerodynamic parameters to form an interpolation table; Step 4: The guidance instructions with time lag effect for each channel obtained in step 1 γ is brought into the interpolation table formed in step 3 to quickly obtain the various parameters that are difficult to measure or calculate in the equilibrium state, and realize the rapid fitting of the reverse difference; the angle of attack, sideslip angle, and balance rudder deflection angle of the guidance instruction at the current Mach number are obtained through the interpolation table, and the various aerodynamic parameters of the aircraft at this time are obtained through the interpolation table, that is, the rapid fitting of the nonlinear state of the aircraft is realized, and the simulation reproduction of the aircraft as a particle in the entire guidance process is realized according to the rapid fitting result of the nonlinear state.

2. A method for rapid matching of aircraft nonlinear states according to claim 1, characterized in that: Step 1 is implemented as follows: Construct a three-loop autopilot, ignoring the dynamic characteristics of the servo, angular velocity gyro, structural filter and overload meter, that is, according to the definition The servo transfer function is taken as -1; the three-loop pilot system is disconnected at the position of the servo, and the open-loop transfer function of the three-loop pilot system is obtained; Among them: a yc is the pitch channel overload command; k ac , k g ,ω i , A1, A2, cs are autopilot parameters; ω m is the projectile crossing frequency; μ m is the missile damping coefficient; T α is the step time; a y It is an overload signal; is the pitch angle velocity; Ballistic inclination angular velocity; Introduce intermediate variables K'0, K' acω , K' gω ,make: Then introduce intermediate variables B1, B2, B3, let The open-loop transfer function of the three-loop autopilot system is simplified to: The closed-loop transfer function of the three-loop autopilot system is obtained: The closed-loop characteristic equation of the three-loop autopilot system is obtained: Assume that the three poles required by the pole configuration consist of a negative real root and a pair of oscillating roots, and assume that the natural frequency of the oscillating root is ω, the damping is μ, and the time constant of the first-order lag link is τ, then the expected closed-loop characteristic equation of the three-loop autopilot system is: The pole configuration should be completed with the above two equations being equal, thus: In the formula, τ is known, according to ω and ω m Calculate B3; given B3, τ, ω, ω m and μ, calculate B1 and B2; substitute B1, B2, B3 into the above formula to obtain the solution K'0, K' acω , K' gω The expression Given K'0, K' acω , K' gω , we get the autopilot design parameter k g ,ω i , k ac Calculation formula: According to formula (10), it is known that K'0, Get parameter k g ; Known K' gω , k g , we get the parameter ω i ; Known K' acω , ω i Get k ac ; Thus, the design parameter k is obtained g ,ω i , k ac ; According to the above-mentioned pole configuration calculation formula, the expected closed-loop characteristic equation of the three-loop autopilot system is: Since there is a pair of oscillating roots in equation (11), which will have an adverse effect on the system performance, the oscillating roots are discarded, the left side of the expression is taken and the Laplace inverse transform is performed to obtain the general formula: Where y refers to the aircraft normal phase overload signal a y Or lateral overload signal a z ; Discretize equation (12) to get the following equation: The above closed-loop control system is applicable to both the pitch channel and the yaw channel of the aircraft. The pitch channel overload instruction is a yc , the time lag is The yaw channel overload command is a zc , the time lag is The closed-loop control system is a three-loop autopilot system; Construct a closed-loop control loop of the rolling pilot, and derive the closed-loop transfer function of the closed-loop control system as follows: in, is the rudder response parameter; J x is the X-axis moment of inertia; is the damping feedback coefficient; The solution required by the pole placement method is a pair of real roots. Assuming the damping of the pole placement function is μ and the natural frequency is ω, the desired closed-loop characteristic equation of the roll pilot system is obtained as Performing the inverse Laplace transform gives the general formula: Wherein, y in formula (16) refers to the aircraft rolling command γ; The closed-loop control system is discretized by the discretization method, and the closed-loop characteristic equation (15) (16) is solved by the second-order Runge-Kutta method. The specific steps are as follows: Introducing a new variable in equation (16) Transform the original equation into a system of two first-order differential equations: ① Prediction: Calculate at t n The predicted values ​​k1 and l1 at: ②Update: Use k1 and l1 to calculate temporary update values ​​of y and v and ③ Slope: Calculate new slopes k2 and l2 based on the temporary update values: ④Correction: Use k1, k2, l1, l2 to calculate the correction value y of y and v n+1 and v n+1 : ⑤Iteration: Update time t n+1 =t n +h, and repeat the above steps; Among them, h is the simulation step, t is the simulation time, k1, k2 are process quantities; According to equations (17) to (21), the continuous system is converted into a discrete system, which is convenient for numerical solution in the aircraft guidance process, and the second-order Runge-Kutta method is used to improve the accuracy of the roll command of the roll pilot. By substituting the rolling command γ into y in equations (17) to (21), we get γ * Then y in the formula n+1 , which is the rolling instruction with time lag effect.

