CN115659748A - CFD technology-based aircraft attitude control law design method - Google Patents
CFD technology-based aircraft attitude control law design method Download PDFInfo
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Abstract
The invention discloses an aircraft attitude control law design method based on a CFD (computational fluid dynamics) technology, belongs to the technical field of computational fluid mechanics, flight dynamics and control, and shortens the flight test period and reduces the cost. The static and dynamic aerodynamic characteristics of the aircraft are calculated by means of a CFD simulation technology, and static stability and dynamic stability derivatives are obtained; obtaining a transfer function between the deflection angle and the attitude angle of a longitudinal or transverse lateral movement rudder of the aircraft according to the static derivative and the dynamic derivative, and designing a controller to realize the control of the pitching, rolling and yawing attitudes of the aircraft; and (3) obtaining a time-varying curve of the attitude angle and the rudder deflection angle of the aircraft by using a CFD and control coupling simulation technology, and verifying whether the control parameters meet the design requirements. If not, the corrected aircraft system transfer function and control parameters are further obtained through system identification, and verification is carried out by using a CFD and control coupling simulation technology to obtain better control parameters.
Description
Technical Field
The invention belongs to the technical field of computational fluid dynamics, flight dynamics and control, and particularly relates to an aircraft attitude control law design method based on a CFD (computational fluid dynamics) technology.
Background
Agility and over-stall maneuverability are important characteristics of high maneuvering aircrafts, in recent years, with rapid development of aviation science and technology, fifth-generation fighters represented by American F-22 and Chinese fighter 20 are greatly improved in the aspects of combat performance, maneuverability and the like, advanced air-to-air missiles such as American AIM-120, russian K-77M, china PL-15 and the like have maneuverability of rapid turning and omnibearing precise striking, and better maneuverability can help the fighters to occupy active positions during combat, and enhance the combat capability. Along with the increasingly prominent importance of air control rights in modern wars, the aircraft has higher requirements on agility and maneuverability, for example, the over-stall maneuverability of a new generation of fighter has become a necessary requirement, and future unmanned combat aircraft cancels the physiological overload limitation on people, so that the maneuverability and the agility of the aircraft are much higher than those of the existing manned fighter, the combat performance requirement of a new generation of air-to-air missile is improved, the aircraft has omnibearing three-dimensional attack capability, and in addition, the novel aerospace aircraft has higher maneuverability and can carry out large-scale hypersonic maneuvering flight in the adjacent space.
Meanwhile, when the aircraft flies under the conditions of high speed, large attack angle and super maneuvering, the complex unsteady flow and near-field vortex interference behaviors of the aircraft easily cause the nonlinear coupling phenomenon of the aircrafts in multiple disciplines such as pneumatics, movement, control and the like, the flight safety is seriously threatened, and the design difficulty of a control system is further increased. The traditional control system design adopts an iteration mode of trial flight and improvement, firstly decouples the unsteady aerodynamic problem and the rigid body dynamic problem based on the linear small disturbance hypothesis, acquires static and dynamic test data of the aircraft through a wind tunnel, establishes a pneumatic database, establishes a corresponding pneumatic model through parameter identification, further completes the design of the control system based on the pneumatic model, and finally improves and optimizes the control system in the trial flight process of the aircraft. The whole design iteration process is costly to complete and long in period, and due to the fact that unsteady aerodynamic characteristics and multidisciplinary cross coupling problems are not considered, the controllable safety range is manually limited in order to guarantee flight safety in the design process of the control system, and therefore the flight performance is reduced. In addition, the increased development performance requirements of aircraft and the application of new technologies also make flight tests face higher and higher technical and safety risks. Therefore, it is important to develop comprehensive design, verification and optimization using multidisciplinary simulation methods at the aircraft design stage.
Disclosure of Invention
The invention provides an aircraft attitude control law design method based on a CFD (computational fluid dynamics) technology, which is used for developing and designing an aircraft attitude control law by applying a relatively mature CFD numerical technology, is simple and convenient to realize and high in calculation efficiency, can obtain control parameters close to a real aircraft system, and can observe pneumatic and motion parameters and flow field state changes of an aircraft under the action of the control law, so that reference is provided for design and improvement of the control law of the aircraft, the flight test period is favorably shortened, the cost is reduced, and the method has important engineering value.
