RU2124688C1 - Method of combined control of flying vehicle - Google Patents

Abstract

FIELD: control of unmanned flying vehicles; control of artillery projectiles with homing heads whose flight trajectory consists of three sections: ballistic section, section of flight at constant angle of inclination to longitudinal axis to level (gliding section) and homing section. SUBSTANCE: method consists in limiting angles of inclination of control members (for example, aerodynamic control surfaces) which is effected in its turn due to limiting the input signal to control surface drive in gliding section. Level of signal has magnitude of control signal required for motion of projectile over reference trajectory which is below real trajectories. EFFECT: reduction of the dispersion of artillery projectile in gliding section arising to end of ballistic section of flight due to bringing gliding trajectory to reference trajectory before beginning of homing. 4 dwg

Description

 The invention relates to the field of control of unmanned aerial vehicles (LA) for military purposes, in particular to the management of artillery shells with homing heads (GOS) and allows to reduce the dispersion of shells formed on the ballistic section of the flight, which, in turn, increases the probability of hitting a target.

 The probability of hitting the target of shells with GOS, first of all, depends on the accuracy of the output of shells into the zone of capture of GOS radiation from the target.

 Homing shells with a long firing range (15 - 20 km) the main part of the flight time moves along a ballistic trajectory, i.e. are managed. As a result of this, scattering of shells inevitably occurs due to technical dispersion (due to variations in drag coefficient, angle and launch velocity, projectile mass) and firing preparation errors (errors in determining firing range, propellant charge temperature, wind, pressure and air temperature, method of calculating settings, anticipation of the movement of the target, etc.).

A known method of bringing a projectile to the target target area of the GOS, proposed at the stage of preliminary development of the Copperhead artillery shell with a semi-active laser GOS, including launching a projectile from a launcher (155 mm howitzer) at an angle θ ₀ to the horizon with an initial speed of V ₀ , ballistic flight and homing in the final section of the trajectory until the projectile hits the target [1].

 The method has the disadvantage that when firing at long ranges along a ballistic trajectory, such projectile heights are set that exceed the heights of the lower cloud cover, i.e. between the GOS and the target, a cloud interference is established. This interference does not allow optical GOS to receive radiation from the target, thereby significantly reducing the capture range of the GOS target.

 Small values of the target range of the GOS target and large values of shooting errors (primarily technical dispersion of shells) do not allow for a high probability of hitting the target.

The Copperhead artillery projectile guidance method proposed at the engineering stage [2] allows increasing the target range of the GOS target, which includes launching the aircraft from the launch device at an angle θ ₀ to the horizon with an initial speed V ₀ , ballistic flight, flight section with a given pitch angle ( inertial guidance area) and homing before the projectile hits the target.

 The use of inertial guidance in the considered method makes it possible to reduce projectile flight altitudes when firing at long ranges (i.e., to provide more gentle flight paths until the target is captured), thereby increasing the target's target range. In addition, the inertial guidance section allows you to increase the range of the projectile by 25% compared with the method described in [1].

 However, the method according to [2] has a significant drawback, which is that the technical dispersion of the shells formed at the end of the ballistic flight section will be preserved throughout the inertial guidance section until the start of homing, since the projectile flight in the inertial guidance section (planning) is carried out at a constant, predetermined angle of inclination of the longitudinal axis of the projectile, which is provided by the pitch stabilization system, see FIG. 1.

 Large values of the technical dispersion of the projectile, formed at the end of the ballistic section of the trajectory, preserved on the inertial guidance section until the start of homing, reduce the likelihood of hijacking the target of the seeker, reduce the likelihood of the projectile falling into the region of initial misses selected during the homing process, which generally reduces the probability of hitting the target .

