RU2047888C1 - Device for coordinated regulation of angular lateral movement of flying vehicle - Google Patents

Abstract

FIELD: instrumentation engineering. SUBSTANCE: device has roll channel and lateral g-load channel. Each channel is provided with subtraction element and summing amplifier. These members form control signals for roll and lateral g-load. Device is provided with sine function forming unit, cosine function forming unit, inverting amplifier, maneuvering jet engine control signal shaper, integrator, non-linear element with dead zone, two limiters, tangent function signal former, two correction function generators, three amplifiers, three multipliers, roll channel output summer, scaler and divider. EFFECT: extended functional capabilities at deficit of aerodynamic control. 5 dwg

Description

 The invention relates to the instrument-making industry and can be used in the development of on-board control systems (BSAU) for aircraft (aircraft) with an airplane circuit in conditions of large wind disturbances and coordinated turns with large roll angles.

 Analogs include automatic control systems [1, 2, 3] The yaw and roll control channels in the analogs contain subtraction elements and summing amplifiers that form control actions on the aircraft actuators according to the driving actions and signals of the state sensors.

 The disadvantage of analogues is the lack of solutions to the expansion of the flight area in conditions of wind gusts, leading to the achievement of aerodynamic control devices (ADU) braking articulated moments.

 Analogs also have the following disadvantages. The control system [1] presents the formation of lateral movement channels: yaw and roll separately in the decomposition plan and with a cross connection between them for the implementation of coordinated control; cross-connection is made in the form of chains from the rudder signal as the formation of a defining influence on the roll channel.

 The disadvantage of this implementation is the side loading of the yaw channel in it, caused by the introduction of the aircraft into the roll.

 The disadvantages should also include the lack of means of ensuring coordinate invariance with respect to slip (slip angle β and its derivatives) in a coordinated turn. Control systems for the lateral movement of aircraft along the roll and yaw channels are considered [2, 3]. In [2], the formation of the control law is also given, which ensures that the slip angle β is equal to zero. In this system, an excess component of the control signal in the yaw channel from the roll of the aircraft along the roll also occurs, weakening the effectiveness of coordinated control and reducing the accuracy characteristics of short-period and trajectory motion, which are significant disadvantages of the analogues.

 The closest in technical essence solution is the development of self-propelled guns [4] The device contains a series-connected roll channel subtraction element, the subtracting and summing inputs of which are respectively the device job input and feedback input for the roll angle of the device, and the total roll channel amplifier, the second input of which is feedback input on the angular velocity of the roll of the device, a series-connected element of subtraction of the side overload channel, subtracting and summing the inputs of are the input of the device reference and the feedback input of the device for lateral overload, and the summing amplifier of the side overload channel, the second input of which is the feedback input for the angular velocity of the yaw of the device.

 The disadvantage of the prototype is also the lack of means to expand the area of zero-control control of the main yaw channel during wind gusts, leading to the achievement of braking values of the hinge moments of the actuators ADU.

 The aim of the invention is to expand the functionality of the device with a deficiency of ADU.

 The solution to this problem will be the expansion of the flight zone in speed and altitude.

=ω_y+Z^β(β+β_w) +

cosθ·γ

=M

+M $\genfrac{}{}{0ex}{}{\text{β}}{\text{y}}$ (β+β_w)+M

=M

+M

+M $\genfrac{}{}{0ex}{}{\text{β}}{\text{x}}$ (β+β_w)+M

(1)

=ω_x-tgθ·ω_y
n_z

z^β(β+β_w)+z

+z

Уравнение шарнирных моментов руля направления (РН) имеет вид:
М_ш= М_ш ^β˙(β+β_w)+М_w ^δн˙δ_н, (2) здесь θ угол тангажа, характеризующий как параметр продольного канала для бокового;
β, ω_y, ω_x,γ,n_z координаты бокового движения ЛА: угол скольжения, угловая скорость рысканья, угловая скорость по крену, угол крена и боковая перегрузка, соответственно;
ρ_w возмущающий фактор ветровой порыв;
v скорость полета ЛА;
g ускорение свободного падения;
δ_н,δ_э углы отклонения исполнительных органов в каналах рысканья и крена, соответственно;
Z^β,М_y ^ωy,M_y ^β,M_y ^δн,М_х ^ωх,M_x ^ωy
М_х ^β, М_х ^δэ,М_ш ^β,M_ш ^δн- динамические коэффициенты. Из системы (1) видно, что член

cosθ·γ определяет движение координат ЛА, обусловленное введением ЛА в крен, т.е. координированным управлением.The linearized equations of the lateral movement of the aircraft along the dominant components, for example, in accordance with [5, 6] and taking into account the impact of the wind gust, can be represented by a system of differential equations in the form:

