UA82858C2 - Method for device for control of directed rocket by means of drive that follows orientation of trajectory - Google Patents

Abstract

The invention concerns a method for piloting a craft steerable by means of an actuator, servocontrolled in attitude on a trajectory and subjected to external perturbations which consists, at least in pitch, in first defining a piloting law including a first order correcting term, a transfer term comprising parameters characteristic of the craft and of the external perturbations (A6, K1), a gain in attitude (Kp) and a gain in attitude speed (Kv); and relationships correlating the attitude speed gain and the parameters of the correcting term based in particular on attitude gain and on terms representing rapidity ( ) and damping ( ) instructions, and for at least one phase of the craft flight, in selecting the attitude gain and the rapidity and damping terms and deducing therefrom the parameters of the correcting term and the speed gain of the piloting law.

Description

The present invention relates to a method and device for controlling a propelled rocket equipped with orientation control means, in particular, a launch vehicle with oriented nozzles.

When controlling the flight of a launch vehicle, a distinction is made between guidance and control. In general, it can be said that - the guidance system issues commands that allow you to perform the task (given orientation values). These commands allow you to determine the movement of the center of gravity of the launch vehicle so that it reaches a predetermined goal (usually an intermediate orbit), - the control system executes the commands coming from the guidance system, stabilizing the movement of the launch vehicle around its center of gravity and taking into account changes in in the external environment (disturbances such as wind, gusts, etc.) and in the internal environment (dispersion and errors in the parameters of the launch vehicle, etc.).

The two previous provisions can be clarified as follows.

Based on the known location of the target and the current position of the launch vehicle, which is measured by inertial sensors, the guidance system calculates a trajectory to which the marking of the given value of the orientation of the launch vehicle must correspond (its origin is set at the center of gravity of the launch vehicle, and one of its axes is the direction with which the axis of the rocket should coincide; inertial measurements of the launch vehicle determine the orientation of the launch vehicle, determined relative to a fixed reference point on the ground in a coordinate system considered to be Galilean).

When no perturbations act on the launch vehicle, the control loop responds instantly, the axis of the rocket coincides with one of the axes of the given coordinate system.

The control system is an automatic adjustment circuit, the first task of which is to stabilize the movements of the launch vehicle around its center of gravity (the concept of stability), as a rule, by adjusting the direction of the jet thrust. On the other hand, the control system, as far as possible, realizes the given value of the orientation coming from the guidance system (the control system is an automatic "executive" system) and provides a certain immunity of the launch vehicle in relation to disturbances acting on it (the control system is a control circuit ). For this purpose, the functional control unit determines the most acceptable direction of jet thrust on the basis of inertial orientation measurements (the same ones used by the guidance system).

A disturbance such as wind is an obstacle to the launch vehicle's flight, creating motion around its center of gravity.

In fact, the wind also creates the overall motion of the launch vehicle (progressive motion or drift). Since the guidance system takes into account the current position of the launch vehicle (using inertial measurements) to produce a new trajectory up to the end point, it would not be a mistake to consider only the movement around the center of gravity to produce data intended for the effective operation of the functional control unit.

The control function must take into account some features, in particular, the fact that the launch vehicle is not a stationary system in its essence. its physical characteristics (position of the center of gravity, mass, inertia) and aerodynamic characteristics change depending on the flight. Transient phases are interrelated with the control function, including such as stage separation, the presence of external disturbances (wind) and internal disturbances (for example, displacement of the jet thrust direction). Finally, as in any physical system, the launch vehicle parameters are affected by errors.

The control function is set based on the flight task strictly taking into account the above elements. Naturally, this flight task includes: - stabilization of the launch vehicle; - alignment of the launch vehicle and minimization of the angle of attack a in the flight phase in the layers of the atmosphere (Fig. 1); - following the given orientation value produced by the guidance system; - elimination of the consequences of external and internal disturbances; - taking into account the characteristics of the power steering.

