RU183670U1 - Rotating homing missile - Google Patents

Abstract

A rotating homing missile belongs to defense technology and can be used in single-channel missiles launched from the shoulder, as well as from various carriers, including, among other things, volley fire at aerial targets.

A rotating homing missile contains an engine, a wing unit, an instrument compartment, including a rocket housing position sensor and a rocket control electric drive, an optical homing head, including an adjustable gyroscope with correction winding, a bearing winding, a gyro rotor position sensor winding, a reference signal generator, a descent sensor and autopilot with a missile pivot in the initial section.

The proposed rotating homing missile allows launching both from the shoulder and from various carriers, including in the presence of target designation signals from the carrier for each missile, which facilitates aiming and carrier control during launch, reducing the preparation and launch time, the ability to carry out both volley launch of missiles for one target, and simultaneous launch of missiles for two targets from one carrier, with automatic rotation of the rocket (missiles) to the required lead angles for each A specific combination of starting conditions.

Description

A rotating homing missile is a defense technique and can be used in rolls of single-channel controlled rockets that are launched from the shoulder, as well as from various carriers, including volley fire against aerial targets.

Known rotating homing missile (Technical description and instruction manual 9K38 TO. Moscow, "Military Publishing", 1987, p. 22-27) [1] as part of MANPADS "Igla" (9K38). The indicated rotating homing missile contains an engine, a wing block, a steering compartment, an optical homing head (OGS), which includes an initially adjusted corrected gyroscope with direction finding and correction coils, an autopilot with a command generator for launching a missile at the initial flight phase (FCR), and missile descent sensor.

Known rotating homing missile (patent for invention No. 2216707 of the Russian Federation) [2]. Said rotating homing missile comprises an engine, a folding tail stabilizer, a hardware compartment with a tracking electric drive for the rocket controls, a missile and OGS sensor, including an adjustable gyro with direction finding and correction coils initially oriented at a given bearing angle, an autopilot whose functional diagram contains a series-connected strip bar filter, nonlinear correction device, first adder, first nonlinearity zone limiter, phase det a ctor, a notch filter, a second adder and a second nonlinearity zone limiter, the output of which is connected to a servo drive, a PCR is connected to the second input of the first adder, a reference voltage generator is connected to the second input of the phase detector, and a rocket angular velocity sensor is connected to the second input of the second adder, the FCR is made in the form of a first sequential branch connected to a bearing bearing coil, consisting of the first and second controlled keys, a third adder connected to the second input of the first sum a torus, and a second sequential branch, consisting of an amplitude detector, a low-pass filter, a comparator with a given threshold value and a bistable element with a given initial state, while the third key is connected to the output of the first key, the output of which is connected to the second input of the third adder, and control input - with the output of a bistable element, the comparator output is connected to the control input of the second managed key, and a time relay is connected to the control input of the first managed key, the input to torogo connected to the missile gathering tow.

The FKR generates a control signal at the initial portion of the flight, which ensures the automatic rotation of the aforementioned missiles to the required lead and elevation angles, provided that the "capture" of the target and the subsequent launch of the rocket occurs by turning the launch tube on the target so that the target is in sight MANPADS devices, the target axis of which is aligned with the axis of the OGS gyroscope (in this case, if there is a target in the field of view of the MANPADS sighting device, the target will also be in the OGS field of view).

The aforementioned missiles have the following disadvantages:

- when placing missiles on various mobile carriers for launching, it is necessary to deploy the launcher or the carrier on the target, which complicates the design of the launcher, makes it difficult to control the carrier and aim, and also increases the preparation time for launching the rocket;

- volley launch of missiles from one carrier can be carried out only for one purpose, since aiming is carried out only by the launcher;

- the phase detector filter used in the autopilot of the aforementioned missiles emphasizes the useful component of the output signal of the phase detector at the speed of the rocket, but at the same time, changes its phase shift when the speed of the rocket changes, which leads to a phase imbalance of the control signal, resulting in is a deterioration in the accuracy of rockets.

