RU2279119C1 - Adaptive system for controlling flight height of an aircraft - Google Patents

Abstract

FIELD: engineering of onboard systems for automatic control of aircrafts, which during flight realize turns at significant angles of attack.

SUBSTANCE: control system contains aircraft as control object, steering drive, device for measuring angular position by pitch, adding amplifier, speed force indicator, indicator of height and height alternation speed, flight speed indicator, height signal set-point device, discrepancy block, first signal limiting block, functional transformer, second signal limiting block, functional device for limiting signal, first set-point device of bearing signal, first division block and first multiplication block, while functional transformer contains two of each of the following: set-point device of bearing signal, division block, multiplication block and adder.

EFFECT: expanded functional capabilities, limitation of angle of attack of an aircraft.

2 cl, 3 dwg

Description

The invention relates to airborne systems for the automatic control of aircraft, which realize turns with significant angles of attack during flight.

Aircraft control systems are known that include a control signal adjuster, a pitch angle sensor, a pitch angle sensor and a summing amplifier in the pitch channel, which generate control actions on the steering gears of the aircraft according to the driving actions and signals of the state sensors [1].

The disadvantage of this implementation is the limited control capabilities that require means to limit the angle of attack of the aircraft.

Known solutions include the introduction of additional filters in the control channel to limit the angle of attack [2]. However, these filters solve particular problems, as described in [2], for example, in conditions of wind gusts; at the same time, filters in a direct circuit weaken the passage of control signals.

Known decisions on the formation of a flight altitude control system include the control system according to [3]. This system also includes a pitch control loop, which includes units similar to [1] described above, and a flight altitude control loop containing a sensor for flight altitude and rate of change (for example, a radio altimeter), an altitude signal adjuster, and a height mismatch unit and a summing amplifier for generating a flight altitude control law.

Closest to the proposed invention is a control system for the pitch channel of the aircraft, comprising a pitch angle meter and a steering gear [4].

The disadvantages of the known system are limited functionality in conditions of non-stationary parameters and the lack of means to limit the angle of attack of the aircraft.

The technical problem solved in the proposed control system is the expansion of functionality and limitation of the angle of attack.

The specified technical result is achieved by the fact that in the known control system of the flight altitude of the aircraft containing the aircraft as a control object, steering gear, pitch angle meter, pitch angle meter and summing amplifier, an additional pressure head sensor, height sensor and altitude change rate, flight speed sensor, altitude adjuster, mismatch unit, first signal limiting unit, functional converter, second unit signal limiting, signal limiting functional device, first reference signal adjuster, first division block and first multiplication block, the output of the height signal adjuster through the mismatch unit is connected to the input of the summing amplifier, the first and second outputs of the height sensor and the rate of change of height are connected respectively to the second input of the block mismatch and through the first block of multiplication - with the second input of the summing amplifier, the output of the first block signal limitation is connected through series-connected the functional converter and the second signal limiting block to the input of the steering gear, and the input of the first signal limiting block is connected to the output of the summing amplifier, the pressure head sensor is connected to the second input of the functional converter directly and through the signal limiting functional device to the second input of the first signal limiting block, meter the pitch angular position and the pitch angular velocity meter are connected respectively to the third and fourth inputs of the functional about the converter, the first reference signal adjuster and the flight speed sensor are connected to the first and second inputs of the first division block, respectively, and the output of the first division block is connected to the second input of the first multiplication block. In addition, the functional converter comprises a second reference signal generator and a second division unit connected in series, a first adder, a second multiplication unit and a second adder connected in series, the output of which is a functional converter output, a third reference signal generator, a third division unit and a third multiplication unit the output of which is connected to the second input of the second adder, the output of the second division unit is connected to the second input of the second multiplication unit, when this, the first input of the first adder, the second input of the second division unit, the second input of the first adder and the second input of the third multiplication unit are the first, second, third and fourth inputs of the functional converter, respectively.

Figure 1 shows the structural diagram of the control system, figure 2 shows a block diagram of a functional Converter, figure 3 presents a static characteristic of the functional device of the signal limiter.