In order to achieve the purpose, the invention adopts the following technical scheme:
a method for designing an aircraft attitude control law based on a CFD technology comprises the following steps:
s10, carrying out grid division on the engineering object model according to the geometric shape of the engineering object model;
s20, solving aerodynamic force and moment coefficients of the model under different attitude angles (a pitch angle, a yaw angle and a roll angle) and rudder deflection angles (an elevator deflection angle, an aileron deflection angle and a rudder deflection angle) by using a CFD (computational fluid dynamics) technology, and calculating a static stability derivative and a dynamic stability derivative under an initial trim state;
s30, obtaining a transfer function between the rudder deflection angle and the attitude angle of the longitudinal or transverse lateral movement of the aircraft according to the static and dynamic stability derivative in the reference state obtained in the step S20;
s40, designing an attitude motion controller and setting a control target according to a transfer function of a motion rudder deflection angle and an attitude angle of the aircraft, and obtaining control parameters meeting design requirements through simulation of a control system to serve as initial control parameters;
s50, verifying control parameters by using a CFD and control system coupling simulation technology to obtain curves of changes of an aircraft attitude angle and a rudder deflection angle along with time and changes of unsteady aerodynamic force, moment and flow field characteristics;
s60, judging whether the time-varying curves of the attitude angle and the rudder deflection angle of the aircraft meet the design requirements, if so, obtaining control parameters reaching the design target and finishing, otherwise, executing the step S70;
and S70, carrying out system identification on a response curve of the aircraft attitude angle changing along with the rudder deflection angle, obtaining the corrected transfer function and corresponding control parameters, and continuously executing the step S50 until the design requirements are met.
In the above steps, step S20 specifically includes: under the assumption of a continuous medium, based on the mass, momentum and energy conservation law, a three-dimensional Reynolds average N-S equation under a rectangular coordinate system is established as follows:
wherein Q is a conservative variable, F (Q) and F v Respectively adopting a finite volume method for convection flux and viscous flux, constructing a universal lattice point format control body and a face-based data structure on a structural grid, adopting a mature second-order windward format Roe format for calculation of the convection flux, adopting a gradient decomposition method for the viscous flux to solve the gradient at the face center of a unit, and adopting a Gaussian-Green method to reconstruct the flow field gradient so as to achieve second-order spatial calculation accuracy; the time derivative term is dispersed by adopting an efficient implicit format, an accurate Jacobian matrix is constructed, se:Sup>A differential control equation is converted into se:Sup>A linear equation set of se:Sup>A large-scale sparse coefficient matrix which can be solved, efficient LU-SGS is adopted for iterative solution, an S-A turbulence model is adopted for sealing se:Sup>A fluid control equation, and therefore the aerodynamic force and the moment of the aircraft under different attitude angles and rudder deflection angles are obtained through calculation;
according to the required initial attitude angle, acquiring a trim rudder deflection angle by adopting a polynomial curve fitting mode, wherein a polynomial function is as follows:
y＝ax+b (2)
calculating by the CFD solving technology to obtain aerodynamic force, moment and static stability derivatives in a balancing state;
the derivative of the dynamic stability is obtained by making the aircraft do forced oscillating motions in the roll, pitch and yaw directions respectively, taking the pitch motion as an example, the forced pitch motion equation is as follows:
wherein α and θ represent an angle of attack and a pitch angle, respectively, α 0 At an initial angle of attack, theta m To the oscillation amplitude, ω is the oscillation frequency.
When the aircraft does small-amplitude forced pitching oscillation movement, the pitching moment coefficient C m Can be expressed in taylor expansion form:
wherein C is m0 Initial pitching moment coefficient in steady state, C mα For the purpose of the derivative of the static stability in pitch,
and C mq The washout time difference derivative and the pitch damping derivative, respectively.
Simplifying the process of forced pitch equations of motion yields:
where the reduction frequency k = ω c/(2V), c is the aircraft characteristic length, V is the flight speed, and substituting equation (5) into equation (4) and omitting the higher order terms may result:
the time-varying curve of the aerodynamic moment of the aircraft during the forced oscillation motion is obtained through the CFD technology iterative solution, the Fourier series y = A + Bsin (ω t) + Ccos (ω t) is adopted to fit the curve, A, B and C can be obtained, and therefore the pitching combined dynamic derivative is obtained:
the heaving motion of the aircraft may be represented as:
z(t)＝z m sin(ωt) (7)
wherein z is m The amplitude of pitching ups and downs movement.