As a rule, the control algorithm is entered in the form of recurrent equations (repeatedly) into a digital computer. Based on the information coming from the sensors (orientation, orientation rate, acceleration) and navigation devices, the guidance algorithm calculates the given orientation values that are implemented by the control loop and which are thus used in the control algorithm.

Usually, the control rules are formed on the basis of the assumption that the pitch (horizontal axis), yaw (vertical axis) and roll (longitudinal axis) axes can be separated. This assumption allows synthesizing (formulating) the rule using a uniaxial model valid for small angles. The effect of combining three rules (one for each axis) is subsequently tested by triaxial simulation.

Figure 2 shows a uniaxial functional control scheme. The control algorithm is considered as a "black box" that contains the information that comes from the sensors of the launch vehicle (an inertial unit for measuring the orientation and a gyrometer for measuring the velocity of the orientation) and determines the command for the uniaxial drive, in this case the nozzle.

The uniaxial scheme assumes that the motion of the launch vehicle is motion in a plane. On the other hand, the drive position is adjusted automatically. It assumes a position that corresponds to the given nozzle adjustment value expressed in elongation.

Prior art

The main difficulty faced by the control system designer is to use the physical conditions that the control rule must satisfy to define a mathematical criterion that can then be optimized. For such use, it is necessary to depart from the physical essence of the conditions in order to find an adequate mathematical expression.

For example, robustness (from a physical point of view) is expressed as the damping of the time response of a control loop that is perturbed or commanded. From a mathematical point of view, it can also be said that mirage stability (damping quality) is related to the overvoltage factor of the control circuit (resonance of the control circuit at a given frequency). A clear connection between attenuation and overvoltage coefficient exists only in so-called academic systems (second-order systems).

Despite these limitations, modern control rule designs use the concept of overvoltage rather than the concept of damping to take into account the problem of stability, since this concept of overvoltage most closely matches the criteria developed based on the minimization of the mathematical norm.

Since the relationship between overvoltage and damping is not direct, damping is only determined later by time simulation.

The following synthesis algorithms are used in known control rules: - OS method (A4, M45, AB); - the Noo method (AB).

Each algorithm uses synthesis parameters that allow optimizing the rule so that it meets the requirements and conditions necessary for control (stability, alignment efficiency, reliability of stability and operation).

The specified methods of synthesis are intended for the generalization of various problems using the concept of a standard system. Most problems of analysis as well as control synthesis can be written in the form of the same standard scheme, with the help of which you can formulate the conditions in the form of a mathematical criterion that tries to minimize (Neo or Nos norm). Thus, the concept of a standard problem is of great importance for the generalization of problems in the form of a single mathematical criterion. However, such a generalization occurs at the expense of a departure from the physical value that the resulting criterion will have.

On the other hand, reduction to the standard form, as a rule, leads to the complication of the algorithm.

Since the parameters of the standard are not necessarily suitable for solving the given problem, it is necessary to allow additional degrees of freedom (where the order increases) for the correct retranscribing of the given tasks.

The latest methods used at this time make an undeniable contribution to the analysis and synthesis of management problems. On the other hand, the criterion obtained as a result, on the basis of which the synthesis of management tasks is carried out, leads to the use of the "trial and error" method to find a satisfactory result.

The solutions obtained by these algorithms do not allow the relationship between the frequency behavior and the time behavior to be achieved other than by modeling, in particular, a high level of closed-loop overvoltage correlated with satisfactory damping. Thus, the optimal rule can be obtained only through successive iterations of "synthesis-simulation-verification".

The technical task of this invention is to create a method and device for a guided missile using a tracking drive in relation to a trajectory subject to the influence of external disturbances, in which a control algorithm is used, the parameters of which can be correlated with the physical conditions set for control.

The task is solved by "7 creating a method of controlling a guided missile using a drive that monitors the orientation of the UU/O trajectory, which is subject to the influence of external disturbances, the specified method consists in the fact that at least for the pitch (horizontal axis) a control rule is determined in advance, which contains the correction term of the first order, the transition term containing the characteristic parameters of the missile and external disturbances, the orientation gain and the orientation speed gain, determine in advance the correlations between the orientation speed gain and the parameters of the correction term depending on the orientation gain, parameters of the transition member of the rocket and two members representing the given values of speed and damping of the control loop, for at least one flight phase, determine the values of the parameters of the transition member of the rocket and choose the orientation gain factor and the speed and damping members, from which the pairs are derived eters of the correction term and the speed gain, and the drive uses a control rule with these parameter values.