The closest in technical essence to the proposed utility model is a rotating homing missile in the complex 9KZZZ (Operation Manual, Part 1 Technical Description, Part 4 9K333.00.00.000 REZ) [3], selected as a prototype. Said rotating homing missile comprises an engine, a wing unit, an instrument compartment including a rocket housing position sensor and a rocket control electric actuator, an OGS, including an adjustable gyroscope initially oriented at a given bearing angle with a correction winding, a bearing winding, a gyro rotor position sensor coil, and a support shaper signals, descent sensor, autopilot, missile rotation scheme in the initial section (RNU).

Functional diagram of the autopilot of the indicated rotary homing missile (patent for invention No. 2400795) [4] contains a first phase detector, a first filter for extracting the average for the period of the input signal carrier, and a first low-pass filter, for series connection of a second phase detector, and a second middle filter for the period of the carrier of the input signal and the second low-pass filter, while the information inputs of the first and second phase detectors are combined and connected to the correction winding; the first controlled limiter, the first and second adders, the first inputs of which are connected to the first and second outputs of the first controlled limiter, respectively, the second controlled limiter, similar to the first controlled limiter, the first and second modulators and the third adder, as well as the third phase detector connected in series, the third filter the allocation of the average for the period of the carrier of the input signal, the third low-pass filter and the first functional unit connected in series to the fourth phase detector , a fourth filter for extracting the average signal carrier period for the period of the input signal, a fourth low-pass filter and a second function block, while the information inputs of the third and fourth phase detectors are combined and connected to the bearing winding, and the outputs of the first and second functional blocks are connected to the second inputs of the first and second adders respectively.

The rocket adopted as a prototype uses a single-channel control system. To create a control force in any direction of space using one pair of rudders, a forced rotation is set relative to the longitudinal axis, while the rocket hull position sensor generates a signal proportional to the current angle of rotation of the rocket hull around the longitudinal axis, which is used as a reference coordinate system associated with the initial position rockets before launch. The signal from the AC correction winding, carrying information about the angular velocity of the target-rocket line of sight in the coordinate reference system rotating at the gyroscope speed, enters the autopilot, where it is converted to direct current signals displaying information about the angular velocity of the line of sight “rocket” -goal ”in a fixed Cartesian coordinate system associated with a rocket is filtered and after non-linear processing is converted into an input signal to the electric drive of the rocket controls in a rotating the rotation frequency of the rocket, the coordinate reference system, providing control of the rocket in two mutually perpendicular planes, which implements the selected guidance method - proportional approach, in which the control signal is proportional to the absolute angular velocity of rotation of the target-rocket line of sight. The phase and amplitude of the control signal supplied to the electric drive of the rocket controls determine the change in direction and intensity of the rocket in flight.

Due to the use in a single-channel rotating rocket [3], a structurally associated sensor of the position of the rocket body and autopilot [4], a change in the rotation frequency of the rocket adopted as a prototype does not change the amplitude and phase of the control signal.

However, the prototype has the same disadvantages as its analogues:

- when placing the rocket on various mobile carriers for launching, it is necessary to deploy the launcher or the carrier on the target, which complicates the design of the launcher, complicates the management of the carrier and aiming, and also increases the preparation time for launching the rocket.

- volley launch of missiles from one carrier can be carried out only for one purpose.

The objective of the proposed technical solution is to create a rocket that provides the possibility of its launch both from the shoulder and from various carriers, as in the case when the "capture" of the target and the subsequent launch of the rocket occurs by turning the axis of the OGS gyroscope relative to the longitudinal axis of the rocket along the outer axis relative to the missile target designation signal from the carrier so that the target is in the OGS field of view, regardless of the position of the sighting axis of the MANPADS sighting device, and in the case when the "capture" of the target and the subsequent launch of the missile It occurs by turning the launch tube to the target so that the target is in the field of view of the MANPADS sighting device, with the automatic rotation of the rocket at the required lead and elevation angles with minimal dependence on the influence of various kinds of external disturbing influences on the rocket trajectory in the initial flight area.

The technical result of the proposed utility model is to facilitate aiming and control of the carrier during the launch of the rocket, reducing the time to prepare and launch the rocket, the ability to launch a salvo launch for various purposes from one carrier.