где

δ_m - максимальный угол отклонения рулей;
K₁ - коэффициент передачи рулевого привода;
n_p - расчетная поперечная перегрузка в высшей точке баллистического полета при запуске под углом θ₀-3τ_θ0 с начальной скоростью V₀-3σV₀;

среднеквадратическое отклонение угла запуска снаряда;

среднеквадратическое отклонение скорости запуска снаряда;
ϑ_пр - программный угол наклона продольной оси снаряда;
ϑ - текущий угол наклона продольной оси снаряда;
K - коэффициент передачи датчика угла тангажа.The proposed method of combined control, which allows to increase the probability of hitting a target due to a decrease in technical dispersion due to a programmed change in the angle of inclination of the longitudinal axis of the projectile to the horizon in the inertial guidance section, includes launching the projectile from the launch device at an angle θ ₀ to the horizon with an initial velocity V _{0 of} guiding the projectile along ballistic trajectory, inertial guidance of the projectile with a programmed angle of inclination of the longitudinal axis of the projectile to the horizon, homing to direct projectile hitting a target, characterized in that the value U _Rin control signals to the steering actuator projectile entering the inertial guidance portion is calculated in accordance with the dependencies:

Where

δ _m - the maximum angle of deviation of the rudders;
K ₁ - gear ratio of the steering gear;
n _p is the calculated transverse overload at the highest point of a ballistic flight when launched at an angle θ ₀ -3τ _θ0 with an initial speed of V ₀ -3σV ₀ ;

standard deviation of the projectile launch angle;

standard deviation of the projectile launch velocity;
ϑ _pr - program angle of inclination of the longitudinal axis of the projectile;
ϑ - current angle of inclination of the longitudinal axis of the projectile;
K is the gear ratio of the pitch angle sensor.

 Reducing the technical dispersion of the projectile by the time the homing begins is achieved by reducing the projection planning trajectory to the reference trajectory.

 As a reference trajectory from the entire set of trajectories determined by the technical dispersion of the projectile, one is selected at which the lowest flight heights are established, i.e. real projectile flight paths on the ballistic section will always be located above the reference flight path.

 Technically, a programmatic change in the angle of inclination of the projectile to the horizon is carried out by introducing restrictions on the angles of deviation of the projectile controls (for example, aerodynamic rudders), which, in turn, are achieved through restrictions imposed on the level of the input signal to the steering gear.

The magnitude of the limitation of the level of the input signal to the steering gear (U _ogre ) is determined by the value of the input signal required for the projectile to move along the reference path. This value of the input signal is determined by the disposable overload of the projectile and provides compensation for acceleration of gravity and inertial guidance.

In real launches, projectiles moving along trajectories that exceed the reference one have less overload due to lower air density. Accordingly, for planning with a given angle of inclination of the trajectory at a flight height greater than the height of the reference path, a large control command is required. The limitation of the command leads to an undercompensation of gravity and, as a consequence, curvature of the trajectory in the direction of the reference (Fig. 2). As you approach the supporting trajectory, the velocity of the projectile and the density of air increase, and therefore, the overload also increases. This process continues until the overload developed by the action of the U _ogre command compensates for the effect of acceleration of gravity. This will happen when the projectile is on the supporting trajectory or in the immediate vicinity of it. Further, the movement of the projectile occurs near or along this trajectory.

 To illustrate the stated control method, we give an example.

начальная скорость ЛА,

угол пуска θ₀ = (45±15)^o;
δ_m = 10^o; K₁ = 5^o/В; K = 1 В/0;
угол наклона траектории на участке инерциального наведения θ_пр = -25^o;
Параметры траектории на баллистическом участке могут быть определены по зависимостям [3, стр. 38, 74]:

ϑ‗‗θ (2.6)
где
V_s, Y_s - скорость и высота в вершине траектории.Initial data:
aircraft mass, m = 50 kg;
mid-sectional area, S = 0.0181 m ² ;
drag coefficient, C _x = 0.4;
lift coefficient, C $\genfrac{}{}{0ex}{}{\text{α}}{\text{y}}$ = 27;
coefficient of stabilizing moment, m $\genfrac{}{}{0ex}{}{\text{α}}{\text{z}}$ = -1.2;
coefficient of control torque, m $\genfrac{}{}{0ex}{}{\text{δ}}{\text{z}}$ = 0.6;
air density at sea level, ρ ₀ = 1.226 kg / m ³ ;
acceleration of gravity,

aircraft initial speed

starting angle θ ₀ = (45 ± 15) ^o ;
δ _m = 10 ^o ; K ₁ = 5 ^o / V; K = 1 V / 0;
the angle of inclination of the trajectory in the area of inertial guidance θ _CR = -25 ^o ;
The trajectory parameters on the ballistic section can be determined by the dependencies [3, p. 38, 74]:

ϑ‗‗θ (2.6)
Where
V _s , Y _s - speed and height at the top of the trajectory.