According to the concept of aerodynamic derivative, when the aircraft makes small-amplitude ups and downs oscillation, the aerodynamic force of the aircraft can be expressed as:
and the change of the attack angle during the ups and downs motion can be expressed as:
substituting the above equation into equation (9) and omitting the higher order terms can result:
the differential derivative in washing flow can be obtained by adopting the same identification method as the combined dynamic derivative
And pitch resistanceDerivative of Ni C mq ；
Step S30 specifically includes: the transfer function between the deflection angle and the attitude angle of the rudder moving in the longitudinal direction and the transverse direction can be obtained according to the static derivative and the dynamic derivative in the trim state obtained in the step S20, and the deflection angle delta of the elevator e And the form of the transfer function for angle of attack α is as follows:
wherein
The natural frequency ω and the damping ratio ξ are respectively expressed as:
in the formula, rho represents air density, V represents flight speed, Q represents dynamic pressure, S and c represent wing reference area and average geometric chord length of the aircraft respectively, m represents mass of the aircraft, and I represents y Is moment of inertia, C D Is the coefficient of resistance in the reference state, C Lα And C mα Respectively the slope of the lifting line and the slope of the pitching moment curve,
and
the pitch moment damping derivative and the washout time difference derivative are represented separately,
the derivative of the elevator pitch moment.
Transverse lateral motion aileron deflection angle delta a And roll angle phi is in the form of a transfer function:
wherein the parameters are as follows:
wherein the content of the first and second substances,
in the formula
Respectively expressed as:
in the above formula, I x And I z Is moment of inertia, I xz Is the product of inertia, alpha is the angle of attack, theta e A pitch angle in a reference state, V is a flight speed, Q is a dynamic pressure, S and b are respectively a wing reference area and a wing span length of the aircraft, m is a mass of the aircraft,
and
side force derivatives and roll rate side force derivatives,
the derivative of the roll static stability is,
and
roll damping derivatives, cross-kinetic derivatives and roll steering derivatives,
and
respectively a course static stability derivative and a course damping derivative;
step S40 specifically includes: designing attitude motion controllers and limiting the yaw rate and range of the rudder deflection angle, e.g. the controller being a PID controller, by adjusting the PID parameter k p ,k i And k d The response of the system attitude angle along with the rudder deflection angle meets the control requirement through the simulation of a control system, so that initial control parameters are obtained;
step S50 specifically includes: the simulation technology of CFD coupling control is adopted to complete the simulation of aerodynamic force parameters and motion states of the aircraft under the controlled conditions, namely, the aerodynamic force and the moment of the flow field at each time step are obtained by applying the CFD solving technology and are used as input parameters for solving a six-degree-of-freedom motion equation, and the kinetic equation and the kinematic equation of the six-degree-of-freedom motion are as follows:
wherein u, v and w are each independentlyThe components of the flight speed on three coordinate axes; p, q and r are respectively a roll angular velocity, a pitch angular velocity and a yaw angular velocity; f x ，F y And F z The projection component of the resultant force vector of the external force acting on the aircraft on the dynamic system; l, M and N are respectively pneumatic rolling moment, pneumatic pitching moment and pneumatic yawing moment; phi, theta, psi are roll, pitch and yaw angles, respectively, and V represents the flight speed. c. C 1 ～c 9 Respectively expressed as:
solving the six-degree-of-freedom motion equation by adopting a fourth-order Runge-Kutta method, wherein the calculation formula is as follows:
wherein x is a state quantity, h represents a time step, c 1 ～c 4 Are constants of 1,2 and 1, respectively; k is a radical of formula 1 ～k 4 Is represented as follows:
obtaining the motion state parameters of the aircraft, including attitude angle, angular velocity, total force, moment and the like, solving the deviation value of the attitude angle of the current time step and the target attitude angle by the control module, obtaining the offset of the controlled sub-grid according to the control parameters obtained in S40, completing the six-degree-of-freedom motion of the uncontrolled sub-grid and the following motion and the deflection motion of the controlled sub-grid by using a dynamic grid deformation technology, repeating the steps to calculate the next time step, thereby realizing the control of the attitude of the aircraft, obtaining the attitude angle and rudder deflection angle curve of the aircraft changing along with the time, and the changing conditions of aerodynamic force, moment and flow field characteristics, and further verifying whether the control parameters meet the control target;
step S70 specifically includes: if the response of the attitude angle obtained by simulation in the step S50 along with the rudder deflection angle cannot meet the control target under the action of the control parameters obtained in the step S40, adjusting the control parameters; the rudder deflection angle is used as input, the attitude angle is used as output, system identification is carried out, a transfer function which is closer to an actual system after improvement is obtained, control parameters meeting requirements are obtained through control system simulation, and then the control parameters are verified through a CFD and control coupling simulation technology, so that better control parameters are obtained.