According to the invention, there is also proposed a control device designed to implement the method described above, comprising an input for receiving at least one pitch orientation measurement, an output intended for transmitting a command to the drive, at least one automatic pitch control loop connected between the input and the output and holding correction filter of the first order, while the transition term contains the characteristic parameters of the missile and external disturbances, the orientation gain and the orientation speed gain.

It can be noted that such a simple choice of a control rule goes against the modern analysis used by specialists in this field of technology, according to which a simple algorithm, that is, one containing a small number of parameters, is not suitable for normal control. However, contrary to expectations, such a simple choice of control rule will allow the formulation of rules that allow, with some simplification, for example, by limited closed-loop deployment, to correlate the parameters of the rule with physical parameters such as damping (i.e., stability) and auto-adjustment speed (i.e. , efficiency in time), by the orientation amplification factor around the axis under consideration, as well as with the physical parameters of the simulation.

Brief description of the drawings

In the future, the distinctive features and advantages of this invention are explained by the description of the preferred version of the embodiment with references to the accompanying drawings, in which: Fig. 1 depicts a classic scheme of influences on a rocket, which is subject to dynamic pressure; Fig. 2 - a well-known diagram of a single-axis control system; Fig. 3 is a well-known scheme that defines different angles for these axes; Fig. 4 is a diagram of the architecture of the analytical control rule used for this axis according to the invention; Fig. 5 is a variant of the architecture scheme for the case when the angular velocity around the axis under consideration is not measured, according to the invention; fig. b - a curve correlating the amplification factor and the phase of the open circuit in the nominal case and in the so-called cases NU and VU, according to the invention; Fig. 7 and 8 are curves showing the missile's response to orientation, adjustment of the nozzle and the angle of attack of the closed control loop for NL (Ablp, KI tah) and HF (Abtah, KI tip), according to the invention.

A detailed description of preferred embodiments of the invention

According to this invention, the necessary (in terms of complexity) control algorithm is used to ensure control of the missile.

The algorithm is based on a generic structure, the parameters of which are calculated strictly depending on the parameters of the corresponding missile and the necessary control characteristics (trajectory tracking, minimization of the angle of attack, intermediate phases, etc.). The rule clearly takes into account the errors of the missile parameters and the delay of the globalized chain (the delay caused by the dynamics of the drive and the delay of the sensors and the delay due to discretization).

This rule is applicable for non-stationary systems (change in launch vehicle parameters). The non-stationary rule is obtained quickly because its parameters are formal. As a result of the generic structure, there are no problems associated with interpolation techniques that are widely used in launch vehicle control systems.

An indisputable advantage of this method is that the synthesis parameters are physical parameters of a higher level (closed-loop attenuation, passband). The reduction to these parameters of the corresponding physical quantity and the corresponding problem is known. Thus, regulation is carried out directly on the basis of physical conditions established by the developer.

This follows from the following explanations, which, for the sake of simplicity, are limited to the control of the hard flight mode of the rocket. The possible presence of flexible modes can be taken into account when choosing a "cautious policy", i.e., avoiding the excitation of flexible modes by loop gain of the control rule (for example, by adding an attenuator filter).

Developing a control rule requires the use of a model of the launch vehicle, called a synthesis model. The uncorrected dynamics of the open circuit can be represented by two combined equations describing the movements during orientation and wear of the launch vehicle (Fig. 3).

The development of a control rule requires the selection of an appropriate model of the launch vehicle, called the synthesis model.

The uncorrected dynamics of the open circuit can be represented, with reference to the values defined in Fig. 3, in the form of two combined equations describing the movement of the orientation and wear of the missile.

