SI8212278A8 - Process for controlling aerodynamical body - Google Patents

Description

The present invention relates to a method for controlling an aerodynamic body, such as a projectile or grain, upon firing a target. Using the output signal of the guidance device, which is a measure of the instantaneous value of the angular error between the fixed axis of the body, λ

primarily the axis of symmetry of the body, and the line of sight from the body to the target, the body is guided by the flight path to the target, responding to a variable control signal that depends on the speed of the angular change of the line of sight.

2. Technical problem:

The object of the invention is to propose a process of the aforementioned type for managing a projectile without the need for a gyro device.

5 »State of the art:

Missiles previously had an guidance device that used the gyro device to determine the angular velocity change required to calculate the angular velocity change <o line of sight by expression (o - S + to determine the angular error 6 »between the position of the projectile and the line of sight to the target. ·

In order to reduce costs, it is advisable to eliminate expensive giro.

R. Description of solution to a technical problem:

According to the invention, the object of the invention is achieved by, on the basis of the relations observing the aerodynamic behavior of the body with respect to the target, the value of the signal representing the speed of the angular change of the line of sight, on the one hand, and the value of the signal representing the speed and

angular changes in body position). on the other side.The two values of signal 3e combine to form the value of the angular error signal. the size of the relationships mentioned ·

The invention is further described in more detail with reference to the accompanying drawings.

Figure 1 is a single plane view of a guided missile, by means of proportional navigation, against a moving target for interception, and some essential sizes are shown.

Figure 2 is a single-channel block diagram of an earlier system for proportional projectile navigation and shows the mode.

Figure 3 shows a single-channel block diagram of the invention, showing its mode of operation, and having a similar arrangement as Figure 2.

The invention can be referred to all types of missiles, for example, guided or artillery projectiles, equipped with guided-turn tools. Figure 1 shows such a projectile M flying on a path to a target-vehicle T moving along a Pg path. The S - ^ - 5 ^ visibility in the four positions I, II, III, and IV shows that the projectile does not approach the target at the same time that the line of sight gradually becomes more parallel as the projectile approaches the target.

At position I, the projectile has a velocity V in the flight direction, <3 being the angle of line of sight between the line of sight S and the inertial reference cabinet R, θ defines the extinguishing position of the projectile between its axis (related to its body) A, which is here the axis of symmetry of the projectile, and the inertial reference direction R. C is the angle of error between the fixed axis A of the projectile, and the line of sight S, It is seen that the angle of error £ _ is obtained from the angle of the line see jivo3'ti <o and angular position 0, by the expression :

ζ. s: G * - θ '

3.

Figure 2 is an operating block diagram of an example of an earlier missile system, type aa proportional navigation, which uses a 1 * guidance device. angles €> line of sight and angular position O, age of block 3 *, result in aU / aO. angular gr. .This angle is measured with a guidance device l ^r whose output signal is a measure of £ the instantaneous angular error between the fixed axis of symmetry A and the line of sight S.

As mentioned in the introduction, this kind of system requires the gyro 2 'used here to determine the angular position of the projectile. MEASURING © _ra i £ sabiraju is to get variable _m ugla line visibility, a hundred business differentiation gives veličinu & _m Ugaona speed changes of the last line vidijivosti.Pomocu's greatness in block 4 * calculated signal koji represents kontrolnu (upravljačku) promenljivu u based on the law of control uoc · <y in accordance with the principle of proportional navigation, where '· c is a constant. The signal representing the control variable in water is not shown to the projectile control unit in block 3 *, and the control variable can be obtained by damaging the control surface.

In the above description of the earlier mode, still in the description of the invention, for the sake of simplicity, it is presented that orojectile moves on the same vertical or horizontal plane, which opens the channel for inclination or for turns. and in practice a missile is also a man? that it also operates on the second plane that is governed by the first first-order (aerodynamic) relationship of the projectile's aerodynamic behavior, which was further used in the published embodiment of the invention, was described to describe the movement in the vertical plane and could therefore be neglected the effect of gravity. It is therefore obvious that the relations describing the movement of the projectile perpendicular to the vertical plane are no longer obsolete.