= ω _y + Z ^β (β + β _w ) +

cosθ γ

= M

+ M $\genfrac{}{}{0ex}{}{\text{β}}{\text{y}}$ (β + β _w ) + M

= M

+ M

+ M $\genfrac{}{}{0ex}{}{\text{β}}{\text{x}}$ (β + β _w ) + M

(1)

= ω _x -tgθ ω _y
n _z

z ^β (β + β _w ) + z

+ z

The equation of the hinged moments of the rudder (PH) has the form:
M _W = M _W ^β ˙ (β + β _w ) + M _w ^δ н˙δ _n , (2) here θ is the pitch angle, which characterizes as a parameter of the longitudinal channel for the lateral;
β, ω _y , ω _x , γ, n _z coordinates of the aircraft lateral movement: glide angle, yaw angular velocity, roll angular velocity, roll angle and lateral overload, respectively;
ρ _w disturbing factor wind gust;
v flight speed of the aircraft;
g acceleration of gravity;
δ _n , δ _{e the} deviation angles of the executive organs in the yaw and roll channels, respectively;
Z ^β , M _y ^ω y, M _y ^β , M _y ^δ n, M _x ^ω x, M _x ^ω y
M _x ^β , M _x ^δ e, M _W ^β , M _W ^δ n - dynamic coefficients. From system (1) it can be seen that the term

cosθ · γ determines the motion of the coordinates of the aircraft due to the introduction of the aircraft into the roll, i.e. coordinated management.

0), получим, что установившееся значение угловой скорости ω_yγ от воздействия γ составит

=

cosθ·γ (3)
Сигналы управления в канале крена (элеронов) и рысканья (руля направления) формируются в виде
σ_кр ^o=a_γ(γ-γ_k)+a_ωx˙ω_x, (4)
σ_н ^o=a_nz(n_zk-n_z)+a_ωy(ω_y+σ_k), где a_γ,a_ωx,a_nz и a_wy соответствующие передаточные числа;
γ_k и n_zk управляющие сигналы по крену и перегрузке, соответственно.It can also be seen from (1) that, considering the decomposition motion with respect to the action of γ and also evaluating the lateral motion with allowance for the assumption of small slip (β,

0), we obtain that the steady-state value of the angular velocity ω _yγ from the action of γ is

=

cosθ γ (3)
The control signals in the roll channel (ailerons) and yaw (rudder) are formed in the form
σ _cr ^o = a _γ (γ-γ _k ) + a _ωx ˙ω _x , (4)
σ _n ^o = a _nz (n _zk -n _z ) + a _ωy (ω _y + σ _k ), where a _γ , a _ωx , a _nz and a _{wy are the} corresponding gear ratios;
γ _k and n _zk control signals for roll and overload, respectively.

This process corresponds to the establishment of the signal ω _y = ω _{yγ of the} induced signals γ ≠ 0.

In conditions of multi-mode control, in the presence of large levels of command signals γ _k , n _zk and disturbing factors β _w , a control deficit arises due to the achievement of the hinged moments of the braking value. This is especially acute in the steering gear of the launch vehicle, which is much less aileron in terms of aircraft efficiency and quite loaded in terms of design and operating factors. Design features include, for example, the presence of an air brake (VT) in modern aircraft such as aerospace aircraft (VKS), made technologically combined with the rudder; the disclosure of VT leads to a decrease in the capabilities of the LV, in particular, to a decrease in the braking values of the hinge moments. One of the means of expanding control capabilities is the use of additional resource management sources of maneuverable engines (MD) of a reactive control system (DCS) [7]
However, such direct involvement of the MD is also limited by the MD's own resource and the conditions of their work: duration, frequency of inclusion, and maintenance of necessary pauses. The indicated limitations in such a solution determine its major drawback and may make it unacceptable, which requires the development of means for the correct management of MD.