The technical result is achieved by the fact that in the proposed rocket, as well as in the closest to it, selected as a prototype, containing an engine, a wing unit, an instrument compartment, including a position sensor for the rocket body and an electric drive for the rocket controls, an optical homing head, including an adjustable gyroscope with correction winding, bearing winding, gyro rotor position sensor winding, reference signal shaper, descent sensor and autopilot with missile rotation scheme at the initial section (RNU), fun the national circuit of which contains a first phase detector connected in series, a first filter for extracting the average for the period of the input signal carrier and a first low-pass filter, a second phase detector in series, a second filter for extracting the average for the period of the input signal carrier and the second low-pass filter, while the inputs of the first and second phase detectors are combined and connected to the correction winding; the first controlled limiter, the first and second adders, wherein the first output of the first controlled limiter is connected to the first input of the first adder, and the second output of the first controlled limiter is connected to the first input of the second adder, the second controlled limiter, similar to the first controlled limiter, the first and second modulators and the third adder, while the output of the first adder is connected to the first input of the second controlled limiter, the output of the second adder is connected to the second input of the second control removable limiter, the first output of the second controlled limiter is connected to the information input of the first modulator, the second output of the second controlled limiter is connected to the information input of the second modulator, the output of the first modulator is connected to the first input of the third adder, and the output of the second modulator is connected to the second input of the third adder, the output of which connected to an electric drive rocket controls; connected in series to a third phase detector, a third filter for extracting an average for a period of an input signal carrier, a third low-pass filter and a first function block, sequentially connected for a fourth phase detector, a fourth filter for extracting an average for a period of an input signal carrier, a fourth low-pass filter and a second function block, at the same time, the information inputs of the third and fourth phase detectors are combined and connected to the bearing winding, and the outputs of the first and second functional blocks are connected inens with the second inputs of the first and second adders, respectively; the control inputs of the first and third, as well as the second and fourth phase detectors are interconnected and connected respectively to the first and second outputs of the reference signal shaper, the control inputs of the first and second modulators are connected respectively to the third and fourth outputs of the reference signal shaper, and the winding of the rotor position sensor the gyroscope and the position sensor of the rocket body are connected respectively to the first and second inputs of the driver of the reference signals, the RNU circuit is made in the form of connected x to the output of the first low-pass filter of the first sequential branch, consisting of the first managed key, the first storage device, the fourth adder, the second managed key and the fifth adder, the output of which is connected to the first input of the first controlled limiter, and the second connected to the output of the second low-pass filter a serial branch consisting of a third managed key, a second storage device, a sixth adder, a fourth managed key and a seventh adder, the output of which o connected to the second input of the first controlled limiter, the output of the first managed key connected to the second input of the fifth adder, and the output of the third managed key connected to the second input of the seventh adder; the second inputs of the fourth and sixth adders are connected respectively to the output of the first additional control signal generating unit and the output of the second additional control signal generating unit, and the third inputs of the fourth and sixth adders are connected respectively to the output of the third low-pass filter and the output of the fourth low-pass filter; and the control inputs of the first, second, third, and fourth controlled keys are combined and connected to the output of the timer, the input of which is connected to the descent sensor.

The essence of the proposed technical solution is presented in the drawings, where Fig. 1 shows the layout of the proposed rocket, Fig. 2 shows the structural diagram of the OGS.

In FIG. 1 and FIG. 2 adopted the following notation:

1 - engine;

2 - wing block;

3 - rocket housing position sensor;

4 - electric drive rocket controls;

5 - OGS;

6 - adjustable gyroscope;

7 - correction winding;

8 - bearing winding;

9 - winding of the gyro rotor position sensor;

10 - driver of the reference signals;

11 - vanishing sensor;

12 - the first phase detector;

13 is a first filter selection average for the period of the carrier of the input signal;

14 - the first low-pass filter;

15 - second phase detector;

16 is a second filter selection average for the period of the carrier of the input signal;

17 - the second low-pass filter;

18 - the first controlled limiter;

19 - the first adder;

20 - second adder;

21 - the second controlled limiter;

22 - the first modulator;

23 - second modulator;

24 - the third adder;

25 - the third phase detector;