3. A method for rapid matching of aircraft nonlinear states as claimed in claim 2, characterized in that: Step 2 is implemented as follows: Construct the conversion relationship from the ground coordinate system to the projectile coordinate system, construct the conversion relationship from the ground coordinate system to the ballistic coordinate system; construct the conversion relationship from the velocity coordinate system to the projectile coordinate system; construct the conversion relationship from the ballistic coordinate system to the velocity coordinate system; The Mach number and heading angle of the aircraft are measured by the accelerometer and gyroscope inside the aircraft. Combined with the aircraft kinematic equation, the equations for solving the aerodynamic parameters of the aircraft are derived as follows: Among them, α is the angle of attack, β is the sideslip angle, and γ v is the velocity tilt angle, is the pitch angle, ψ is the yaw angle, γ g is the roll angle, θ is the velocity inclination angle, θ v is the velocity deflection angle; Combined with the flight dynamics equation, the normal force, lateral force and drag coefficient of the aircraft are expressed as follows Where: P is the thrust of the aircraft engine, m is the mass of the aircraft, a yc is the normal overload instruction of the aircraft, a zc is the lateral overload instruction of the aircraft, a xc It is the aircraft roll channel command; Aircraft guidance commanda y 、a z , γ are the normal force, lateral force, and roll command obtained by equation (25) according to the closed-loop control system after pole configuration in step 1, namely, the pitch channel command, yaw channel command, and roll command with time lag effect; the roll command is the pitch channel overload command, yaw channel overload command, and roll channel overload command; The aircraft guidance instructions include normal overload instructions, lateral overload instructions and rolling channel instructions.

4. A method for rapid matching of aircraft nonlinear states as claimed in claim 3, characterized in that: Step 3 is implemented as follows: An interpolation table is established based on the aerodynamic data of the longitudinal plane of the aircraft. The main data of the interpolation table include the angle of attack α, Mach number Ma, and pitch moment M z ; According to the pitch aerodynamic moment expression of the aircraft: Where: S is the reference area, L is the wingspan, q is the dynamic pressure, m z is the pitch moment coefficient of the aircraft, is the pitch static stability derivative, is the elevator control efficiency, δ z is the elevator deflection angle; if and only if m z = 0, the pitch moment of the aircraft is 0, that is, the actual motion of the equivalent aircraft is the motion of a particle; The variation law of the pitching moment coefficient within this range is obtained through small step iteration and multiple interpolation calculations of Mach number, angle of attack and rudder deflection angle. 1) If there is m z =0, the corresponding elevator deflection angle is the trim rudder deflection angle; 2) If m z ≠0, then take two adjacent digits And its corresponding rudder deflection angle, and then calculate the current m by Newton interpolation method z = 0 corresponding to the specific balanced elevator deflection angle value; Balance yaw rudder deflection angle δ y 、Balance roll rudder deflection angle δ x , obtained by interpolation; By substituting the wind tunnel data into equations (22) to (25) in step 2, other aerodynamic parameters are obtained, and then the balance elevator deflection angle δ z 、Balance yaw rudder deflection angle δ y 、Balance roll rudder deflection angle δ x , combined with the various aerodynamic parameters obtained in this step to form an interpolation table as follows: (1: Normal overload command - Mach number - balance elevator deflection angle; (2: Normal overload command - Mach number - equilibrium angle of attack; (3: Lateral overload command - Mach number - balance rudder deflection angle; (4: lateral overload command - Mach number - balance sideslip angle; (5: Normal overload command, lateral overload command, roll command - aerodynamic parameters; The data after the wind tunnel blowing includes normal force, lateral force, thrust, drag, attitude angle, Mach number, ballistic inclination angle, and ballistic deviation angle; The other aerodynamic parameters include angle of attack, sideslip angle, velocity inclination angle, velocity of each axis of the projectile coordinate system, normal force coefficient, lateral force coefficient, and drag coefficient.

5. A method for rapid matching of aircraft nonlinear states as claimed in claim 4, characterized in that: In step 4, the angle of attack, sideslip angle, and balance rudder deflection angle of the guidance instruction at the current Mach number are obtained through interpolation tables (1, (2, (3, and (4); and various aerodynamic parameters of the aircraft at this time are obtained through interpolation table (5); the various aerodynamic parameters include normal force, lateral force, thrust, drag, attitude angle, Mach number, trajectory inclination angle, trajectory deflection angle, angle of attack, sideslip angle, velocity inclination angle, axial speed of the missile body coordinate system, normal force coefficient, lateral force coefficient, and drag coefficient.

6. A method for rapidly matching the nonlinear state of an aircraft as claimed in claim 1, 2, 3, 4 or 5, characterized in that: Step 5: Rapidly simulate the nonlinear state obtained in step 4 to improve the accuracy and efficiency of aircraft guidance control.

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