Figure 3 shows the invention with reference to a solution having proportional navigation. The block diagram of Figure 3 contains blocks 1,3 and 4 having analog functions as well as corresponding blocks in Figure 2 marked with prim tags.

'i.

In order to avoid the need for expensive giro, a computer unit lo, based on relationships describing the aerodynamic behavior of the projectile relative to the target, was used to determine the magnitude of the signal that is predictive or the approximate value of the angular change rate of the line visible. form a more or less approximate mathematical model of the aerodynamic behavior of the missile relative to the target. In the opioid first-order solution, these relations, as can be seen, ou known earlier, however exclude the fact that other similar relationships can be used, within the scope of the invention.

In the first step, the computing unit lo establishes the value of the Q signal, which is an approximation of the angular velocity poz of the position (missile) of the projectile, by calculating the approximate value of e / for aerodynamic attack angle projel / ULa ^ used 'in the second step of the computing unit ·

In the second step, the computing unit lo sets the value "

signal 5, by expression - the ratio of the angular rate of change of a projectile to the line of sight, which is an approximation of the rate of angular change of the line of sight. This signal value is used as an input signal to unit 4 to establish a signal in control variables by the control law, here uec · & in accordance with the principles of proportional navigation.

A signal to control variables that has been previously determined, alternatively a deflection to ^ control surfaces or the like given as a measured signal from a control apparatus in block 3 »serves as an input signal to a computer unit 10.

Two set values of signal G? and (o are combined as shown in unit 20 for determining signal € which is approximately the value of the error angle. At junction point 11 based on ddnos • * · · <£ »S - ute the mentioned values of the two signals result in a signal ** ** which is the approximate value L of the angular error angle change velocity. Later, 'integrational', as shown, in block 16 denoted by the Laplace integration operator, the aforementioned sig / s is found.

Controllable. into which is specified in block 4 through the control law, gives, in the curve of the ambient waves and the aerodynamics of the projectiles, in accordance with aa block 3, the error angle G, which is measured by £ through guidance device 1 in the same way as it was early. It should be noted that the guidance device may be, and may be, initially fixed to the body of the projectile. , because the aim of the inventor is not to have a giro device.

updating the appropriate size is as follows:

Correction or

* z> Θ i A a C b - Γ * Ί Θ £ A and t and, + • s f I II _ A * Θ A o ( A σ A ε A -V H · ¢ -Δ.Θ A Zkk A Δ & A δ £ A A A Δ b * b _t K - -t -t-1 I t

Here, the index t denotes the corrected value in ande vmmeu and indeka t-1 denotes the previous value. The correction factors k ^ - kg are here the coefficients that ou are dependent on the

Z \ £. , aa on the one hand, and from the rotation in the corresponding magnitudes aa on the other hand. Each correction factor k ^ - kg is a pound-mass of the type kf (a ^, a ^ »a-jj» V, r, u), According to the towers, these are variables in the projectile control procedure, it also expresses the forks of the path, as shown in Figure 3 by the presence of block 14. A convenient note for calculating the correlated correction factors ki - kg is via Kalma.n - fil.ter; v, e.g., Introduction to Stonaatic Control Theory, Chapter 5-4, by Karl J. AstrUin, Academic Press, Nev / York, London 1970.

Unit 2o illustrates a successive correction or update. ~. a i. Avelines £. , .. Correction value Δc. The switch shown between the output of said junction point and the output of the integrator 16 illustrates the introduction of a corrected value of size £ ^. Updating other sizes is not shown in detail but 00 is similar.

Λ

The value of the signal £ determined as an approximate value of the error angle is combined by subtracting at the junction point 12 with the value of the signal £ measured in £ le error, resulting in a signifier In difference corresponding to the difference.