First of all, if we limit ourselves to small angles, the pitch moment can be determined by equations (1) given in the Appendix, while the forces acting transversely to the airspeed axis can be determined using equations (2).

In addition, different angles (Fig. 3) are related by equation (3), which, after differentiation by time, gives equation (4).

After eliminating y in equations (2) and (4), the system of equations (5) is obtained, where 5 is the Laplace variable.

If we assume that the terms identified in condition (C) are small for values contained in the passband of the control loop, then expression (5) can be arrived at through expression (6).

This expression shows that the system is a second-order system with first-order dynamics, which is unstable when pulsating Av .

The system of equations (6) is a synthesis model, on the basis of which the foundations of analytical synthesis will be developed.

It follows that the transition of the command, that is, the transition between the controlled variable (6) and the command (adjustment of the nozzle dD) is expressed by the term Ki/(52-Av).

The behavior of the servo and the measuring chain can be modeled using the general net delay of the chain in series with the transition of the command, from which expression (7) is obtained, where C" from the point of view of automation is the mathematical expression of the physical system that must be embodied in the circuit.

The net delay in the mentioned expression is obtained by approximating the RAOE of the first order.

Inclusion in the circuit is carried out in such a way that the corresponding closed circuit meets the requirements of stability and efficiency.

Examples of the generic architecture of an analytical control rule

Fig. 4 shows the generic architecture of the analytical control rule. It is presented with the assumption that orientation measurements as well as orientation velocities are available, but as will be shown below, this assumption is not a limitation of the method.

In this architecture, the delay t combines the circuit delay, the delay due to sampling - blocking, as well as the delay equivalent to the phase lag of the steering wheel. This delay may also include a filter phase delay designed to locally attenuate the loop gain of the control rule to ensure gain reliability in the flexible launch vehicle mode.

In the direct circuit 5--o01/5-952, the correction term of the first order and the coefficient of amplification of the speed of orientation (Ku) and the coefficient of amplification of orientation (Kr) are introduced. In this case, the order of the control rule is 1.

We evaluate the first dynamics of the method in a closed circuit for its fixation on the model of the second order of parameters (C, op). This technique allows you to carefully condense the efficiency-stability trade-off.

Damping provides a measure of stability, while op (speed auto-adjustment) generates efficiency over time. Thus, the parameters ((C, op) represent "higher level" adjustment parameters.

After that, a limited expansion of the closed loop to the third order is carried out. A system of three equations is obtained, on the basis of which three of the four control parameters of the control rule (К, Фи, ог) are derived depending on (Кр, С, оп) using equations (8).

In order not to clutter the presentation, the previous solutions were formulated without taking into account the general delay of the contour t. When it is present, these solutions become much more complicated. However, solutions are always available and can be easily integrated into a rule synthesis computer program. At the same time, the following observations can be formulated: the pole of the corrective member of the direct chain is unstable (о2-«0);

parameters (K, Phi, yug) are a function of Kr. From the point of view of mathematics, it would be possible to determine K and two other parameters depending on the fourth. However, it will be more correct to leave Kr as a free parameter, since it can be chosen independently of the dynamics of the closed loop. A zero value of Kr helps to minimize the angle of attack of the missile, while a non-zero value will lead to tracking and adjusting the orientation of the launch vehicle; in Fi UAb, ul of the correction member compensates for the stable dynamics of the rocket; at Авб-0, the pole and zero of the correction term of the direct circuit tend to 0. The stability of the system is ensured by the phase advance, which is well known in classical control methods and which is defined as

Kr-8Ku.

Defining the rule parameters shows that they were calculated based on higher-level parameters (damping and closed-loop speed). This makes it possible to establish a relationship between attenuation and the stability margin of automatic regulation, expressed in the form of the gain limit Mo (recall that the gain limit is the value of the gain coefficient of the open circuit when its input and output are phase-shifted by 180", while the correct instability is obtained at the value of the amplification limit - 1), which is obtained using equation (9).

Note that the proof of the gain limit is carried out by (, (attenuation of the closed loop). In particular, it is minimal at zero orientation gain coefficient (Ke-0). Its value is obtained from equation (10).