Has the beneficial effects that: the invention provides an aircraft attitude control law design method based on a CFD (computational fluid dynamics) technology, which is used for developing and designing an aircraft attitude control law by applying a relatively mature CFD technology, is simple and convenient to realize and high in calculation efficiency, can obtain control parameters close to a real aircraft system, can observe aerodynamic and motion parameters and flow field state changes of an aircraft under the action of the control law, can verify the feasibility of the control law on one hand, and can research the abnormal aerodynamic and kinematic laws of the aircraft on the other hand. Compared with the prior art, the method for designing the attitude control law of the aircraft based on the CFD technology can provide reference for the design of the control law in the design stage of the aircraft, shorten the design period of a control system, verify the designed control law, reduce the cost of a flight test and improve the flight safety.
Drawings
FIG. 1 is a flow chart of the method of the present invention;
FIG. 2 is a diagram of an engineering object geometry model and a mesh architecture provided in an embodiment of the present invention;
fig. 3 is a comparison between a CFD corresponding to a unit step pull-up mode and a control coupling simulation result and an MATLAB simulation result under the initial control parameter provided in the embodiment of the present invention;
fig. 4 is a comparison between a CFD corresponding to a unit step pull-up mode and a control coupling simulation result and an MATLAB simulation result under the modified control parameters provided in the embodiment of the present invention;
fig. 5 is a comparison between CFD corresponding to four pull-up modes (step pull-up, sinusoidal pull-up, constant-speed pull-up, and two-stage constant-speed pull-up in sequence from top to bottom) and a control coupling simulation result and an MATLAB simulation result provided in the embodiment of the present invention;
fig. 6 shows aerodynamic force, moment, and flow field changes obtained by CFD and control coupling simulation corresponding to five pull-up modes (unit step pull-up, sinusoidal pull-up, constant-speed pull-up, and two-stage constant-speed pull-up in sequence from top to bottom) provided in the embodiment of the present invention.
Detailed description of the invention
The invention will now be described more fully hereinafter with reference to the accompanying drawings and specific examples in which:
a method for designing an aircraft attitude control law based on a CFD technology is disclosed, a flow chart is shown in figure 1, the embodiment is designed aiming at the attitude control law of the unfolding of a certain type of winged missile, a PID controller is designed to enable the pitch angle of the missile to change as required under the control of a rudder deflection angle, and the method comprises the following steps:
s10, respectively generating grids for the missile body, the missile wing and the calculation domain of the missile according to the appearance of the engineering object model, wherein the grids are mutually overlapped to establish a nested relation as shown in figure 2;
s20, solving the pneumatic power and moment coefficients of the model under different pitch angles and elevator deflection angles by using a CFD (computational fluid dynamics) technology, and calculating the static stability derivative and the dynamic stability derivative under an initial balancing state (the attack angle is 0 degrees, and the rudder deflection angle is 0 degrees):
under the assumption of a continuous medium, based on the mass, momentum and energy conservation law, a three-dimensional Reynolds average N-S equation under a rectangular coordinate system is established as follows:
wherein Q is a conservative variable, F (Q) and F v Respectively adopting a finite volume method for convection flux and viscous flux, constructing a universal lattice point format control body and a face-based data structure on a structural grid, adopting a mature second-order windward format Roe format for calculation of the convection flux, and adopting a gradient decomposition method for solving the viscous fluxThe gradient at the center of the unit face is reconstructed by using a Gaussian-Green method to achieve second-order spatial calculation precision; the time derivative term is dispersed by adopting an efficient implicit format, an accurate Jacobian matrix is constructed, se:Sup>A differential control equation is converted into se:Sup>A linear equation set of se:Sup>A large-scale sparse coefficient matrix which can be solved, iterative solution is carried out by adopting an efficient LU-SGS, and an S-A turbulence model is adopted to seal se:Sup>A fluid control equation, so that the aerodynamic force and the moment of the aircraft under different attitude angles and rudder deflection angles are obtained through calculation;
according to the required initial attitude angle, acquiring a trim rudder deflection angle by adopting a polynomial curve fitting mode, wherein a polynomial function is as follows:
y＝ax+b (16)
calculating to obtain aerodynamic force, moment and static stability derivative in a balancing state by the CFD solving technology;
the derivative of the dynamic stability is obtained by making the aircraft do forced oscillating motions in the roll, pitch and yaw directions respectively, taking the pitch motion as an example, the forced pitch motion equation is as follows:
wherein α and θ represent an angle of attack and a pitch angle, respectively, α 0 At an initial angle of attack, theta m To the oscillation amplitude, ω is the oscillation frequency.
