RU2644842C2 - System of automated modal control in aircraft longitudinal axis - Google Patents

Abstract

FIELD: aviation.

SUBSTANCE: system of automated modal control in an aircraft (AC) longitudinal channel comprises a pilot stick/pitch selector, a pitch angle autopilot computer, a servo-actuator, a pitch-angle pickoff, a marginal condition limiter, a pitch angular rate sensor, a balancing unit, a modal control algorithm computer (MCAC), an air data system, connected in a certain way. The servo-actuator contains a hydraulic drive and a minimum signal selector. The marginal condition limiter contains a maximum attack angle selector and an angle-of-attack restriction machine computer. MCAC contains a signal conditioner of averaged reduced lift force coefficient, a program unit of system transfer function by pitching angular velocity signal, a discrepancy unit of pitching angular velocity, a control signal conditioner.

EFFECT: improving flight safety by improving the characteristics of aircraft control due to creation of an aircraft control synthesis algorithm in a longitudinal plane.

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Description

The invention relates to the field of automatic control systems (ACS) of the pitch angle of an aircraft (LA).

Known self-propelled guns that ensure the development of a given pitch angle of an aircraft with the help of an autopilot acting on the angle of deviation of the elevator of the aircraft [1].

To implement the well-known aircraft control methods, modal control methods based on the choice of the poles of the transfer function of a closed system or the roots of its characteristic equation are used. If all the components of the state vector of the object can be measured and the model of the dynamics of the object is linear, then providing a given location of the roots of a closed system does not cause difficulties. The required location of the roots of the characteristic equation for each object is determined separately.

If the transfer function of a closed system does not contain zeros, then the solution of the problem of the required location of the roots of the characteristic equation can be solved using the classical method of standard coefficients. Further, the assessment of the correspondence of the obtained coefficients of the denominator of the transfer function is estimated by the nature of transients in a closed system.

Existing methods of automatic control of aircraft require adjustments in flight of gear ratios of control laws depending on changes in the angle of attack when the aerodynamic and mass inertia characteristics of the object change.

Closest to the technical result achieved, chosen as a prototype, the self-propelled gun is adopted by the pitch angle of the aircraft, which implements the astatic control law with high-speed feedback, containing a pitch angle adjuster and a pitch angle autopilot calculator, a servo drive whose output signal determines the angle of deviation of the elevator apparatus, the pitch angle sensor of the aircraft, having an output connected to the second input of the pitch angle calculator autopilot [2].

The problem that this invention solves is achieved by the fact that the automatic pitch angle control system and the aircraft angle of attack limits include a pitch angle adjuster and a pitch angle autopilot calculator, a servo drive whose output signal determines the angle of deviation of the aircraft elevator, angle sensor pitch of an aircraft having an output connected to a second input of a pitch angle calculator autopilot, serially connected maxi the angle of attack, the calculator of the automatic machine for limiting the angle of attack and the selector of the minimum signal, the output of which is connected to the input of the servo drive, the angular velocity sensor of the roll of the aircraft having an output connected to the second input of the calculator of the automatic machine for limiting the angle of attack, the output of the calculator of the autopilot pitch angle is connected to the second input minimum signal selector.

The disadvantage of this invention is the need to adjust in flight the gear ratios of the control laws depending on the change in the angle of attack, when the aerodynamic and mass inertia characteristics of the object change, and the result of the operation of such a regulator does not satisfy the requirements of the pilot to ensure a minimum of mental and physical stress when working on ground and air targets.

The technical result of the invention proposed by the authors is to solve the problem of improving the control characteristics of an aircraft by constructing an algorithm for synthesizing an aircraft control in the longitudinal plane.

(

- частная производная нормальной перегрузки по углу атаки). Для получения процессов в контуре управления, близких к апериодическим, на основании критерия Шомбера-Гертсена можно сформировать требования к желаемым значениям параметров объекта

и

(

- производная приведенной к произведению массы ЛА на скорость установившегося полета аэродинамической подъемной силы крыла по углу атаки,

- производная приведенного к моменту инерции относительно связанной оси Oz момента тангажа по углу атаки). При

удается получить корни характеристического многочлена системы с большими отрицательными вещественными частями, тем самым выполнить требования к апериодичности переходных процессов в контуре «летчик-система автоматического управления». При этом соотношение