 The goal stated above with the resolution of the DCS factors is achieved as follows. The following composition is introduced into the device described previously based on the prototype.

 A sine function generating unit, a first amplifier and a first multiplier connected to a feedback angle input of the device in series, a cosine function forming unit, a divider and a second amplifier connected to a feedback input of a pitch pitch of a device, the output of which is connected to the second input of the first multiplier , the control input of the divider is connected to the input by the flight speed of the device, the output of the first multiplier is connected to the third input of the summing channel amplifier side overload, as well as a series-connected inverting amplifier, the input of which is connected to the output of the summing amplifier of the side overload channel, a driver for controlling maneuverable jet engines, the output of which is the output of a device for controlling jet engines, an integrator, a nonlinear element with a dead band, a first limiter and an output a side overload channel subtraction element, the second input of which is connected to the output of the summing amplifier of the side lane channel load, and the output is the output of the device for controlling the aerodynamic drive of the rudder, the tangent function signal generating unit, the first corrective functional converter and the third amplifier, the second multiplier, the second limiter and the roll bank output adder, the second input of which is connected to the output of the summing channel amplifier roll, and the output is the output of the device for controlling the aerodynamic drive of electrons, and the scale amplifier is connected in series Tel, whose input is connected with the input of airspeed device, the second correction function generator and a third multiplier, a second input coupled to an output of a nonlinear element with a dead zone, and an output to a second input of the second multiplier.

 Comparative analysis with the prototype shows that the claimed device differs from it in the introduction and composition of new channels and common links. Thus, the claimed device meets the criteria of the invention of "novelty."

от угла γ или γ_k, т.е. сформирован канал c σ_к

cosθ·sinγ·K_к (5) где сигнал sinγ определяет полную проекцию подъемной силы на горизонтальную плоскость [5] (линеаризацию синусной функции определяет угловое приращение γ ); К_к коэффициент, определяющий степень усиления сигнала γ для компенсации ω_yγ и достижения Inv β

Так, при К_к=1, получим сигнал, компенсирующий компоненту (3) с точностью до динамики и установившегося отклонения. Формируется сигнал с усилением для синхронизации сигналов σ_k и ω_yγ т.е. с учетом отличия от установившегося состояния по (4) и для достижения технически потребной инвариантности.An additional component of the yaw channel control signal σ _{k is} formed in the device, taking into account the compensation of the dominant signal ω _yγ according to (3) and the compensation of occurrence

from the angle γ or γ _k , i.e. channel c σ _to

cosθ · sinγ · K _to (5) where the signal sinγ determines the full projection of the lifting force on the horizontal plane [5] (the linearization of the sinus function determines the angular increment γ); K _to coefficient determining the degree of amplification of the signal γ to compensate for ω _yγ and achieve Inv β

So, when K _k = 1, we get a signal that compensates component (3) up to dynamics and steady-state deviation. A gain signal is generated to synchronize the signals σ _k and ω _yγ i.e. taking into account the differences from the steady state according to (4) and to achieve the technically necessary invariance.

0,9÷1,15), где особо выражены экстремумы и смена полярности (знакопеременных) динамических коэффициентов ЛА что важно и для динамики процессов. Знак и уровень отклонения элеронов выбраны в связи с изложенным с точки зрения минимизации β. Анализ системы уравнений (1) показал с принятой точки зрения зависимость требуемого отклонения элеронов от величины tgθ также нашедшую отражение в сформированном техническом решении.Direct channel of DCS: an inverting amplifier-driver of a control signal of choice MD is introduced by analogy with known solutions, for example, according to [7]
However, such a solution, as noted earlier, has limited application, because the constant involvement of the MD DCS is unacceptable. The developed and introduced additional channels are aimed at a flexible, comfortable and strictly dosed application of MD DCS along with ADUs while satisfying the conflicting requirements of the limited inclusion of MD and the effectiveness of DCS. Indeed, the integrated channel forms the integral signal of the operating time of the DCS engines and the transfer of the integral component from the DCS to the ADU, in turn, this transmission is formed for two ADU channels: via the LV and ailerons. With limited capabilities of the launch vehicle (small braking torque), the integrated component of the DCS is transmitted only to ailerons and the dynamics of the use of ailerons are greater, the less the capabilities of the launch vehicle. Monitoring M according to (2), taking into account the knowledge that there is a greater value of the first component (2) with respect to β, requires minimizing the angle β, which is especially important in the transonic region of the flight (Mach numbers correspond to M