When the aircraft does small-amplitude forced pitching oscillation movement, the pitching moment coefficient C m Can be expressed in taylor expansion form:
wherein C is m0 Initial pitching moment coefficient in steady state, C mα For the purpose of the derivative of the static stability in pitch,
and C mq Differential derivatives of the washout time difference and pitch, respectivelyA damping derivative.
Simplifying the process of forced pitch equations of motion yields:
where the reduction frequency k = ω c/(2V), c is the aircraft characteristic length, V is the flight speed, equation (5) is substituted into equation (4) and higher terms are omitted:
the variation curve of the aerodynamic moment of the aircraft during forced oscillation motion along with time is obtained through iterative solution of the CFD technology, a Fourier series y = A + Bsin (ω t) + Ccos (ω t) is adopted to fit the curve, A, B and C can be obtained, and therefore the pitching combined dynamic derivative is obtained:
C m0 ＝A，
the heaving motion of the aircraft may be represented as:
z(t)＝z m sin(ωt) (21)
wherein z is m Is the amplitude of the pitching ups and downs movement.
According to the concept of aerodynamic derivative, when the aircraft makes small-amplitude ups and downs oscillation, the aerodynamic force of the aircraft can be expressed as follows:
and the change of the attack angle during the ups and downs motion can be expressed as:
substituting the above equation into equation (9) and omitting higher order terms can result:
the differential derivative during washing can be obtained by adopting the same identification method as the combined dynamic derivative
And pitch damping derivative C mq 。
S30, obtaining a longitudinal motion transfer function of the aircraft according to the static stability derivative and the dynamic stability derivative obtained in the step S20, and establishing an input-output relation between an elevator deflection angle and a pitch angle:
elevator rudder deflection angle delta e And the form of the transfer function for angle of attack α is as follows:
wherein
The natural frequency ω and the damping ratio ξ are respectively expressed as:
in the above formula, ρ represents the air density, V represents the flying speed, Q represents the dynamic pressure, and S and c represent the respective valuesReference area of the wing of the aircraft and the mean geometric chord length, m is the mass of the aircraft, I y Is moment of inertia, C D Is the coefficient of resistance in the reference state, C Lα And C mα Respectively the slope of the lifting line and the slope of the pitching moment curve,
and
the pitch moment damping derivative and the washout time difference derivative are represented separately,
is the derivative of the elevator pitch moment.