может меняться на ±30%, важно, чтобы это соотношение не выходило за область, удовлетворяющую критерию. На основании этого соотношения формируются линейные дифференциальные уравнения желаемой системы.The technical result is achieved by constructing the desired model based on the Schomber-Gertsen criterion [3], which implies the separation into two types of aircraft control into pitch and reload, the value is taken abroad

(

- partial derivative of normal overload by angle of attack). To obtain processes in the control loop that are close to aperiodic, based on the Schomber-Gertsen criterion, it is possible to formulate requirements for the desired values of the object parameters

and

(

- the derivative of the product of the mass of the aircraft by the speed of a steady flight of the aerodynamic lifting force of the wing by the angle of attack,

is the derivative of the pitch moment with respect to the connected axis Oz of the moment of inertia with respect to the angle of attack). At

it is possible to obtain the roots of the characteristic polynomial of the system with large negative real parts, thereby fulfilling the requirements for the aperiodicity of transients in the “pilot-automatic control system” circuit. Moreover, the ratio

can vary by ± 30%, it is important that this ratio does not go beyond the region that meets the criterion. Based on this relationship, linear differential equations of the desired system are formed.

The figure 1 presents a functional diagram of a system of automated modal control (ACS) in a longitudinal channel:

limit mode limiter (OPR) 1, consisting of a setter of the maximum angle of attack and a computer of the automaton of limiting the angle of attack;

pilot handle / pitch angle adjuster (RL / ZT) 2;

angular velocity sensor (TLS), incorporating a pitch velocity sensor 3;

pitch angle autopilot calculator (WAUT) 4;

the modal control algorithm calculator (VAMU) 5;

a servo drive (SP), which includes a hydraulic drive and a minimum signal selector 6;

balancing unit (BB) 7;

system of air signals SHS 8;

pitch angle sensor (FLS) 9.

The figure 2 presents the structural diagram of the self-propelled guns, where:

1 - ODA - limiter limit modes;

2 - RL / ZD - pilot's handle / pitch angle adjuster;

3 - TLS - angular velocity sensor (ω _z );

4 - VAUT - calculator of autopilot pitch angle;

5 - VAMU - a modal control algorithm calculator, which includes blocks:

;BFS - control signal conditioning unit

;

FFT - transfer function block;

BN - residual block;

BFSU - control signal conditioning unit (u);

6 - SP - servo drive;

7 - BB - balancing unit;

8 - SHS - system of air signals;

9 - FLS - pitch angle sensor;

LA - aircraft;

x _m is the state vector of the model, A _m is (n × n) and B _m is (n × 1) are the matrices of known parameters of the desired model, u (t) is the control, n _y is the normal overload, q is the pressure head, F ₁ — signal from the pilot’s handle, ω _z — output signal from the aircraft pitch angle sensor, ω _m — output signal from the desired reference model, k _i —scale coefficients of the control law.

, блок оптимальной по критерию Шомбера-Гертсена передаточной функции системы по сигналу угловой скорости тангажа, блок невязки по угловой скорости тангажа (Δω_z), к первому входу которого подключен сигнал с выхода датчика угловой скорости тангажа (ω_z), ко второму входу - сигнал угловой скорости тангажа с выхода блока передаточной функции (ω_zm), при этом сумма масштабированных астатического сигнала (Δω_z/p) и сигнала невязки (Δω_z) с выхода блока невязки поступает на вход сервопривода и на первый вход блока передаточной функции, второй вход которого соединен с выходом датчика ручки летчика (задатчиком угла тангажа), и блок, формирующий сигнал управления (u) по правилу:An automated modal control system in the longitudinal channel of the aircraft contains a serially connected pilot handle / pitch adjuster and a pitch angle autopilot calculator, a hydraulic drive and a minimum signal selector as part of the servo drive, the output signal of which determines the angle of deviation of the aircraft elevator, and the aircraft pitch angle sensor, having an output connected to the second input of the pitch angle autopilot calculator, limit mode limiter (OPR), consisting from a series-connected adjuster of the maximum angle of attack, the calculator of the automatic machine for limiting the angle of attack, the output of which is connected to the input of the servo drive, the sensor of angular pitch velocity (TLS) of the aircraft having an output connected to the input of the calculator of the automatic machine for limiting the angle of attack, the output of the calculator of the autopilot of the pitch angle is connected to the second input of the minimum signal selector, as well as the balancing unit. In addition, to ensure a given control quality in the entire field of application of the aircraft when changing the aerodynamic and mass inertia characteristics of the pitch angle, a modal control algorithm calculator is introduced, containing a program block for generating a signal of the averaged reduced lift coefficient