0.9 ÷ 1.15), where extremes and a change in the polarity (alternating) of the dynamic coefficients of the aircraft are particularly pronounced, which is also important for the dynamics of processes. The sign and level of deviation of the ailerons are selected in connection with the above from the point of view of minimizing β. Analysis of the system of equations (1) showed, from the accepted point of view, the dependence of the required deviation of the ailerons on the value of tgθ also found reflection in the generated technical solution.

 Indeed, we consider the system (1) for the steady state from the point of view of the introduced minimization criterion β for the conditions for the activation of the DCS in the LV channel with the introduction of the cross integrated component of the DCS in the LV channels and ailerons. For research under these conditions, it is correct to accept the deviation of the steering elements of the ADU equal to the signals, i.e.

δ _e = σ _cr ,
δ _n = σ _n , and also assuming that σ _n is a direct integral signal from the DCS, and σ _cr, we set it as dependent on σ _n , i.e.

σ _cr = K _f ˙σ _n , where K _f functional coefficient.

cosθ·γ=0
-М_y ^ωy ˙ω_y-M_y ^β˙β=M_y ^δн ˙δ_н (6)
-М_x ^ωy ˙ω_y-M_x ^ωx ˙ω_x-M_x ^β˙β=M_y ^δэ˙К_ф˙δ_н,
-tgθ˙ω_y+ω_x= 0. Так как ω_x=tgθ˙ω_y и учтя, что 2-е и 3-е уравнения определят ω_y и β а 3-е γ рассматриваем 2-е и 3-е уравнения:
-М_y ^ωy ω_y-M_y ^β˙β=M_y ^δн ˙σ_н
-(М_х ^ωх tgθ+M_x ^ωy )ω_y-M_x ^β˙β= M_x ^δэ ˙К_ф˙σ_н (7)
Из (7) вспомогательный определитель по β равен: Δβ

· σ_н (8)
Δβ=0 при
-M_y ^ωy M_x ^δэ K_ф+M_y ^δн M_x ^ωx tgθ+M_y ^δн M_x ^ωy0.In the proposed analysis, the signal σ _{n is} considered as a perturbation in this way, and the dominant output coordinate β _mouth . Thus, in estimates of forced motion, system (1) takes the form
ω _y + z ^β

cosθ γ = 0
-M _y ^ω y ˙ω _y -M _y ^β ˙β = M _y ^δ n ˙δ _n (6)
-M _x ^ω y ˙ω _y -M _x ^ω x ˙ω _x -M _x ^β ˙β = M _y ^δ e˙K _f ˙δ _n ,
-tgθ˙ω _y + ω _x = 0. Since ω _x = tgθ˙ω _y and taking into account that the 2nd and 3rd equations determine ω _y and β and the 3rd γ we consider the 2nd and 3rd equations:
-M _y ^ω y ω _y -M _y ^β ˙β = M _y ^δ n ˙σ _n
- (M _x ^ω x tgθ + M _x ^ω y) ω _y -M _x ^β ˙β = M _x ^δ e ˙K _f ˙σ _n (7)
From (7), the auxiliary determinant with respect to β is equal to: Δβ

Σ _n (8)
Δβ = 0 for
-M _y ^ω y M _x ^δ e K _f + M _y ^δ n M _x ^ω x tgθ + M _y ^δ n M _x ^ω y0.

tgθ (10) Из (10) следует явная зависимость К_ф и signК_ф от θ. Для анализа и синтеза коэффициента К_ф выражение (10) преобразует к виду с разделением характерных членов по эффективности и демпфированию: K_ф

tgθ +

(11)
Благодаря такому разделению видна достаточность в явном виде зависимости К_ф от tgθ и можно пояснить достаточность такой реализации с грубостью к скоростному напору и скорости в числах Маха, что важно с точки зрения минимума реализационных средств на борту ЛА. Действительно, в соответствии, например, с [6] и [5] динамические коэффициенты объекта прямо пропорциональны соответствующим производным аэродинамических коэффициентов и обратно пропорциональны скоростному напору, т.е. в доминирующем плане K_ф