Transverse lateral motion aileron deflection angle delta a And roll angle phi is in the form of a transfer function:
the parameters in the formula are as follows:
wherein the content of the first and second substances,
in the formula
Respectively expressed as:
in the above formula, I x And I z Is moment of inertia, I xz Is the product of inertia, alpha is the angle of attack, theta e A pitch angle in a reference state, V is the flying speed, Q is the dynamic pressure, S and b are respectively the wing reference area and the wing span length of the aircraft, m is the mass of the aircraft,
and
side force derivatives and roll rate side force derivatives,
the derivative of the roll static stability is,
and
roll damping derivatives, cross-kinetic derivatives and roll steering derivatives,
and
respectively a course static stability derivative and a course damping derivative;
s40, aiming at a transfer function of the motion of the aircraft, adopting a PID controller and setting unit step lift as a control target, and obtaining a PID parameter meeting the requirement through simulation of a control system as an initial control parameter;
s50, verifying the PID parameters obtained in the step S40 by using a CFD and control system coupling simulation technology to obtain a change curve of the pitch angle and the rudder deflection angle of the missile along with time as shown in figure 3:
the flow field aerodynamic force and moment of each time step obtained by applying the CFD solving technology are used as input parameters for solving a six-degree-of-freedom motion equation, and the kinetic equation and the kinematic equation of the six-degree-of-freedom motion are as follows:
in the formula, u, v and w are components of the flight speed on three coordinate axes respectively; p, q and r are respectively a rolling angular velocity, a pitch angular velocity and a yaw angular velocity; f x ，F y And F z The projection component of the resultant force vector of the external force acting on the aircraft on the dynamic system; l, M and N are respectively pneumatic rolling moment, pneumatic pitching moment and pneumatic yawing moment; phi, theta, psi are roll, pitch and yaw angles, respectively, and V represents the flight speed. c. C 1 ～c 9 Respectively expressed as:
solving the six-degree-of-freedom motion equation by adopting a fourth-order Runge-Kutta method, wherein the calculation formula is as follows:
where x is the state quantity, h represents the time step, c 1 ～c 4 Are constants of 1,2 and 1, respectively; k is a radical of 1 ～k 4 Is represented as follows:
obtaining the motion state parameters of the aircraft, including attitude angle, angular velocity, total force, moment and the like, solving the deviation value of the attitude angle of the current time step and the target attitude angle by the control module, obtaining the offset of the controlled sub-grid according to the control parameters obtained in S40, completing the six-degree-of-freedom motion of the uncontrolled sub-grid and the following motion and the deflection motion of the controlled sub-grid by using a dynamic grid deformation technology, repeating the steps to calculate the next time step, realizing the control of the attitude of the aircraft, obtaining the attitude angle and rudder deflection angle curve of the aircraft changing along with time, and the changing conditions of aerodynamic force, moment and flow field characteristics, and further verifying whether the control parameters meet the control target;
s60, judging whether the variation curve of the pitch angle and the rudder deflection angle of the aircraft along with time meets the design requirement, wherein the pitch angle of the missile does not reach the control target under the action of the group of PID parameters driven rudder deflection angle, so that the transfer function of the system needs to be corrected, and S70 is continuously executed;
and S70, identifying the response curve of the aircraft attitude angle changing along with the rudder deflection angle obtained in the S50, obtaining a corrected transfer function and obtaining corresponding control parameters, and obtaining the curve of the attitude angle and the rudder deflection angle changing along with time through CFD and control system coupling simulation, wherein as shown in FIG. 4, the corrected control parameters can enable the attitude angle of the missile to change according to the specified requirements and are very close to the simulation result of MATLAB software. Based on the corrected transfer function, CFD and control coupled simulation of other four pull-up modes (step pull-up, sinusoidal pull-up, constant-speed pull-up, and two-stage constant-speed pull-up) is further realized, as shown in fig. 5, the attitude angle can be changed according to the target value by the control parameters corresponding to the four pull-up modes, and the attitude angle almost coincides with the MATLAB simulation value. Fig. 6 shows the changes of aerodynamic force, moment and flow field corresponding to the five pulling-up modes, and it can be seen that, in the pulling-up stage of the missile pitch angle, as the pitch angle increases, the pressure below the missile under each pulling-up mode gradually increases, the lift coefficient increases and shows the same change trend as the pitch angle, and the oscillation amplitude of the pitch moment coefficient gradually decreases; when the pitch angle reaches a target value, the lift resistance, the pitch moment coefficient and the flow field tend to be stable. Therefore, the method for designing the attitude control law of the aircraft based on the CFD technology can design, verify and improve the control parameters under the condition of considering the unsteady pneumatics and motion characteristics of the aircraft to obtain better control parameters, and can be used for evaluating the control law and improving the control performance.
The above description is only for the preferred embodiment of the present invention, but the scope of the present invention is not limited thereto, and any changes or substitutions that can be easily made by those skilled in the art within the technical scope of the present invention will be covered by the present invention without departing from the principle of the present invention.