The adaptive control system of the flight altitude of the aircraft (Fig. 1) contains serially connected signal height 1 (ZSV), mismatch unit 2 (BR), summing amplifier 3 (SU), first signal restriction unit 4 (1 BOS), functional converter 5 (FP), the second signal limiting block 6 (2 BOS), steering gear 7 (RP) and the aircraft as a control object 8 (LA), a pressure head sensor 9 (SDS) is connected to the second input of the functional converter 5 directly and through the functional device limited Signal 10 (FCF) - with the second input of the first block of signal restriction 4, the pitch angle meter 11 (IUPT) and pitch angle meter 12 (IUST) are connected respectively to the third and fourth inputs of the functional transducer 5, the first and second outputs the height sensor and the rate of change of height 13 (DViSIV) are connected respectively to the second input of the mismatch unit 2 and through the first multiplication unit 14 (1 control unit) with the second input of the summing amplifier 3, the first reference signal adjuster 15 (1 AIA) and the speed sensor flight 16 (DSP) are connected to the first and second inputs of the first division unit 17 (1 OBD), respectively, and the output of the first division unit 17 is connected to the second input of the first multiplication unit 14.

Functional converter 5 (figure 2) contains a second connected reference signal setter 18 (2 AIA) and a second division unit 19 (2 DBs) connected in series to the first adder 20, the second multiplication unit 21 and the second adder 22, the output of which is the output of the functional the converter, connected in series with the third reference signal master 23, the third division unit 24 and the third multiplication unit 25, the output of which is connected to the second input of the second adder 22, the output of the second division unit 19 is connected to the second input of the second multiplication unit 21, wherein the first input of the first adder 20, the second input of the second division unit 19, the second input of the first adder 20 and the second input of the third multiplication unit 25 are the first, second, third and fourth inputs of the functional converter 5, respectively.

The control system operates as follows.

, формируемый системой управления на основе сигналов задатчиков 1 и 15 и датчиков первичной информации 11, 12, 13 летательного аппарата 8, подается на рулевой привод 7, отклонения руля которого воздействуют на летательный аппарат, изменяя его положение в соответствующем направлении. Выходными параметрами летательного аппарата являются: угол тангажа ϑ, угловая скорость

, скорость полета V, высота полета Н, скорость изменения высоты полета

и скоростной напор q.Control output

, formed by the control system based on the signals of the setters 1 and 15 and primary information sensors 11, 12, 13 of the aircraft 8, is fed to the steering gear 7, the steering deviations of which affect the aircraft, changing its position in the corresponding direction. The output parameters of the aircraft are: pitch angle ϑ, angular velocity

, flight speed V, flight altitude H, rate of change in flight altitude

and pressure head q.

Primary information sensors measure and form the corresponding signals of these parameters:

- pitch angle meter 11 - pitch angle signal ϑ _and ;

- pitch angular velocity meter 12 — pitch angular velocity signal ω _zи ;

- flight speed sensor 16 - signal of flight speed V _and ;

;- altitude and rate of change of altitude sensor 13 - signals of the flight altitude N _and and its rate of change

;

- pressure sensor 9 - signal of the pressure head q _and .

являются координатными, по ним формируются основные контуры управления и стабилизации по высоте и углу тангажа.The measured signals ϑ _and , ω _{z and} , H _and and

are coordinate, they form the main contours of control and stabilization in height and pitch angle.

The signals q _and , V _and are parametric and form the channels of adaptive tuning parameters (gear ratios and restrictions) of the main circuits:

;- along the pitch contour - these are gear ratios in block 5 for pitch K _ϑ and angular velocity

;

, определенный блоком 10 по зависимости на фиг.3 для сигнала, сформированного контуром высоты в блоке 4, и перестройки передаточного числа

по скорости изменения высоты

в функции сигнала скорости полета V_и, сформированной блоками 15, 16 и 17.- along the height contour - this is the level of restriction

determined by block 10 according to FIG. 3 for a signal generated by a height contour in block 4 and gear ratio adjustment

by rate of change of height

as a function of the signal of flight speed V _and formed by blocks 15, 16 and 17.

The functioning of the control system is as follows.

Block 1 generates a signal of a given height H _ass. . The mismatch unit 2 generates a mismatch ΔН in the form:

where H _and is the signal from the height sensor and the rate of change of height 13.

. Суммирующий усилитель 3 формирует базовый сигнал управления контура высоты для подачи в контур управления по тангажу ϑ_у в виде:The mismatch ΔН is supplied to the summing amplifier 3, to the second input of which the signal component

. The summing amplifier 3 generates a basic control signal of the height loop for feeding into the pitch control loop ϑ _у in the form:

where K _n - gear ratio for the mismatch ΔН;

- компонента сигнала

, формируемая блоками 15, 17, 14 и датчиками 13 и 16.

- signal component

formed by blocks 15, 17, 14 and sensors 13 and 16.

формируется блоками 15, 16, 17 и составляет:In this case, the gear ratio

formed by blocks 15, 16, 17 and is:

where a _o is the base coefficient corresponding, for example, to a flight with an average speed V _o , i.e.

a V _and - measured flight speed at the output of the sensor 16.