If the dynamics of a closed circuit is chosen, which is defined as Фп-Ав (in this case, the launch vehicle is stabilized), then the gain limit is determined by expression (11).

This expression establishes a direct relationship between the damping and the stability margin.

The above-mentioned method can be applied even if the orientation velocity measurement is not available. In this case, it is evaluated using a clean (the presence of a direct transmission term at high frequencies) or a strictly clean (limitation of the filter gain at high frequencies with a steepness of at least 20dB per decade) high-pass filter. Let's choose a high-pass filter with a limit of 20dB per decade. Its expression is determined by equation (12).

In this case, the control rule becomes a single-circuit rule, the architecture of which is shown in Fig.5.

As in the case of the feedback signal of the orientation velocity, the expressions of the control rule (ой, ог, Ку) can be completely determined by expanding the closed loop to the third order and depending on the parameters Кр, оп, С, Ос, etc.

Selection of adjustment parameters

Orientation tracking necessarily leads to the measurement of the angle of attack, which can be difficult from the point of view of the joint forces acting on the launch vehicle. Therefore, the selection of the orientation gain is carried out as follows.

It is necessary to set a non-zero value of Kr if it is necessary to ensure good tracking of the orientation (start, flight phase outside the atmosphere).

A low or zero value of K» must be chosen to minimize the angle of attack in the critical phases of flight corresponding to the maximum dynamic pressure.

The value of Соп determines the speed of response of the response of the control circuit. CO pulsation should be selected in such a way that the alignment of the launch vehicle at maximum dynamic pressure is compatible with the allowable angle of attack and the characteristics of the servo. Thus, the control parameter sop is closely related to the problem of the efficiency of the control rule. When choosing (Op), it is necessary to take into account errors in the parameters of the launch vehicle (the problem of reliability with efficiency).

The stability and reliability of stability are set by parameter C, so that all launch vehicles, those characterized by variations in Av and Ki parameters, are stabilized at attenuation, at least equal to the attenuation specified by the developer.

Non-stationary development of the rule

Considering the various conditions that arise during the flight (launch, trajectory tracking, rocket stability in critical phases, launch accuracy (civilian launch vehicle) or firing accuracy (military missile)), the most acceptable rule results from the interpolation of several linear rules, which are calculated at different moments of the flight. In this case, the rule parameters become functions of time. Since the structure of the correction term is generic, these parameters can be interpolated directly, while they retain the same physical essence from one moment to another.

At the moment of its flight, the following determine: - parameters of the launch vehicle Av (Y) K) (0; - orientation gain coefficient Kr (0; - speed of the control circuit op); damping C (X), which can be chosen constant for the entire flight.

The parameters of the rule, corresponding to Her moment, are estimated on the basis of the formulas obtained in the limited expansion of the closed circuit.

The synthesis process is repeated for n flight points of the stage under consideration. The general non-stationary rule is obtained by interpolation of n stationary rules.

Extending the structure of the rule

The analytical rule has two generic basic forms depending on the presence or absence of the feedback signal of the orientation speed, which are obtained using equations (13) and (14).

These forms can be expanded, although the essence of the decision is determined by one of the two basic forms. The first proposal consists in complicating one or another basic form by multiplying them using a filter of any order. This filter is designed either to locally limit the frequency gain to avoid exciting the flexible modes of the launch vehicle at certain frequencies, or to attenuate the loop gain starting at some frequency determined by the developer.

It goes without saying that another type of filter can be used, which does not perform any special function other than the apparent complication of one of the two basic forms.

Physical essence and typical parameter values

Taking into account the analytical expression of the parameters Ku, о, and ебог, some of them can be given a given physical value.

Pulsation oi is equal to ud due to the dynamics of the launch vehicle itself. They make her compensation. In particular, it can be written that the pulsation so is proportional to the dynamic pressure at the moment of flight under consideration and inversely proportional to the inertia of the launch vehicle. its value is initially difficult to quantify, as it depends on the characteristics of the launch vehicle.