tgθ +

(12) где m_i ^j(M) производные i-х аэродинамических моментов по j-м параметрам.(nine)
Therefore, K _f

tgθ (10) From (10) there follows an explicit dependence of K _f and signK _f on θ. For analysis and synthesis of the coefficient K _f, expression (10) transforms to a form with the separation of characteristic terms by efficiency and damping: K _f

tgθ +

(eleven)
Due to this separation, the sufficiency of the dependence of K _f on tgθ is clearly visible and the sufficiency of such an implementation with coarseness to the velocity head and speed in Mach numbers can be explained, which is important from the point of view of the minimum implementation means on board the aircraft. Indeed, in accordance with, for example, [6] and [5], the dynamic coefficients of an object are directly proportional to the corresponding derivatives of the aerodynamic coefficients and inversely proportional to the velocity head, i.e. dominantly K _f

tgθ +

(12) where m _i ^j (M) are derivatives of i-aerodynamic moments with respect to j-th parameters.

K₁

K₂ (13)

K₃ тогда К_ф= К₁(К₂tgθ+К₃)=К₁К₂tgθ +К₁К₃=Аtgθ+B (14) где А=К₁К₂, В=К₁К₃.Moreover, based on the sufficient correlation of similar coefficients:
efficiencies m _y ^δ n and m _x ^δ e
and damping m _x ^ω xm _y ^ω y and m _x ^ω y can be obtained quite convenient for the limited possibilities of implementing the relationship by setting:

K ₁

K ₂ (13)

K ₃ then K _f = K ₁ (K ₂ tgθ + K ₃ ) = K ₁ K ₂ tgθ + K ₁ K ₃ = Аtgθ + B (14) where A = K ₁ K ₂ , B = K ₁ K ₃ .

 Signs that distinguish the claimed technical solution from the prototype are not identified in other technical solutions in the study of this and related areas of technology and, therefore, meet the criterion of "significant differences".

 The invention is illustrated in FIG. 1-5.

 In FIG. 1 is a block diagram of a device as part of an aircraft lateral motion control system, i.e. object, actuators and information sensors circled by a dotted line to the right. The device has the following composition: roll channel subtraction element 1, summing roll channel amplifier 2, side overload channel subtraction element 3, lateral overload channel amplifier 4, sine function generating unit 5, first amplifier 6, first multiplier 7, cosine generating unit 8 , divider 9, second amplifier 10, tangent function signal conditioning unit 11, first corrective functional converter 12, third amplifier 13, second multiplier 14, second limiter 15, output adder channel and bank 16, a large-scale amplifier 17, a second corrective functional converter 18, a third multiplier 19, an inverting amplifier 20, a driver signal for maneuvering jet engines 21, an integrator 22, a nonlinear element with a dead zone 23, the first limiter 24, the summing amplifier of the side overload channel 25 .

The composition of the managed object is as follows: aerodynamic drive of ailerons 26, aerodynamic drive PH 27, block of jet engines 28, the object itself as a dynamic link 29, roll angle sensor 30, roll angle sensor 31, lateral overload sensor 32, yaw rate sensor 33, pitch angle sensor 34, flight speed sensor 35. In the block diagram of FIG. 1 denotes: ω _{x the} measured roll angular velocity signal, γ _{k the} roll command command signal, γ is the roll angle signal, σ _cr ^o , σ _cr the control signals of the roll channel, δ _e the aileron deviation, n _z for lateral overload, n _z is the measured signal of lateral overload, ω _y is the measured signal of the angular velocity in the yaw channel, θ is the measured signal of the pitch angle, v is the measured signal of flight speed, σ _n ^o , σ _n are the control signals of the side overload channel, σ _to compensatory I component of the manipulated variable signal, δ _n deviation RN, Δγ mismatch in roll channel, Δn _z mismatch channel overload, σ _n ^θ intermediate signal tgθ component; σ _cr ^and, σ _n ^and integral components of signals in the roll channel and overload, respectively, , M _{y is the} moment created by the MD DCS, U is the jet engine control signal, M is the flight speed signal in Mach numbers; ± σ ₁ ; ± σ ₂ ; ± ρ ₁ ; ± U ₁ , ± U _N signals of the settings set directly at the corresponding links (15; 24; 23; 21); β _w disturbance acting on the aircraft (wind gust); To (M) the signal corresponding to the gear ratio. Blocks 22 and 23 are circled by a dotted line as constituting an integrating device. In FIG. 2 is a graph of the dependence of N _y (U) of block 21 in the form of a multi-stage relay characteristic that reflects the sequence of connecting maneuverable jet engines as the control settings exceed U _i (N is the number of engines on each side of the aircraft, determining the right and left turns of the aircraft yaw).