26 is a third filter selection average for the period of the carrier of the input signal;

27 - the third low-pass filter;

28 - the first functional unit;

29 - the fourth phase detector;

30 - the fourth filter selection average for the period of the carrier of the input signal;

31 - the fourth low-pass filter;

32 - the second functional unit;

33 - the first managed key;

34 - the first storage device;

35 - the fourth adder;

36 is a second managed key;

37 - fifth adder;

38 - the third managed key;

39 is a second storage device;

40 - sixth adder;

41 - the fourth managed key;

42 - seventh adder;

43 - the first block generating an additional control signal;

44 is a second block for generating an additional control signal;

45 - timer.

The functioning of the nodes of a rotating homing missile is as follows.

In the initial state, the first and third managed keys 33 and 38 are closed, and the second and fourth managed keys 36 and 41 are open. The rocket control signal from correction winding 7 at the gyro rotor rotational speed corresponding to the magnitude and direction of the angular velocity of the target-rocket line of sight is fed to the information inputs of the first and second phase detectors 12 and 15:

where U _SK - the amplitude of the signal corresponding to the angular velocity of the line of sight "missile-target";

ω _gear - rotational speed of the gyro rotor;

t is the current time;

γ is the phase corresponding to the direction of the angular velocity of the missile-target line of sight.

The control inputs of these phase detectors from the first and second outputs of the reference signal generator 10 respectively receive orthogonal reference signals at the rotational speed of the gyro rotor. These reference signals are generated on the basis of trigonometric transformations of the signals from the winding of the position sensor of the rotor of the gyroscope 9 and from the position sensor of the rocket body 3 to the first and second inputs of the driver of the reference signals 10:

where ω _gear , ω _r , ω _gon - rotational speed of the gyroscope and rocket rotor and their total frequency;

θ _dates — angular error of the rocket hull position sensor;

t is the current time.

The output signals of the first and second phase detectors 12 and 15, displaying information about the magnitude and direction of the angular velocity of the line of sight "missile-target" contained in the signal from the correction winding 7, in a fixed Cartesian coordinate system, are fed to the first and second filters the carrier period of the input signal 13 and 16, respectively, which further do not pass the harmonics of the signals at frequencies that are multiples of the rotational speed of the gyro rotor:

where U _SK - the amplitude of the signal corresponding to the angular velocity of the line of sight "missile-target";

γ is the phase corresponding to the direction of the angular velocity of the missile-target line of sight;

θ _dates is the angular error of the rocket hull position sensor.

Then the signals are fed to the first and second low-pass filters 14 and 17, which determine the noise band of the rocket’s autopilot, and then through the first and third controlled keys 33 and 38 to the first and second storage devices 34 and 39 and simultaneously through the fifth and seventh adders 37 and 42 - to the first controllable limiter 18, which protects the circuit from overloads, including impulse noise.

On the first and second adders 19 and 20, the missile control signal generated on the correction winding 7 is added to the signal from the outputs of the first and second functional blocks 28 and 32, respectively, intended, if necessary, to correct the missile control signal according to information about the bearing angle and its derivatives contained in the signal from the winding of the bearing 8, which is fed to the information inputs of the third and fourth phase detectors 25 and 29 and passes, respectively, through the third and fourth filters to select the average for d carrier input 26 and 30, third and fourth low-pass filters 27 and 31:

where U _SP - the amplitude of the signal corresponding to the value of the angle of the bearing;

η is the phase of the bearing signal;

θ _dates is the angular error of the rocket hull position sensor.

From the outputs of the first and second adders 19 and 20, the control signal is supplied to a second controlled limiter 21, similar to the first controlled limiter 18, and then to the information inputs of the first and second modulators 22 and 23. To the control inputs of these modulators from the third and fourth outputs of the reference signal former 10, respectively, the orthogonal reference signals at the rocket speed are received, obtained on the basis of the signal from the position sensor of the rocket body 3 to the second input of the reference oscillator caught 10:

where ω _p , is the frequency of rotation of the rocket;

θ _dates — angular error of the rocket hull position sensor;

t is the current time.