The ripple expression og is a function of the speed chosen for the control loop. Just as in the previous case, it can be specified that the greater the value of sog, the faster the control loop will work.

In practice, the speed of the contour is chosen in such a way that it does not force the vehicle (launcher) to go beyond its own UAB dynamics.

The expression of the amplification factor Ku is relatively complicated. In addition to members Ki, Av, Kr and op appears. damping term C, which is a parameter that establishes the stability and reliability of the stability of the control loop. From the point of view of quality, the degree of stability of automatic regulation is higher, the greater the numerical value of Ku, , which is proportional to b.

Example

The example below shows the application of the analytical synthesis of the steady-state control rule for the launch vehicle X. The values of the Av and Ki parameters at the maximum dynamic pressure (corresponding to the fixed trajectory) are obtained from:

Av-Z35a15c2

Ki - 30310 s?

Majorant for the stability of the launch vehicle corresponds to:

Av-Avtakh-50 s?

Ki-Kitip-20 s?

The control rule is synthesized at the maximum dynamic pressure with the aforementioned parameters of the launch vehicle. The damping of the dynamics of the closed circuit is set by:

СхО,З (this is the necessary attenuation corresponding to the majorant)

In order to minimize the angle of attack at the maximum dynamic pressure, a zero position amplification factor is set, i.e.:

Cre-0

The choice of the op parameter (which sets the speed of automatic adjustment) is determined by checking that its value does not exceed the speed and deviation of the turn that the steering wheel can achieve.

It is assumed that the value of op is known and is equal to:

Fp-Z, 5 rad/s

If this does not match the assumption, the correction term synthesis step is repeated until a value (FO) is found that is compatible with the rudder characteristics verified by time simulation.

The parameters of the correction term are obtained using the equations underlying the invention. We have (for Kr - 0) oi idbyaah ho i 7 reve th na dos ku te UZH tab) Od E YUA V v a

Kosha and me

The application of the formulas assumes that the control chain does not introduce a delay. If this assumption is not verified, then it is necessary to re-evaluate the analytical expressions of the parameters of the correction term, taking into account the presence of chain delay.

The gain limit at low frequency is obtained from the analytical expression

Massive iv, hek V ххх о Zm i.e. 2 dB at. What 35

Timing simulations show that a maximum nozzle rotation of 5.6" and a maximum nozzle rotation speed of 30" are obtained. These figures are obtained for the majorant from the point of view of stability (the case of Avtakh,

Kitip, which is called the NU case).

The frequency characteristics are shown in Fig.b. In particular, with the help of modeling, the amplification limit "calculated in the case of the majorant (Avtakh, Kitip) is found again.

The time characteristics obtained by calculated regulation are shown in the form of curves in Fig. 7 and 8.

The analytical control rule allows you to quickly perform stability and efficiency problems for a set of launch vehicles described by variations of Av and Ki parameters. In contrast to the methods of the Os or No type, it is possible to abandon the frequency iteration, which is in the further verification of stability reserves. Time iteration, which allows you to choose the speed of the control loop, is still necessary, since the problem of efficiency is directly related to the nature of disturbances acting on the launch vehicle (wind, gusts, internal disturbances, etc.). The non-stationary expansion is straightforward and poses no particular problems from the point of view of synthesis.

If necessary, the analytical control rule can be extended to other vehicles besides launch vehicles.

When considered in principle, it can be noted that it is produced on the basis of a synthesis model derived from the equations of flight mechanics, while the most significant assumption is the instability of the aircraft. Therefore, there is no limit to the fact that this type of rule can be applied to aircraft that are not stable on one or more axes of flight (pitch, yaw and roll). To do this, it is enough to apply the flight mechanics equations for unstable aircraft. This will make it possible to obtain a new synthesis model from which the parameters of the analytical control rule can be derived.

