RU2338235C1 - Method for generating of flying aircraft angular motion adaptive control signal - Google Patents

Abstract

FIELD: instrument making.

SUBSTANCE: method comprises following operations, i.e. measurement of current angular position and angular speed signals, generation of angular position reference signal, generation of signal of mismatch between the angular position preset and current signals. The method includes also generating the control signal equal to the sum of the mismatch and angular speed signal components to initiate actuator, measuring kinetic head signal and generating adaptive signal be proportional scaling the kinetic head signal.

EFFECT: invariance of stabilisation quality and increased control accuracy.

Description

The invention relates to airborne automatic control systems for substantially unsteady unmanned aerial vehicles.

A known method of generating a control signal for an aircraft, in which the signals of the driving action, comparison, summation are generated and the signals of the angle and angular velocity are measured [1].

The disadvantage of this method is the limited functionality in the conditions of non-stationary parameters of the aircraft caused by changes in speed and altitude.

Closest to the proposed invention is a method of generating an adaptive signal for controlling the angular motion of an unsteady aircraft, which consists in measuring the current signals of the angular position and angular velocity, setting the reference signal of the angular position, generating a mismatch signal between the given signal and the current signal of the angular position, a control signal equal to the sum of the components of the mismatch signals and the angular velocity to act on the actuator [2].

The disadvantage of this method is the limited functionality for stability and accuracy, due to the lack of means of invariance to solve the problem of unsteadiness of the aircraft in conditions of a significant change in flight conditions in speed and altitude.

Solved in the proposed method, the technical problem is to ensure the invariance of the quality indicators of stabilization processes and improving the accuracy of control characteristics. The proposed formation of the method provides for the adaptation of the parameters of the regulatory part of the law of stabilization and increased stability and quality in a wide range of parameters of the aircraft.

где m^δ - производная коэффициента эффективности по отклонению рулей; s и l - характерные геометрические параметры летательного аппарата - площадь и длина, сохраняя при этом значения λ_min=Aq_min при q<q_min и λ_max=Aq_max при q>q_max, где q_min и q_max - соответственно расчетные сигналы минимального и максимального значений скоростного напора, формируют сигнал адаптивно перестраиваемого передаточного коэффициента К₁ по рассогласованию, равный

A₁=const, формируют сигнал адаптивно перестраиваемого передаточного коэффициента К₂ по угловой скорости, равный К₂=А₂К₁, A₂=const, при этом константы A₁, А₂ определены по условиям устойчивости и качества переходных процессов, формируют компоненту сигнала рассогласования, равную произведению сигналов рассогласования и адаптивно перестраиваемого передаточного коэффициента К₁ по рассогласованию и формируют компоненту сигнала угловой скорости, равную произведению сигналов угловой скорости и адаптивно перестраиваемого передаточного коэффициента К₂ по угловой скорости.The specified technical result is achieved by the fact that in the known method of forming an adaptive signal for controlling the angular motion of an unsteady aircraft, which consists in measuring the current signals of the angular position and angular velocity, setting the reference signal of the angular position, generating a mismatch signal between the given signal and the current signal of the angular position, form a control signal equal to the sum of the components of the mismatch signals and angular velocity, to affect the actuator oystvo further measured dynamic pressure q signal, generating adjustment signal λ by proportional scaling signal dynamic pressure λ = Aq, A = const,

where m ^δ is the derivative of the coefficient of efficiency for the deviation of the rudders; s and l are the characteristic geometric parameters of the aircraft — the area and length, while maintaining the values λ _min = Aq _min for q <q _min and λ _max = Aq _max for q> q _max , where q _min and q _max are the calculated signals, respectively the minimum and maximum values of the pressure head, form a signal adaptively tunable gear ratio K ₁ mismatch equal to

A ₁ = const, form the signal of the adaptively tunable gear coefficient K ₂ at an angular velocity equal to K ₂ = A ₂ K ₁ , A ₂ = const, while the constants A ₁ , A _{2 are} determined by the conditions of stability and quality of transients, form the component the mismatch signal equal to the product of the mismatch signals and the adaptively tunable transfer coefficient K ₁ according to the mismatch and form the component of the angular velocity signal equal to the product of the angular velocity signals and the adaptively tunable transfer coefficient K ₂ by angular velocity.

Indeed, this ensures the development of control signals with maximum quality in a wide range of altitude and flight speed changes through the introduction of adaptation tools and the implementation of the adaptive control law.

Consider the generalized formation of the proposed method for the formation of an adaptive signal for controlling the angular motion of an unsteady aircraft in full on the example of one channel. The basic equations of the angular motion of an aircraft, for example, according to [3] are described in the form:

where a, b are the dynamic coefficients of the aircraft for damping and efficiency, respectively;

φ is the angle;

ω _φ is the angular velocity;

δ is the angle of deviation of the steering surfaces by the actuator.