At the outputs of the first and second modulators 22 and 23, a control signal is generated at the rocket speed, which is fed to the third adder 24, and then to the electric drive of the rocket controls 4:

where U _ap - the amplitude of the control signal;

ω _p , is the frequency of rotation of the rocket; t is the current time;

γ is the phase corresponding to the direction of the angular velocity of the missile-target line of sight;

θ _dates is the angular error of the rocket hull position sensor.

In the initial phase of the flight, the missile is controlled by signals generated by the RNU circuit.

At the time of the descent of the rocket, the descent sensor 11 generates a start signal for the timer 45, tuned to the estimated delay time. At the same time, at the moment of the rocket descent, the first and third controlled keys 33 and 38 open, while the rocket control is turned off by the signal from the correction winding 7, and the second and fourth controlled keys 36 and 41 are closed. In this case, signals are stored on the first and second storage devices 34 and 39, containing information about the magnitude and direction of the angular velocity of the line of sight "missile-target" at the moment of the descent of the rocket, which determines the required lead angle. At the fourth and sixth adders 35 and 40, these signals are summed with an additional control signal supplied to the second inputs of these adders from the outputs of the first and second blocks for generating an additional control signal 43 and 44, respectively. An additional control signal does not depend on the firing conditions and is necessary to prevent contact of the rocket with the ground at the initial stage of the flight. The specified signal, for example, can be proportional to some structurally specified initial angle between the axis of the OGS gyroscope and the longitudinal axis of the rocket, which provides the initial angle of elevation of the rocket during launches on low flying targets:

where K ₁ , K ₂ - transmission coefficients of the additional control signal;

U _cn0 is the signal amplitude corresponding to the initial angle of elevation of the rocket.

Thus, the signals at the first and second inputs of the fourth and sixth adders 35 and 40 form a signal of the program bearing, which is proportional to the required angles of lead and elevation of the rocket, respectively. The third inputs of the fourth and sixth adders 35 and 40 from the outputs of the third and fourth low-pass filters 27 and 31, respectively, receive a signal proportional to the current angle of the bearing in the fixed Cartesian coordinate system, which is formed on the basis of the signal from the coil of bearing 8, which at the time of descent missiles unambiguously characterizes the position of the target, which is in the field of vision of the OGS, relative to the axis of the missile, including when the "capture" of the target is carried out by an external target signal from the carrier.

The polarity of the signals arriving at the fourth and sixth adders 35 and 40 and the transmission coefficients are selected in such a way that a control signal proportional to the difference between the signal of the program bearing proportional to the required lead and elevation angles and the signal proportional to the current angle of the bearing is generated at the outputs of these adders. the latter strives to become equal to the angles of lead and elevation:

where K ₃ is the transmission coefficient of the fourth and sixth adders 35 and 40, respectively;

U _{cr_c} is the signal amplitude corresponding to the value of the angular velocity of the line of sight "missile-target" at the time of missile launch;

γ is the phase corresponding to the direction of the angular velocity of the line of sight "missile-target" at the time of the descent of the rocket;

K ₁ , K ₂ - transmission coefficients of the additional control signal;

U _cn0 is the signal amplitude corresponding to the initial angle of elevation of the rocket;

U _SP - the amplitude of the signal corresponding to the value of the angle of the bearing;

η is the phase of the bearing signal;

θ _dates is the angular error of the rocket hull position sensor.

The specified control signal then goes according to the scheme to the electric drive of the missile control bodies 4, while the axis of the missile at the initial section of the flight unfolds at the anticipated meeting point of the missile for the purpose.

After the time set on timer 45 expires, the second and fourth controlled keys 36 and 41 open until the end of the flight, the RNU circuit is completely turned off, and the first and third controlled keys 33 and 38 are closed until the end of the flight, and the rocket is controlled by a signal from correction winding 7, the corresponding magnitude and direction of the angular velocity of the missile-target line of sight.