Mainly: - for the flight phase, the values (Letakh, Kitip) of the parameters of the transition term, majorizing external disturbances, are determined; - a simulation acceptance test is used to select the orientation gain and velocity and damping terms for the entire flight time and these selected parameters are used throughout the flight if they have passed the reliability check, and if not, then the flight time is divided into several phases; - the transition term is determined for external disturbances, mainly arising from the action of atmospheric air; - on the longitudinal axis and on the vertical axis, the control rules are applied, determined by the same method as the control rule on the horizontal axis.

Appendix: Formulas be shva m -k and what g

Event and ; tour" Ob va ReviKY vazu nya aud 2 tu tu y

Where t: the mass of the launch vehicle

A: dynamic pressure

Sga: gradient of lifting force

I: inertia along the horizontal axis 5: the reference surface of the apparatus for the lifting force

Re: jet thrust of the device

And there is a distance (center of gravity - aerodynamic focus)

And distance (center of gravity - center of rotation of the nozzle)

Vera kasa svi a 3

Anne M

I) beufat days for ME otit de u ushu (wind gradient) from U

Where, (The speed of M2 is the projection of the speed of the launch vehicle on the vertical axis)

Be ibbsSe Kt

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MK, bo, giy p 6 - byu vysh gi (b l/ same va meti) dok r IZHU

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Wow pitches: - 17 (rule with feedback signals of speed 6 and orientation 6) (13) d-e- (ЖЕК сх) (5 /no" kg/more (know ! (rule without feedback signal of speed 8) (14)

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FiOZ

Claims (11)

1. A method of controlling a guided missile using an actuator that monitors the orientation of a trajectory sensitive to the influence of external disturbances, which consists in predetermining a control rule containing a correction term of the first order, a transition term, at least for the pitch (for the horizontal axis) , containing the characteristic parameters of the missile and external disturbances (Ab, Ki), the coefficient (Kr) of the orientation enhancement and the coefficient (К,) of the orientation speed enhancement, predetermine the correlation relationships between the orientation speed enhancement coefficient and the parameters of the correction term depending on the coefficient amplification of the orientation, parameters of the transition member of the rocket and two members representing the given values of velocity (n) 1 attenuation (5 ), at least for one phase of the rocket flight, determine the values of the parameters of the transition member and choose the coefficient of amplification of the orientation and members of the velocity 1 attenuation, according to which derive the parameters of the correction term and the amplification factor speed, and a control rule with these parameter values is applied to the drive. 2. The method according to claim 1, which differs in that the correlation rules are determined on the basis of a limited expansion of the equation of the closed circuit formed by the rocket and its drive. 3. The method according to any of paragraphs And or 2, which differs in that for the flight phase, the values (Avetakh, Kitip) of the parameters of the transition term, which is magnified by external disturbances, are determined. 4. The method according to any of claims 1-3, which is characterized by the fact that the control rule additionally provides for the use of a filter. 5. The method according to any one of claims 1-4, which is characterized by the fact that the control rule is applied using orientation and orientation speed measurements. b. The method according to any one of claims 1-4, which is characterized in that the control rule is applied using an evaluation filter designed to estimate the speed of orientation in the control rule. 7. The method according to any one of claims 1-6, which is characterized in that a simulation validation test is applied to select the orientation gain and velocity and damping terms throughout the flight time and these selected parameters are used throughout the flight if the selected parameters have passed validation , and if not, then the flight time is divided into several phases. 8. The method according to any of claims 1-7, which differs in that the transition term is determined for external disturbances, mainly arising under the influence of atmospheric air. 9. The method according to any one of claims 1-8, characterized in that the control rule additionally includes a delay term. 10. A method according to any one of claims 1-9, characterized in that yaw (longitudinal axis) 1 roll (vertical axis) control rules determined according to the same method as 1 pitch control rule (horizontal axis ). 11. A control device designed to implement the method according to any of claims 1-10, containing an input for receiving at least one pitch orientation measurement (on the horizontal axis), an output for transmitting a command to the drive, at least one circuit for automatic pitch adjustment, which connected between input and output 1 contains a correction filter of the first order, while the transition term contains the characteristic parameters of the missile and external disturbances, the orientation gain 1 the orientation speed gain.