Claims (9)
1. A method for designing an aircraft attitude control law based on a CFD technology is characterized by comprising the following steps:
s10, carrying out grid division on the engineering object model according to the geometric shape of the engineering object model;
s20, solving the pneumatic power and moment coefficients of the model under different attitude angles and rudder deflection angles by using a CFD (computational fluid dynamics) technology, and calculating a static stability derivative and a dynamic stability derivative under an initial balancing state;
s30, obtaining a transfer function between the rudder deflection angle and the attitude angle of the aircraft in longitudinal or transverse lateral movement according to the static and dynamic stability derivative obtained in the step S20;
s40, aiming at a transfer function of a motion rudder deflection angle and an attitude angle of the aircraft, designing an attitude motion controller and setting a control target, and acquiring control parameters meeting design requirements through simulation of a control system to serve as initial control parameters;
s50, verifying control parameters by using a CFD and control system coupling simulation technology to obtain curves of changes of an attitude angle and a rudder deflection angle of the aircraft along with time and change conditions of unsteady aerodynamic force, moment and flow field characteristics;
s60, judging whether the time-varying curves of the rudder deflection angle and the attitude angle of the aircraft meet the design requirements, if so, obtaining control parameters meeting the design requirements and ending, otherwise, executing the step S70;
s70, carrying out system identification on a response curve of the aircraft attitude angle changing along with the rudder deflection angle, obtaining the corrected transfer function and corresponding control parameters, and continuing to execute the step S50 until the design requirements are met.
2. The method for designing an attitude control law of an aircraft based on a CFD technology according to claim 1, wherein step S20 is specifically: solving a steady viscosity RANS equation by adopting CFD technical numerical values, and calculating aerodynamic force and moment of the aircraft under different attitude angles and rudder deflection angles; acquiring a trim rudder deflection angle and aerodynamic force, moment and static stability derivatives in a trim state by adopting a polynomial curve fitting mode according to the required initial attitude angle; the time-varying aerodynamic force condition of the aircraft in forced oscillation motion is obtained by iteratively solving the unsteady RANS equation, and Fourier series is adopted for fitting, so that the dynamic derivatives in the pitching, yawing and rolling directions are identified.
3. The method for designing the attitude control law of the aircraft based on the CFD technology according to claim 1 or 2, wherein the calculation of the aerodynamic force and the moment of the aircraft under different attitude angles and rudder deflection angles specifically comprises:
under the assumption of a continuous medium, based on the law of conservation of mass, momentum and energy, a three-dimensional Reynolds average N-S equation under a rectangular coordinate system is established as follows:
wherein Q is a conservative variable, F (Q) and F v Respectively adopting a finite volume method for convection flux and viscous flux, constructing a universal lattice point format control body and a face-based data structure on a structural grid, adopting a mature second-order windward format Roe format for calculation of the convection flux, adopting a gradient decomposition method for the viscous flux to solve the gradient at the face center of a unit, and adopting a Gaussian-Green method to reconstruct the flow field gradient so as to achieve second-order spatial calculation accuracy; time derivative terms are entered using efficient implicit formattingAnd (3) performing line dispersion, constructing an accurate Jacobian matrix, converting se:Sup>A differential control equation into se:Sup>A linear equation set of se:Sup>A large sparse coefficient matrix which can be solved, performing iterative solution by adopting an efficient LU-SGS, and sealing se:Sup>A fluid control equation by adopting an S-A turbulence model, so that the aerodynamic force and the moment of the aircraft under different attitude angles and rudder deflection angles are calculated.
4. The method for designing the attitude control law of the aircraft based on the CFD technology as claimed in claim 3, wherein the obtaining of the trim rudder deflection angle and aerodynamic force, moment and static stability derivatives in the trim state is specifically as follows: according to the required initial attitude angle, acquiring a trim rudder deflection angle by adopting a polynomial curve fitting mode, wherein a polynomial function is as follows:
y＝ax+b (2)
and calculating by a CFD (computational fluid dynamics) solving technology to obtain derivatives of aerodynamic force, moment and static stability in a balancing state.