Thus, in the proposed rocket, the control signal generated by the RNU circuit is covered by feedback along the angle of the bearing, which ensures the simultaneous performance of two functions:

- automatic rotation of the rocket to the required lead and elevation angles with a minimum dependence of the influence of various kinds of external disturbing influences on the rocket trajectory in the initial phase of flight;

- the ability to launch a rocket both from the shoulder and from various carriers, as in the case when the "capture" of the target and the subsequent launch of the rocket occurs by turning the axis of the OGS gyroscope relative to the longitudinal axis of the rocket on the target along the target signal from the carrier external to the rocket so that the target is in the OGS field of view, regardless of the position of the sighting axis of the MANPADS sighting device, and in the case when the "capture" of the target and the subsequent launch of the rocket occurs by turning the launch tube on the target so that The target was in the field of view of the MANPADS sighting device.

As a result, the proposed rotating homing missile allows launching both from the shoulder and from various carriers, including in the presence of target designation signals from the carrier for each missile, which facilitates aiming and control of the carrier during launch, and reduces preparation and holding time launch, the ability to carry out both volley launch of missiles at one target, and simultaneous launch of missiles at two targets from one carrier, with automatic rotation of the rocket (missiles) to the required lead angles For each specific combination of starting conditions.

Claims (1)

A rotating homing missile containing an engine, a wing unit, an instrument compartment including a rocket housing position sensor and a rocket control electric drive, an optical homing head including an adjustable gyroscope with a correction winding, a bearing winding, a gyro rotor position sensor winding, a reference signal conditioner, a descent sensor and an autopilot with a missile pivoting circuit in the initial section (RNU), the functional diagram of which contains a first phase detector connected in series, the first filter is the selection of the average for the period of the carrier of the input signal and the first low-pass filter, connected in series with the second phase detector, the second filter is the selection of the average of the period of the carrier of the input signal and the second low-pass filter, while the information inputs of the first and second phase detectors are combined and connected to the winding correction; the first controlled limiter, the first and second adders, wherein the first output of the first controlled limiter is connected to the first input of the first adder, and the second output of the first controlled limiter is connected to the first input of the second adder, the second controlled limiter, similar to the first controlled limiter, the first and second modulators and the third adder, while the output of the first adder is connected to the first input of the second controlled limiter, the output of the second adder is connected to the second input of the second control removable limiter, the first output of the second controlled limiter is connected to the information input of the first modulator, the second output of the second controlled limiter is connected to the information input of the second modulator, the output of the first modulator is connected to the first input of the third adder, and the output of the second modulator is connected to the second input of the third adder, the output of which connected to an electric drive rocket controls; connected in series to a third phase detector, a third filter for extracting an average for a period of an input signal carrier, a third low-pass filter and a first function block, sequentially connected for a fourth phase detector, a fourth filter for extracting an average for a period of an input signal carrier, a fourth low-pass filter and a second function block, the information inputs of the third and fourth phase detectors are combined and connected to the bearing winding, and the outputs of the first and second functional blocks are connected Nena to second inputs of the first and second adders, respectively; the control inputs of the first and third, as well as the second and fourth phase detectors are interconnected and connected respectively to the first and second outputs of the reference signal shaper, the control inputs of the first and second modulators are connected respectively to the third and fourth outputs of the reference signal shaper, and the winding of the rotor position sensor the gyroscope and the position sensor of the rocket body are connected respectively to the first and second inputs of the driver of the reference signals, characterized in that the RNU circuit is made in the form of a first serial branch connected to the output of the first low-pass filter, consisting of a first controlled key, a first storage device, a fourth adder, a second managed key and a fifth adder, the output of which is connected to the first input of the first controlled limiter, and connected to the output of the second filter low frequencies of the second sequential branch, consisting of a third managed key, a second storage device, a sixth adder, a fourth managed key and a seventh Matora whose output is connected to the second input of the first managed limiter, the output of the first controllable switch is connected to the second input of the fifth adder and the output of the third controllable switch is connected to the second input of the seventh adder; the second inputs of the fourth and sixth adders are connected respectively to the output of the first block for generating an additional control signal and the output of the second block for generating an additional control signal; the third inputs of the fourth and sixth adders are connected respectively to the output of the third low-pass filter and the output of the fourth low-pass filter, and the control inputs of the first, second, third, and fourth controlled keys are combined and connected to the output of the timer, the input of which is connected to the descent sensor.

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