This value of the difference signal of the angular error At is used to correct or update the magnitudes, eg, of variables and desired pararaeters, in ratios (calculations) in a computer unit,

The basis for the first coral of a computer unit are two equals:

Vi © -, + a _o c> L + bu © + a ^ cL + h ₂ u where the pronounced states of magnitude © io (. Correspond to the rapid change of the angular position or aorodinarically at the angle; u is controllable (control of the ditch) it is realized as a deflection of the control surface;

^a l » ^a 2 ' ^? '3 ^are a ^ edical parameters which depend on the shape of the projectile and the arrangement of raaaa, b ^ ib ^ are the parameters of coupling and force.

These equations of state ou approximation of the more complete equations of state, found in, e.g., Bynaraic of Atmospheric Flight ”p. 162,163, by Bernard. htkin, John 'Jiley A Sons Inc., 1972,

The solution of the two equations of state results in an approximate one

ZS-Zs with values of © and X. the angular rate of change of call are, respectively, an awkward angle of attack.

With respect to the param ethers b ^ ib ^ in the state-eaters, it is assumed in this embodiment of the invention that b ^. α 0; bg “0 ie. that b ^ ib _{2 are} suatinally constant.

t ·

J

During short intervals, the approximate values of Θ and oi are determined by the calculus, with a signal in - which is the output control variable - of unit 4, or with the signal in _{ffl of the} control turn as input to the computing unit ·

To determine, in the second step of computing unit 10,

For the approximate values of S * the speed of change of the angle of the line of sight, the following equation of state is used, ie: 6 = «-which is known by itself. In this equation, the sizes having the same • * symbols as above have the corresponding meanings given above, Č and G ? represent the angular acceleration or angular velocity of the line of sight; V is the projectile velocity of a projectile assumed to be known, and as an example it can be constant; r is the distance from the projectile to the target.

In determining the value of signal 6, which is an approximation of the angular velocity of the line of sight, it first determines the approximate value a of the acceleration ^ of the projectile transversely to the line A of the line of sight, from the previously calculated approximation <%. aerodynamic attack angle, the said acceleration is approximated transversely to the axis of symmetry by the expression a «- (a ^ c <+ b ₂ ti) Y, />

The signal of approximate value 5 is then determined according to the expression S = (2VG - a) / r.

The projectile control system is activated at a predetermined target distance from the target, detected by the guidance device, to obtain the initial value of r _Q for the distance to the target. r may, for example, be expressed as r = «r _Q - Vt where t is the time from which the initial value Γθ of the distance was detected.

To determine the distance r _Q at which the projectile control system should start operating, according to Figure 3, since its signal path r_ ^ _{n to} the computer unit lo.This signal path is given information that establishes r _Q and may affect-, others magnitudes that may depend on r _Q , in addition, the signal path V _in the computer unit lo is shown to determine the velocity V in the solution described here.

8.

In relation to the track, it should be mentioned that the equation of state for the signal of the approximate value Θ '· for the devices seeking less accuracy at the terminal distance can be earned by the equation b 0; other words the angular velocity of the line of sight is assumed to be constant at the interval between the measurement of the angular error »

Δ *

The values of the signal & and Θ are determined by the computer unit as above, are used to obtain the control variable signal in and to obtain the signal value £ ·

After integration, this last signal value £ is used to obtain the signal difference value Δί. comparing the value of the measured angular error signal as shown in unit 12.

A,

As shown in Fig. 5 »the signal value £, which is predicted, is also brought to the guidance device 1 to ensure that said device searches for the target in the correct angular zone.

The signal value, the difference Δί, is used in the projectile guidance procedure to successfully correct or update magnitudes as calculable states and parameters, into the equation.ma of the computer unit. This is shown in Figure 3 in the feedback unit 13 for how to determine the variables and states previously. , oC and kf, a certain value of the angular error £, as well as the coupling and force parameters b ^ ib ^ in each.> assigned, a specific correction factor k ^ - kg, as shown in block 15. Each output signal from that block 15 presents · a correction that is specific to each size ·

According to a particular invention of the invention, the aerodynamic parameters “ ^a 5 can be kept constant throughout the entire procedure.