, the block of the transfer function of the system that is optimal according to the Schomber-Gertsen criterion for the pitch angular velocity signal, the residual block in the pitch angular velocity (Δω _z ), to the first input of which the signal from the pitch pitch sensor (ω _z ) is connected, and the signal the angular velocity of the pitch from the output of the transfer function block (ω _zm ), while the sum of the scaled astatic signal (Δω _z / p) and the residual signal (Δω _z ) from the output of the residual block goes to the input of the servo drive and to the first input of the transfer function block, the second input cat It is connected to the output of the pilot’s pen sensor (pitch angle adjuster), and the unit generating the control signal (u) according to the rule:

,

,

в соответствии с критерием Шомбера-Гертсена в зависимости от величины скоростного напора от СВС.based on the data on the pitch angular velocities obtained from the aircraft's angular velocity sensor and from the desired reference model, the parameters of which are adjusted by the software generating unit

in accordance with the Schomber-Gertsen criterion, depending on the magnitude of the pressure head from the SHS.

The proposed control formation system linearizes the final nonlinear model of the object. A significant advantage of the proposed approach is manifested when controlling an aircraft at large angles of attack, when the characteristics of the aircraft are substantially non-linear. As a result, there is no need to enter additional control signals into the ACS.

Literature

1. Bodner V.A. Control systems for aircraft. - M.: Mechanical Engineering, 1973. - 506 p. Page 101, fig. 3.9; 2. Krasovsky A.A. Automatic flight control systems and their analytical design. - M .: Nauka, 1973. - 560 p. Page 179, fig. 5.2; 3. Mikhalev I.A., Okoyemov B.N., Chikulaev M.S. Automatic aircraft control systems. - M.: Mechanical Engineering, 1987. - p. 240. Page 192, fig. 14.2, p. 194, fig. 14.4, p. 198, fig. 14.7, p. 201, fig. 14.9.

2. Patent RU2434785 System for automatic control of pitch angle and limitation of the angle of attack of an aircraft, authors Petunin VI, Abdullina E.Yu., Efanov V.N. IPC V64C 13/18, publ. 11/27/2011.

3. Shomber N., Gertsen W., - “Longitudinal Handing Qualities Griteria: an Evaluation AIAA Paptr, N65-780, 1965.

Claims (7)

An automated modal control system in the longitudinal channel of the aircraft, containing a pilot's pitch / pitch adjuster and a pitch angle autopilot calculator, a hydraulic actuator and a minimum signal selector as part of the servo drive, the output signal of which determines the angle of deviation of the aircraft elevator, and the aircraft's pitch angle sensor having an output connected to the second input of the pitch angle autopilot calculator, limit mode limiter (OPR), state the one from the series-connected adjuster of the maximum angle of attack, the calculator of the automatic machine for limiting the angle of attack, the output of which is connected to the input of the servo drive, the sensor of angular pitch speed (TLS) of the aircraft, having an output connected to the input of the computer of the automatic machine for limiting the angle of attack, the output of the calculator of the autopilot of the pitch angle to the second input of the minimum signal selector, as well as a balancing unit, characterized in that an additional modal control algorithm calculator (VAMU) is introduced, containing th block of the signal averaged reduced coefficient of lift

, the program block of the transfer function of the system that is optimal according to the Schombert-Gertsen criterion according to the pitch angular velocity signal, the residual block of the pitch angular velocity (Δω _z ), the first input of which receives a signal from the pitch angular velocity sensor (ω _z ), the second input - signal angular pitch velocity output from the block transfer function (ω _zm), and the sum of scaled astatic signal (Δω _z / p) and a residual signal (Δω _z) output from the residual block is input to the servo and to the first input of the block transfer function, the second whose input is connected to the output of the sensor knob pilot / pitch set point, in addition, it includes WAYS unit generating a control signal (u) according to the rule:

where: F _l - signal from the handle of the pilot, k _i - scale factors of the control law, ω _z - the output signal of the sensor angular pitch pitch of the aircraft, ω _zm is the output signal from the desired reference model, n _y is normal overload.

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