In FIG. 3 is a block diagram of an implementation of a motor control signal generator 21, taking into account the characteristics of FIG. 2. Here it is indicated: 36i relay elements with a deadband U _i, respectively; 37 adder.

 In FIG. 4 and 5 are examples of piecewise linear dependency graphs implemented by transducers 12 and 18.

, где v скорость полета, измеренная датчиком 35, 10 усиление сигнала

с множителем g, 6 потребное усиление сигнала sinγ т.е. на его выходе имеем К_к˙sinγ 7 умножение сигналов K_к·sin

cosθ=σ_к На выходе усилителя 4 имеем сигнал σ_н ^o. Передаточные числа a_nz и a_ωyвыставляются в усилителе 4. Сигнал σ_кр ^o является основной компонентой сигнала управления аэродинамическим приводом в канале крена элеронами 26. Сигнал σ_н ^o является основной компонентой сигнала упpавления аэродинамическим приводом РН 27 и полным сигналом управления МД 28. Сигнал управления двигателями N_y формируется блоками 20 и 21. Инвертирующий усилитель 20определяет согласование по знаку разворотов ЛА от аэродинамического δ_н и реактивного М_y воздействий, ибо во 2-е уравнение системы (1) в правую часть добавлен член + ε_y, характеризующий ускорение от РСУ, соответствующее М_y, а член от АДУ (M_y ^δн ˙δ_н) при положительном δ_н (и соответственно σ_н ) отрицателен, ибо, как было отмечено ранее, для ЛА M_y ^δн <0; сигнал U с выхода блока 20 поступает на блок 21, с выхода которого сигнал N_y, формируемый, например, по зависимости фиг. 2 в реализации блоками 36 и 37 фиг. 3 поступал в виде дискретного уровня, соответствующего выборке+1, +2, +3,+N, -1, -2, -3, -N} непосредственно на включение соответствующих двигателей блока двигателей 28.The device operates as follows. The main components of the control signals in the roll channels σ _cr ^o and overload σ _n ^o are formed on the basis of the measured signals γ, ω _x , n _z , ω _y , θ of the sensors 30-34 installed on the object 29, and specifying the effects γ _to and n _zк in the following way. The subtraction element of the roll channel 1 highlights the mismatch Δγ = γ-γ _k . The summing amplifier of the roll channel 2 generates a signal σ _cr according to the first equation (4). The transmission coefficients a _γ and a _{ωx are} set in the amplifier 2. The control signal in the yaw channel σ _n ^o is formed according to the second equation (4). Subtraction element 3 forms a mismatch Δn _z = n _zк -n _z , while in the lateral movement of a significant class of aircraft, n _zк = 0 stabilize, then Δn _z = n _z . The signal component σ _k is formed by blocks: 5 forms sinγ- 8 cos θ 9 forms

where v is the flight speed measured by the sensor 35, 10 signal gain

with the factor g, 6 the required signal amplification sinγ i.e. at its output we have K _to ˙sinγ 7 the multiplication of signals K _to · sin

cosθ = σ _к At the output of amplifier 4 we have a signal σ _n ^o . The gear ratios a _nz and a _{ωy are} set in the amplifier 4. The signal σ _cr ^o is the main component of the control signal of the aerodynamic drive in the roll channel ailerons 26. The signal σ _n ^o is the main component of the control signal of the aerodynamic drive PH 27 and the complete control signal MD 28. The signal motor control N _y is formed by blocks 20 and 21. The inverting amplifier 20opredelyaet matching sign reversals of aircraft aerodynamic δ _n and M _y jet impacts, because in 2nd system equation (1) added to the right side tsp n + ε _y, characterizing the acceleration of CSF, M corresponding to _y, and a member of the ROV (M _y ˙δ _n ^δ n) with δ _n positive (and hence σ _n) is negative, since, as previously noted, for aircraft M _y ^δ n <0; the signal U from the output of block 20 enters the block 21, from the output of which the signal N _y , formed, for example, according to FIG. 2 as implemented by blocks 36 and 37 of FIG. 3 arrived in the form of a discrete level corresponding to a sample of + 1, +2, + 3, + N, -1, -2, -3, -N} directly to turn on the corresponding engines of engine block 28.