Параметры траектории на участке инерциального наведения могут быть определены по зависимостям: [4, стр. 128 - 129]

Для опорной траектории 1 (фиг. 3) при

θ $\genfrac{}{}{0ex}{}{\text{*}}{\text{0}}$ = 40^° в вершине траектории (при θ = 0) получим в соответствии с системой уравнений (2,1 - 2,6) b₁=0,0197;

t_s=24,7 c; Y_s=3540 м; X_s=7300 м.The available projectile overload at any point on the trajectory can be determined by the dependence [4, p. 128):

The trajectory parameters in the inertial guidance section can be determined by the dependencies: [4, p. 128 - 129]

For the reference path 1 (Fig. 3) when

θ $\genfrac{}{}{0ex}{}{\text{*}}{\text{0}}$ = 40 ^° at the top of the trajectory (at θ = 0) we obtain in accordance with the system of equations (2.1 - 2.6) b ₁ = 0.0197;

t _s = 24.7 s; Y _s = 3540 m; X _s = 7300 m.

The available overload in accordance with (3) will be equal to 1.8 units, and the restriction level U _ogr from equation (1.1 - 1.3) will be 1.01 V.

34,5 c;

207,5 м/с;

3135 м;

9330 м. При дальнейшем движении под углом θ = -25^o промах ЛА относительно цели, расположенной на дальности 16500 м, составит "минус" 211 м.In accordance with equation (4.5), to maintain the inclination of the trajectory at an angle θ = -25 ^o, it is necessary to set the pitch program angle equal to ϑ = -22.5 ^o . By the beginning of the inertial guidance section (at θ = -25 ^o ), neglecting the control action during the set of a balanced angle of attack, we obtain the following trajectory parameters

34.5 s;

207.5 m / s;

3135 m;

9330 m. With further movement at an angle θ = -25 ^{o the} aircraft’s miss with respect to the target located at a distance of 16500 m will be “minus” 211 m.

θ₀= 50^o, к началу инерциального наведения (θ_u= -25^o), будет иметь

37,4 с;

171,6 м/с;

4970 м;

8700 м; n_p=0,8 ед.For trajectory 2 (Fig. 3) obtained at

θ ₀ = 50 ^o , to the beginning of inertial guidance (θ _u = -25 ^o ), will have

37.4 s;

171.6 m / s;

4970 m;

8700 m; n _p = 0.8 units

With the known guidance method, if there is no restriction on the size of the control command, the aircraft will retain the angle of inclination of the trajectory, starting from 40 s, equal to θ = -25 ^o . A miss for a target located at a distance of 16,500 m will be 1,205 m (trajectory 2, Fig. 3).

то под действием ускорения силы тяжести угол наклона траектории снаряда будет увеличиваться до тех пор, пока растущая с набором скорости и повышением плотности воздуха развиваемая перегрузка снаряда не приблизится к величине g•cosθ. Это произойдет при угле наклона траектории "минус" 37^o в момент времени t=50,4 с. После этого под действием команды U_огр угол продольной оси снаряда начнет приближаться к программному значению, а угол наклона траектории к "минус" 25^o.In the proposed guidance method, due to the fact that the overload developed by the projectile will be limited to the beginning of the inertial guidance section with the value:

then under the influence of gravity acceleration, the angle of inclination of the projectile trajectory will increase until the developed projectile overload, which grows with a set of speeds and an increase in air density, approaches g • cosθ. This will happen when the angle of inclination of the trajectory is "minus" 37 ^o at time t = 50.4 s. After that, under the action of the command U _{ogre, the} angle of the longitudinal axis of the projectile will begin to approach the program value, and the angle of inclination of the trajectory to "minus" 25 ^o .