The control law forming the control signal for the actuator is formed in the form of:

where Δφ is the error signal:

here φ _ass is the signal of the driving action;

To ₁ , To ₂ - gear ratios.

Having adopted the inertia-free refinement by the actuator of the control signal, i.e. putting δ≡σ from equations (1) ÷ (3), we obtain a description of the regulation processes in a closed loop (control law - aircraft) in the form:

or

The characteristic equation of a closed-loop control system in accordance with (5) has the form:

Equation (6) allows us to estimate the parameters of the control law (2) K ₁ and K ₂ with the dynamic coefficients of the aircraft a and c.

To ensure the required stability and quality characteristics, it is necessary to ensure that the stability conditions (invariance) of the coefficients of the characteristic equation are satisfied during the flight, i.e.

In (7), the coefficient a, which characterizes the aircraft’s own damping, is rather small, slightly changes, and is not characteristic, dominant in comparison with the parameter in K ₂ , which determines the damping of the closed control system as a whole. Therefore, the conditions for maintaining the required indicators of stability and quality, based on (7), can be defined as:

where in - the efficiency factor of the controls, equal, for example, according to [3]:

where m ^δ is the derivative of the coefficient of efficiency with respect to the deviation of the rudders δ;

s, l - characteristic geometric parameters of the aircraft: area and length;

J is the moment of inertia of the aircraft;

q - velocity head:

here ρ is the air density at the current flight altitude, ρ = ρ (Н);

V is the flight speed.

The parameter m ^δ in the current state is mainly a function of the Mach number:

where a is the speed of sound at the current altitude; for an aerodynamic aircraft varies within small limits and can be assumed constant - averaged or maximum for a given height range. This coefficient can be considered stable in the vicinity of the balancing values of the current angles of the aircraft.

The moment of inertia J for unmanned aerial vehicles also varies slightly. This circumstance is all the more correct because structurally changing the mass of the aircraft due to fuel burnup also alters the centering characteristics, thereby ensuring the maximum stability of the moments of inertia.

Thus, the main non-stationarity is the velocity head q. Based on equations (8) and on the basis of equations (9) ÷ (11), taking into account the above, to ensure the adaptation processes, it is necessary to form an adaptation function - we denote it by λ:

λ = Aq,

where A = const,

Moreover, for modes with q belonging to non-calculated values q <q _min and q> q _max , the quantity λ is taken at the corresponding values, i.e. generally

From equations (8) we obtain adaptation algorithms:

Most accurately, rather complex laws of identification and adaptation based on (12) can be realized through the use of computer systems. The signals of the identification and adaptation functions are easily implemented algorithmically; all links and blocks can also be implemented on elements of automation and computer technology, for example, according to [4, 5].

The proposed method for the formation of an adaptive control signal for the angular motion of an unsteady aircraft makes it possible to ensure the invariance of quality indicators and improve control accuracy in a wide range of flight conditions.

Information sources

1. Aerodynamics, stability and controllability of supersonic aircraft. Ed. G.S. Byushgens. M .: Science. Fizmatlit, 1998, p. 433.

2. RF patent №2251136, 12.24.02, cl. G05D 1/08.

3. A.A. Lebedev, L.S. Chernobrovkin. Flight dynamics of unmanned aerial vehicles. M.: Mechanical Engineering, 1973, p. 486.

4. V. B. Smolov. Functional information converters. L .: Energy Publishing House, Leningrad Branch, 1981, p. 22, 41.

5. A.U. Yalyshev, O. I. Razorenov. Multifunctional analog control devices for automation. M.: Mechanical Engineering, 1981, p. 107, 126.

Claims (1)

A method for generating an adaptive angular motion control signal of an unsteady aircraft, which consists in measuring the current angular position and angular velocity signals, setting the angular position reference signal, generating a mismatch signal between the given signal and the current angular position signal, generating a control signal equal to the sum of the components mismatch signals and angular velocity to act on the actuator, characterized in that they measure the pressure signal q form an adaptation signal λ by proportional scaling of the pressure head signal λ = Aq, A = const

where m ^δ is the derivative of the efficiency coefficient for rudder deflection, s and l are the characteristic geometric parameters of the aircraft - area and length, while maintaining the values λ _min = Aq _min for q <q _min and λ _max = Aq _max for q> q _max , where q _min and q _max - respectively, the calculated signals of the minimum and maximum values of the pressure head, form a signal of adaptively tunable gear ratio K ₁ according to the mismatch, equal to

A ₁ = const, form the signal of the adaptively tunable gear coefficient K ₂ at an angular velocity equal to K ₂ = A ₂ K ₁ , A ₂ = const, while the constants A ₁ , A _{2 are} determined by the conditions of stability and quality of transients, form the component error signal equal to the product of the error signal and adaptively tunable transfer coefficient K ₁ to the mismatch, and generating a signal component of the angular velocity equal to the angular velocity signals and adaptively tunable transmission coefficient itsienta K ₂ on the angular velocity.