Then (3) can be written as:

- базовый передаточный коэффициент по

, соответствующий скорости V_o.In (4) and (5)

- basic gear ratio

corresponding to the speed V _o .

The value a _{about is} set in the first reference signal setter 15.

. В первом блоке умножения 14 сигналы

и

(последний с датчика 13) умножаются и полученный сигнал

поступает в суммирующий усилитель 3.Relation (3) is formed in the first division block 17, i.e. at its output we have a signal corresponding to

. In the first block of multiplication 14 signals

and

(the last from sensor 13) are multiplied and the received signal

enters the summing amplifier 3.

в функции скорости полета V.Thus, according to (3) and (5), we see the adaptive adjustment of the gear ratio

as a function of flight speed V.

The feasibility and sufficiency of the proposed adaptation can be shown on the basis of the following considerations.

. При этом математически в операторной форме можно записатьThe height contour is formed by the signals of the aircraft H _and and

. In this case, mathematically in operator form, we can write

in the same time

where θ is the angle of inclination of the trajectory:

where α is the angle of attack.

скорость полета V является общим параметром, влияющим на процессы регулирования. Абсолютно корректным для инвариантности процессов регулирования к изменению скорости является введение в контур общего коэффициента, обеспечивающего инвариантность сквозных коэффициентов передачи к этому изменению, т.е. введение сомножителя в передаточные коэффициенты К_H и

, обратно пропорционального скорости полета V. Однако введение этого сомножителя в прямую цепь - по Н, т.е. в коэффициент К_H, существенно сказывается на статической точности, особенно в условиях применения реальных рулевых приводов, имеющих зону нечувствительности. Поэтому величина К_H выбирается К_H=const по двум соображениям:From (6) and (7) it can be seen that for the flight altitude control loop with a closure and, accordingly, regulation with respect to the coordinates H and

flight speed V is a common parameter that affects regulatory processes. Absolutely correct for the invariance of control processes to a change in speed is the introduction into the circuit of a common coefficient that ensures the invariance of the end-to-end transmission coefficients to this change, i.e. the introduction of the factor in the gear ratios K _H and

, inversely proportional to the flight speed V. However, the introduction of this factor into a direct chain is in H, i.e. in the coefficient K _H , significantly affects the static accuracy, especially in conditions of application of real steering gears with a dead zone. Therefore, the quantity K _H is chosen K _H = const for two reasons:

1) ensure accuracy;

, принятым по (3) и (5).2) ensuring sustainability and quality in combination with the coefficient

adopted by (3) and (5).

, которая позволяет соответственно ограничить угол атаки α в переходном процессе.The generated signal ϑ _{y is} limited in block 4 to a certain value

, which allows you to accordingly limit the angle of attack α in the transition process.

в функции от скоростного напора. Эта зависимость в общем виде представлена на фиг.3.In the general case, the aircraft has a tendency to large overruns at low speed heads, which determines the introduction of the corresponding nonlinear dependence of the block

in function of speed head. This dependence is generally presented in figure 3.

с учетом необходимого ограничения базового сигнала ϑ_у.Block 4 outputs a signal

taking into account the necessary limitations of the base signal ϑ _y .

является задающим для части системы управления по тангажу и сформированным на основе сигналов ϑ_и и ω_zи, соответствующих ϑ и ω_z летательного аппарата 8 и поступающих с датчиков 11 и 12.Signal

is the master for the pitch control system part and generated on the basis of the signals ϑ _and and ω _zи corresponding to ϑ and ω _{z of the} aircraft 8 and coming from the sensors 11 and 12.

, ϑ_и и ω_zи поступают в блок 5 для формирования сигнала управления угловой стабилизацией по тангажу. Выходом этого блока является базовый сигнал угловой стабилизации:So the signals

, ϑ _and and ω _zи enter block 5 for generating a pitch angle control signal. The output of this block is the basic signal of angular stabilization:

в функции скоростного напора q≅q_и. Действительно, задатчиками 18 и 23 определено базовое значение передаточных чисел

и

соответственно. Зависимость требуемого изменения (адаптации) передаточных чисел от q_и реализована во втором 19 и третьем 24 блоках деления: на второй блок деления 19 поступает сигнал

от блока 18 и q_и от блока 9, на его выходе формируется сигнал, соответствующий адаптированному передаточному числуThe block diagram of the functional Converter 5 is presented in figure 2. The combination of coordinate signals ϑ _and and ω _zи and parametric K _ϑ and

as a function of the velocity head q≅q _and . Indeed, the adjusters 18 and 23 determined the basic value of the gear ratios

and

respectively. The dependence of the desired changes (adaptation) of the gear ratios of q _and implemented in the second 19 and third 24 blocks fission: a second dividing unit 19 receives the signal

from block 18 and q _and from block 9, a signal corresponding to the adapted gear ratio is generated at its output