As shown in Figure .3, therefore, the required accuracy can be achieved by updating only the parameters b ^ ib ^ together with the sizes β, and £; mentioned sizes in. appropriate time for the future,

The above-discussed invention invention units can be implemented with the help of electronic correspondents who provide very fast computer operations - steps.

The primary and very compact irrapleneration of the invention was achieved by a microprocessor, which, according to the invention, is intended to calculate S. angular errors, as well as the calculation of the correction factors k ^ - kg and correlation quantities, and the microprocessor then serves to feedback the values of Δ £ the difference of the angular error to update the size in question. shown in the figure and illustrated by a digital microprocessor, and projectile units shown in the figure above that cooperate with the microprocessor via signal.

At the beginning of the computational procedure, the variable and parameters are given the initial values determined from the current error angle of the projectile tile, and the previously entered information as r. and V. The calculation in the microprocessor is done. at intervals between moraines, the angular errors to obtain the values and the signal values obtained as a result of the calculation in one computer step are stored as predicted. ian the appropriate sizes to be used successively in calculating the next computational step.

The invention was euistic by calling for a particular solution based on the proportional and common sense. control signal which is dependent on the angular velocity ^{z of the} line of sight, i. u = f (<3), In particular, the code missile has control rockets in place of the control surfaces, uses o rodified proportional navigation, where a guided turn is caused when the control signal exceeds a predetermined value.

Claims (5)

1. Procedure for controlling the aerodynamic body, e.g. guided missile or grenade after firing Ivanj in the flight path to the target for interception, whereby the body emits an output signal (£ m) through the guidance device as a measure of the angular error (£) between the fixed axis of the body and the line (Si ') of visibility from the body to the target, while the body is guided by a variable (u, u) control signal depending on the rate of change of the angle (E) of the line of sight, characterized in that, in the computational unit (10), for processing information about the aerodynamic behavior of the body in relation on the target, which receives as input a control variable (u, u), defines the value of the first signal (Θ), as well as * the rate of change of angle (©) of the line of sight used to obtain the control variable signal (u, u), and the value in t of the second signal, as the rate of change of the angular position (θ) of the body, whereby the value of the third signal (£) is formed, as the approximate value of the angular error (£), from the value of the two signals (, / • S θ-) , and what is the value of the signal (bi) , as the difference between the measure -O _c (') and the approximate value (c) of the angular error (c. ) returns to the computer unit (10). 2. The method of claim 1, wherein, before entering the computing unit (10), the value of the difference signal (xb £) of the angular error is multiplied by a factor (K ^ - Κθ). 3. The method of claim 2, n, a, ie, whereby during correction a correction factor (K ^ - Κθ) is updated to the function of the parameters and variable size of the project. Method according to claim 1, characterized in that the value of the signal () of the rate of change of the angular position changes depending on: e · Ok ,, e + 6X- ₂ c <-h b- | K oL - -e * tbt «·» and where O is the angular position change rate, and -0 is its first derivative in time, d is the aerodynamic attack angle, and its derivation by time, into a control variable, a ^, a, ^, a ^ aerodynamic parameters, ib ₂ coupling and / or force parameters, where the value of the signal (6) of the approximate value of the velocity change angle of the line of sight changes depending on : C2.5- <& h bx ^) · “* ¹ ·» where 6 is the velocity of change of angle of line of sight, and (o its derivation by time, V is the traveling velocity of the projectile, ar its, Emerging from the target, where for lower required accuracy 6 * = o Method according to claim 4, characterized in that when multiplying the difference signal value (A £) by the correction factor (k ^ - kg), before its introduction into the computing unit (10), the factor is updated, whereby the coupling parameters and forces (b ^, b ₂ ), while the aerodynamic parameters are kept constant.