угол тангажа снаряда ϑ = -22,5^o, угол наклона траектории θ = -29^o. С этого момента угол тангажа снаряда и соответственно угол наклона траектории будут находиться вблизи программных значений. Промах относительно цели в предлагаемом способе составит 275 м вместо 1205 м по известному способу (кривая 3, фиг. 3).At time t = 69 s, the projectile will be at a point with coordinates x = 14300 m, y = 1330 m. In this case, the available overload will be equal to n _p = 2.85 units, speed

the pitch angle of the projectile ϑ = -22.5 ^o , the angle of inclination of the trajectory θ = -29 ^o . From this moment, the pitch angle of the projectile and, accordingly, the angle of inclination of the trajectory will be close to the program values. The miss relative to the target in the proposed method is 275 m instead of 1205 m according to the known method (curve 3, Fig. 3).

Technically, the given control method is implemented on an amplifier-limiter with a double amplitude limitation [3]. Limit levels are set immediately before the launch of the projectile so that the amplitude of the first limit is A = U _ogre , and the amplitude of the second limit is B = 0 (Fig. 4).

Then, if the measured angle of inclination of the longitudinal axis of the projectile ϑ will be greater than ϑ _pr , then the control signal at the RP U _{in will be} reset. In the case when ϑ <ϑ _pr , and the required value of U _in does not exceed U _ogre , control is carried out according to the linear law U _in = K • (ϑ _pr -ϑ). When ϑ reaches a certain value of ϑ _ogre , at which K • (ϑ _pr -ϑ)> U _ogre , the control signal is limited at the level of U _ogre .

 The value of the estimated transverse overload of the projectile at the highest point of the reference trajectory can be determined in advance and entered in the shooting tables, or in the case of an automated calculation of the settings, the calculator must implement an algorithm for calculating the available overload.

Approximately the value of the available overload can be determined by the dependence (3) at Y = Y _s ', V = V _s '.

The values of the height of the trajectory Y _s and the velocity at the top of the trajectory V _{s are} determined from the dependences (2.1 - 2.5), setting θ = 0 in them. At the beginning, the approximate values of Y _s and V _{s are set} and b _{1 is} calculated from the dependence (2.5), then dependencies (2.2) and (2.4) determine the adjusted values of b ₁ . As soon as two sequentially calculated values of Y _s 'and V _s ' will differ from each other no more than value assignments (for example 1%), then n _{pr is} calculated according to dependence (3).

 Literature.

 1. The journal "Missile Technology and Cosmonautics", t. 16, N 2, February 1978, "Mir", Moscow, pp. 184 - 194.

 2. The journal "Rocket Technology and Cosmonautics", vol. 18, N 2, February 1980, Mir, Moscow, pp. 128 - 138.

 3. Okunev B.N. The main task of external ballistics and analytical methods for its solution. ONTI, GTTI, Leningrad, Moscow 1934

 4. Dmitrievsky A.A. External ballistics - M.: Mechanical Engineering, 1972, p. 584.

 5. Tetelbaum I. M., Schneider Yu.R. The practice of analog modeling of dynamic systems: a Reference manual - M .: Energoatomizdat, 1987, p. 149.

Claims (1)

A method of combined control of an aircraft (LA), including launching an aircraft from a launch device at an angle Θ ₀ to the horizon with an initial speed V ₀ , pointing the aircraft along a ballistic trajectory, inertial pointing of the aircraft with a programmed angle of the longitudinal axis of the aircraft to the horizon, homing to a direct hit Aircraft in the target, characterized in that the value of the control signals U _I at the input of the steering gear of the aircraft in the inertial guidance section is calculated in accordance with the dependencies

where U _ogre - the value of the limitation of the input signal level;

δ _m - the maximum angle of deviation of the rudders;
ϑ _pr - program angle of the longitudinal axis of the aircraft;
ϑ - current angle of inclination of the longitudinal axis of the aircraft;
K is the gear ratio of the pitch angle sensor;
K ₁ - gear ratio of the steering gear;
n _p is the calculated transverse overload of the aircraft at the highest point of the ballistic flight when launched at an angle Θ ₀ -3σ _θ0 with the initial speed

σ _θ0 is the standard deviation of the launch angle of the aircraft;

standard deviation of the launch speed of the aircraft.

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