от блока 23 и q_и от блока 9, на выходе его формируется сигнал, соответствующий передаточному числу

с учетом адаптации:Accordingly, block 24 receives signals

from block 23 and q _and from block 9, a signal corresponding to a gear ratio is generated at its output

subject to adaptation:

In block 20, a mismatch Δϑ is formed:

и ω_zи, здесь сигнал

поступает с блока 24.In block 21 (the second multiplication block), a component of the error signal for the signal σ _in is formed equal to K _ϑ Δϑ, here the signal K _ϑ comes from block 19; in block 25, a signal component is formed by the angular velocity ω _z equal to

and ω _z , here the signal

comes from block 24.

The adder 22 is formed _{in the} signal σ:

The basic signal of stabilization σ _{in is} limited in block 6 (Fig. 1) to a technical level corresponding to the subsequent activation of steering drives for this channel (pitch-height) and taking into account the possibility of using these steering drives for adjacent channels (heading, roll).

в функции скоростного напора достаточно и оправдано, поскольку, во-первых, просто и экономично, а во-вторых, отражает инвариантность к нестационарному изменению аэродинамических характеристик летательного аппарата в функции от скоростного напора как доминирующего фактора.The construction of the channel adaptation gear ratios K _ϑ and

in the function of the velocity head, it is sufficient and justified, because, firstly, it is simple and economical, and secondly, it reflects the invariance to unsteady changes in the aerodynamic characteristics of the aircraft as a function of the velocity head as the dominant factor.

Thus, the proposed construction of an adaptive control system for the flight altitude of the aircraft allows you to expand the control capabilities of the aircraft and limit the angle of attack.

All blocks of the control system are standard and can be implemented on elements of automation and computer technology, for example, according to [5, 6], as well as program-algorithmically in a computer.

Information sources

1. RF patent No. 1751716, 07/30/92, cl. G 05 B 13/02.

2. Aerodynamics, stability and controllability of supersonic aircraft. / Ed. G.S. Byushgens. M.: Science, Fizmatlit, 1998, pp. 616-618.

3. V.A. Bodner. Theory of automatic flight control. M .: Nauka, 1964, p. 178.

4. I. A. Mikhalev and other systems of automatic control of the aircraft. M.: Engineering, 1987, p.30, 194.

5. A.U. Yalyshev, O. I. Razorenov. Multifunctional analog control devices for automation. M .: Mechanical Engineering, 1981, p. 103.

6. V. B. Smolov. Functional information converters. L .: Energy Publishing House, Leningrad Branch, 1981, p. 55.

Claims (2)

1. Adaptive control system for the flight altitude of the aircraft, containing the aircraft as a control object, steering gear, pitch angle meter, pitch angle meter and summing amplifier, characterized in that it contains a pressure head sensor, a height and speed sensor altitude, flight speed sensor, altitude adjuster, mismatch unit, first signal limiting unit, functional converter, second signal limiting unit, functional unit the signal limiting device, the first reference signal master, the first division unit and the first multiplication unit, the output of the height signal master through the mismatch unit is connected to the input of the summing amplifier, the first and second outputs of the height sensor and the rate of change of height are connected respectively to the second input of the mismatch unit and through the first multiplication unit - with the second input of the summing amplifier, the output of the first signal limiting block is connected through a series-connected functional converter and a second the signal limiting lock to the input of the steering gear, and the input of the first signal limiting block is connected to the output of the summing amplifier, the pressure head sensor is connected to the second input of the functional converter directly and through the signal limiting functional device with the second input of the first signal limiting block, the pitch angle meter and pitch angular velocity meter connected respectively to the third and fourth inputs of the functional transducer, the first master reference of the first signal and the flight speed sensor are connected to the first and second inputs of the first division block, respectively, and the output of the first division block is connected to the second input of the first multiplication block. 2. The adaptive flight altitude control system of an aircraft according to claim 1, characterized in that the functional converter comprises a second reference signal reference unit and a second division unit connected in series, a first adder, a second multiplication unit and a second adder, the output of which is the output of the functional converter connected in series to a third reference signal master, a third division unit and a third multiplication unit, the output of which is connected to the second input of the second sum ora, the output of the second division block is connected to the second input of the second multiplication block, with the first input of the first adder, the second input of the second division block, the second input of the first adder and the second input of the third multiplication block are the first, second, third and fourth inputs of the functional converter, respectively.

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