M

dt

M

dt (15)
Однако эффективным задействование двигателей РСУ в комбинации с АДУ целесообразно считать при отработке достаточно больших сигналов, в связи с чем экономия топлива и ресурса по наработке самих двигателей становится дополнительной актуальной задачей.At the same time, the signal N _y is fed to the integrator 22, from the output of which the integral signal is fed to the link with the deadband 23. The integrated channel corrects the operation of the DCS channel, controlling and optimizing it. As an additional reference to the emerging problem of this optimization, we can indicate [8] where it is said about the process optimality from the point of view of the flow of the working fluid with the admissibility of a one-sided oscillatory cycle determining the compensation of the perturbation f _w (in the form of moment M _c ), i.e. by [8]

M

dt

M

dt (15)
However, it is advisable to consider engaging DCS engines in combination with ADUs when developing sufficiently large signals, in connection with which saving fuel and a resource for operating the engines themselves becomes an additional urgent task.

The main solution of the goal of the invention is to transfer the integral "processing" of DCS engines to ADU bodies. This is done by the integrator 22 and the link with the deadband 23. If the integral signal ∫N _y dt N _y dt is exceeded in terms of the zone level ρ _{1, the} difference signal σ _n is supplied to the ailerons and LV ADUs. It is fed directly to the LV channel with a preliminary restriction on the limiter 24 with a restriction level of ± σ ₂ corresponding to the efficiency of one or two UDs so as not to limit the activation of the DCS, but to compensate for the residual oscillatory cycles. In the summing amplifier of the side overload channel 25, the output signal of this channel is formed
σ _n = σ _n ^o -σ _n ^and , (16) where σ _n ^{and the} signal from the output of the limiter 24. The minus sign reflects the repeated inversion of the DCS signal to match the signal σ _n ^o . The output signal of the ADS RN σ _n is supplied to the steering gear of the RN 27 and the RN is processed, it is deflected by an angle δ _n , as a result of which the process of lateral movement of the aircraft 29 is regulated.

где v скорость полета, измеряемая датчиком 35, а_ср средняя величина скорости звука в области полета (варьирование скорости звука невелико, и может быть учтено при формировании характеристики блока 18 некоторым снижением коэффициента этого блока). Сигнал М поступает на преобразователь 18, характеристика которого (фиг. 5) отражает зону передачи σ_и на элероны: М₂-М₃ активный участок передачи; М₁-М₂ и М₃-М₄ участки сопряжения. С преобразователя 18 сигнал К(М) соответствует передаточному коэффициенту К_о(М) для сигнала σ_и: на умножителе 19 умножаются сигналы К(М) и σ_и с блоков 18 и 23, соответственно. Блоки 11, 12 и 13 отвечают решению поставленной задачи с точки зрения формирования компоненты σ_кр ^и в функции от tgθ в соответствии с (14): с выхода блока 11 снимается сигнал tgθ на преобразователе 12 формируется кусочно-линейная функция σ_и ^θ(tgθ)в соответствии с фиг. 4, при этом А и В параметры, соответствующие (14); усилитель 13 установлен для возможности повышения эффективности элеронов в зоне ослабления РН.Integral component CSF _and σ is also fed into the roll to form a channel therein and σ _cr control components ^and as a parallel impact to RN and only aileron control at full load ROP. This component is formed, as noted earlier, from the point of view of decreasing the sliding angle β and lateral overload in a special flight zone, where the dynamic characteristics of an aircraft are sharply changing, for example, in a transonic one (M = 0.9-1.15). This zone is formed by links 17 and 18. A signal corresponding to the flight speed in Mach numbers M =