5. The method for designing an attitude control law of an aircraft based on a CFD technology according to claim 1, wherein the step S30 is specifically as follows: obtaining a transfer function between the rudder deflection angle and the attitude angle of longitudinal and transverse lateral movement according to the static and dynamic stability derivative in the trim state obtained in the step S20, wherein the elevator deflection angle delta e And the form of the transfer function for angle of attack α is as follows:
wherein
The natural frequency ω and the damping ratio ξ are respectively expressed as:
in the formula, rho represents air density, V represents flight speed, Q represents dynamic pressure, S and c represent wing reference area and average geometric chord length of the aircraft respectively, m represents mass of the aircraft, and I represents y Is moment of inertia, C D Is the coefficient of resistance in the reference state, C Lα And C mα Respectively the slope of the lifting line and the slope of the pitching moment curve,
and
the pitch moment damping derivative and the washout time difference derivative are represented separately,
is the elevator pitch moment derivative;
transverse lateral motion aileron deflection angle delta a And roll angle phi is in the form of a transfer function:
wherein the parameters are as follows:
wherein the content of the first and second substances,
in the formula
Respectively expressed as:
in the above formula, I x And I z Is moment of inertia, I xz Is the product of inertia, alpha is the angle of attack, theta e A pitch angle in a reference state, V is the flying speed, Q is the dynamic pressure, S and b are respectively the wing reference area and the wing span length of the aircraft, m is the mass of the aircraft,
and
side force derivatives and roll rate side force derivatives,
the derivative of the roll static stability is,
and
roll damping derivatives, cross-kinetic derivatives and roll steering derivatives,
and
respectively, a course static stability derivative and a course damping derivative.
6. The method for designing an attitude control law of an aircraft based on a CFD technology according to claim 1, wherein the step S50 specifically comprises: the method comprises the steps of taking pneumatic force and moment of a flow field of each time step obtained through CFD calculation as input, obtaining motion state parameters of an aircraft by solving a six-degree-of-freedom motion equation, further solving a deviation value of an attitude angle of the current time step and a target attitude angle through a control module, obtaining offset of a controlled sub-grid according to the control parameters obtained in step S40, completing six-degree-of-freedom motion of an uncontrolled sub-grid, following motion and deflection motion of the controlled sub-grid by using a dynamic grid deformation technology, repeating the steps after all grids are updated, and then carrying out calculation of the next time step, so that the attitude of the aircraft is controlled, obtaining attitude angle and rudder deflection angle curves of the aircraft changing along with time, and changing conditions of pneumatic force, moment and flow field characteristics, and further verifying whether the control parameters meet control targets.
7. The method for designing attitude control law of aircraft based on CFD technology according to claim 6, wherein the six-degree-of-freedom kinematic equation and the kinematic equation are as follows:
in the formula, u, v and w are components of the flight speed on three coordinate axes respectively; p, q and r are respectively a rolling angular velocity, a pitch angular velocity and a yaw angular velocity; f x ，F y And F z The projection component of the resultant force vector of the external force acting on the aircraft on the dynamic system; l, M and N are respectively pneumatic rolling moment, pneumatic pitching moment and pneumatic yawing moment; phi, theta and psi are respectively a rolling angle, a pitching angle and a yaw angle, and V represents the flight speed; c. C 1 ～c 9 Respectively expressed as:
8. the method for designing the attitude control law of the CFD technology-based aircraft according to claim 7, wherein the six-degree-of-freedom motion equation is solved by a fourth-order Runge-Kutta method, and the calculation formula is as follows:
where x is the state quantity, h represents the time step, c 1 ～c 4 Are constants of 1,2 and 1, respectively; k is a radical of 1 ～k 4 Is represented as follows:
thereby obtaining the motion state parameters of the aircraft.
9. The method for designing an attitude control law of an aircraft based on a CFD technology according to claim 1, wherein step S70 specifically includes: the rudder deflection angle is used as input, the attitude angle is used as output, system identification is carried out, a transfer function which is closer to an actual system after improvement is obtained, control parameters meeting requirements are obtained through control system simulation, and then the control parameters are verified through a CFD and control coupling simulation technology, so that better control parameters are obtained.
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| CN117077296B (en) * | 2023-10-17 | 2024-01-09 | 中国空气动力研究与发展中心计算空气动力研究所 | Control coupling simulation method for aerodynamic structure of aircraft |
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