where v is the flight speed measured by the sensor 35, and _{cp is the} average value of the speed of sound in the flight area (the variation in the speed of sound is small, and can be taken into account when forming the characteristics of block 18 by a certain decrease in the coefficient of this block). The signal M enters the transducer 18, the characteristic of which (Fig. 5) reflects the transmission zone σ _and to the ailerons: M ₂ -M _{3 the} active portion of the transmission; M ₁ -M ₂ and M ₃ -M ₄ mating sections. Since the inverter 18 K signal (M) corresponds to the gear ratio _of K (M) for signal _and σ: a multiplier 19 for multiplying the signals of K (M) and σ _and blocks 18 and 23, respectively. Blocks 11, 12, and 13 correspond to the solution of the problem from the point of view of forming the component σ _cr ^and as a function of tgθ in accordance with (14): the signal tgθ is removed from the output of block 11 on the converter 12 and a piecewise linear function σ _and ^θ (tgθ) is formed in accordance with FIG. 4, with A and B parameters corresponding to (14); the amplifier 13 is installed to increase the efficiency of the ailerons in the area of weakness of the pH.

On the multiplier 14, the signals of the channels tgθ and K (M) are multiplied. The signal from the output of the multiplier 14 is fed to the limiter 15, the level of limitation of which is ± σ ₁ , as well as ± σ _{2 of the} limiter 24, corresponds to the degree of involvement of the ailerons by equivalence to one or two DCS engines.

The signal σ _cr ^and from the output of the limiter 15 is fed to the output adder of the roll channel 16, the output of which the signal σ _cr = σ _cr ^o + σ _cr ^and is fed to the steering gear 26, deflecting the ailerons at an angle δ _e .

 The proposed construction of the device significantly expands the control capabilities over the high-speed range of flight paths of the aircraft, the solution allows the correct combination of ADU and DCS with economical fuel consumption and engine life. As a result of the implementation of the proposal, increased static and dynamic accuracy. The achieved positive effect is confirmed by theoretical analysis, as well as the results of modeling on a full model of an object, actuators and formed control laws.

 The component parts of the device are functionally complete blocks and can be implemented in special. the calculator directly according to the functions performed, and in the analog version using modern elements of automation and computer technology, for example, according to the sources [9, 10, 11] In particular, the implementation of multipliers and dividers according to [9] of the signal generation blocks of trigonometric functions according to [10] corrective functional converters according to [12]

Claims (1)

 DEVICE FOR COORDINATED REGULATION OF ANGULAR LATERAL FLIGHT OF THE AIRCRAFT, containing series-connected element of subtraction of the roll channel, subtracting and summing the inputs of which are the input of the device reference and the feedback input for the roll angle of the device, and the summing amplifier of the roll channel, the second input of which is the input the angular velocity of the roll of the device, a series-connected element of subtraction of the side overload channel, subtracting and summing the input the inputs of which are the input of the device reference and the feedback input of the device for lateral overload, and the summing amplifier of the side overload channel, the second input of which is a feedback input for the angular velocity of the yaw of the device, characterized in that they are connected to it in series connected to the feedback input on the roll angle of the device, a sine function forming unit, a first amplifier and a first multiplier connected in series to a pitch feedback input the device is a cosine function forming unit, a divider and a second amplifier, the output of which is connected to the second input of the first multiplier, the control input of the divider is connected to the input by the flight speed of the device, the output of the first multiplier is connected to the third input of the summing amplifier of the side overload channel, as well as an inverting amplifier connected in series the input of which is connected to the output of the summing amplifier of the side overload channel, the driver of the control signal for maneuverable jet engines, which is the output of the device for controlling jet engines, an integrator, a nonlinear element with a dead zone, a first limiter and an output element for subtracting the side overload channel, the second input of which is connected to the output of the summing amplifier of the side overload channel, and the output is the output of the device for controlling the aerodynamic rudder connected in series to a tangent function signal conditioning unit, a first corrective functional converter, a third an amplifier, a second multiplier, a second limiter, and an output adder of the roll channel, the second input of which is connected to the output of the summing amplifier of the roll channel, and the output is the output of the device for controlling the aerodynamic drive of ailerons, and a scale amplifier, the input of which is connected to the input of the device by flight speed , a second corrective functional converter and a third multiplier, the second input of which is connected to the output of the nonlinear element with a deadband, and the output from th input of the second multiplier.

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