RU2498192C2 - Principle of optic beam guidance of missile launching from mobile carrier - Google Patents

Abstract

FIELD: weapons and ammunition.

SUBSTANCE: principle of optic beam guidance of a missile launching from a mobile carrier includes formation on the carrier of a laser beam with an information control field, guidance on a target of an optic sight, the axis of which is aligned with the information field axis, orientation of the launching system axis in the beam axis direction, launch of the missile with retracted aerodynamic rudders and introduction of the missile to the information field, opening on the missile of an emission receiver and shaping of control commands depending on position of the missile relative to the information field axis, extension of aerodynamic rudders and their deviation. At the moment of the missile ejection an angular position of the launching system axis relative to the coordinate system related to the carrier, as well as angular position of the carrier relative to the earth referenced coordinate system is stored; extension of rudders is performed with time delay relative to the emission receiver opening; at the moment of opening of the emission receiver the beam information field axis is aligned with the launching system position stored at the missile ejection moment, and at the point of time of rudders extension the alignment of laser beam information field axis with the target sighting line is started.

EFFECT: improving accuracy and reducing the time required for placing of the missile in the beam axis.

2 cl, 5 dwg

Description

The present invention relates to the field of development of guidance systems with teleorientation of a rocket in a laser beam and can be used in anti-tank missile systems (ATGMs) mounted on land and air carriers.

Currently, there is a known guided missile guidance method implemented, for example, in the Kastet, Bastion, Sturm, and Cobra anti-tank systems (RD Angelsky, Russian anti-tank systems, Moscow, ACT, Astrel, 2002 ., p. 74-75, 84, 100, 111-114, [1]), including orienting the launching device with the rocket in the direction of the control information field created by the laser beam, moving the rocket into the control field by firing from the starting device, subsequent opening radiation receiver, steering wheel opening, ation control commands which are proportional to the coordinates of the missile in the control field, and the deflection under the action of control commands. The main feature of tele-orientation in the information field of the laser beam control is that the flight of the rocket occurs in two stages. At the first stage, the rocket is shot into the information field of the beam, and after it enters the beam control field, the on-board radiation receiver opens and the second stage begins, at which the rocket flies and is controlled in the control field.

Hereinafter, the control field is understood as the region of determining the coordinates of the rocket relative to the line of sight and this can be, for example, the field of view of the direction finder of the rocket in the command control system implemented in the Sturm, Cobra complexes or the information field of the beam (radio beam or laser beam) ), formed by the guidance device in the beam missile control system, implemented in the Kastet, Bastion, and Vortex complexes.

The operation of the known method of guiding missiles is as follows.

The operator, having detected the target, combines the aiming line (optical axis of the sight) and the zero axis of the coordinate determination equipment (optical axis of the laser radiation source) coinciding with it with the aiming point and launches the rocket. The rocket is moved into the control field by firing from a launching device oriented in the direction of the control field at given angles in the horizontal and vertical planes. The launching device may be a transport and launch container mounted on a ground or air carrier or a tank gun.

After the rocket enters the control field, the radiation receiver mounted on the rocket starts to work, and the coordinate determination equipment senses the frequency-modulated laser radiation and, deciphering it, determines the coordinates of the rocket relative to the axis of the information field of the laser emitter (i.e., relative to the line of sight of the target).

The beginning of the operation of the rocket radiation receiver coincides in time with the moment the protective cover of the radiation receiver opens (the rocket tray may play the role of a protective cover (V.I. Babichev, V.V. Vetrov, A.V. Ignatov, A.R. Orlov, Fundamentals of the device and the functioning of artillery guided shells, ed. TulSU, Tula, 2003, p. 71, [2]), the discharge of which occurs after the rocket leaves the channel of the gun’s barrel). After dumping the pallet, the radiation receiver starts functioning and contact groups are closed, one of which supplies voltage to the electric igniter of the rudder opening mechanism (the rudder opening mechanism has single-acting pyrotechnic drives, the operation of which is based on the operation of electric igniters with the subsequent conversion of the energy of expanding gases into mechanical movement ([2 ], p.61), and the steering wheels open.

Control commands proportional to the coordinates of the rocket relative to the axis of the beam are received at the input of the steering gear unit. The rudders of the steering drive unit, deviating from their middle position, create a control moment acting on the rocket, which leads to the emergence of forces that keep the rocket near the center of the control field for the entire duration of the flight of the rocket to the target.

The described control method is implemented, for example, in a Reflex rocket controlled in a direct laser beam, which is fired by a throwing device from the gun’s bore, having a switch electrically connected to the steering wheel opening mechanism (p. 72-74, [2 ]).

The disadvantage of this guidance method is the significant preparation time for launch, associated with the need to turn the launch device using the rotary and lifting guidance mechanisms to set strictly defined angles between the direction of the shot and the line of sight of the target (LCV).

As a prototype of the claimed method, the guidance method closest in technical essence to the proposed invention was implemented in the MAPATS anti-tank missile system (High-precision weapons of foreign countries, volume 1, anti-tank missile systems, review and analytical manual, GUP KBP, Tula, 2008, p. 316-321), which consists in the formation of a laser beam with an information control field, aiming the optical sight and the beam aligned with the sight on the target, shooting the rocket into the beam, opening the radiation receiver, determining the coordinates of the cancer s relative to the axis of the information field beam forming control commands which are proportional to the coordinates of the missile from the axis of the information field, and the transformation of the rudders disclosure control commands to the deflection.

This guidance system, installed on mobile carriers, for example, on helicopters, has the following disadvantages: significant time for operations related to setting the required angles between the guides of the starting device and the LCF; the requirement of pilot skill to set the required angles by piloting a helicopter; relatively low accuracy of installation of the required angles; the need to transfer the carrier to the dive mode at the target in preparation for firing and launching the rocket; a significant beam size, which is due to the significant dispersion of the missiles when they are launched at low initial speeds.

The objective of the invention is to expand the range of conditions for the use of anti-tank systems when firing from an air carrier with increased pilot aiming errors without organizing precise aiming of the launching device relative to the LCA with air carrier maneuvering, as well as reducing the requirements for missile dispersal at the initial guidance site and reducing the smoke of the LCV.

The problem is solved due to the fact that in the method of guiding the optical beam of a rocket starting from a movable carrier, which includes forming a laser beam with a control information field on the carrier, pointing an optical sight at the target whose axis is aligned with the axis of the information field of the laser beam, orienting the axis a launching device in the direction of the axis of the information field of the laser beam, launching a rocket with folded aerodynamic rudders and entering the rocket into the information field of the laser beam, opening on the cancer e of the radiation receiver and generating signals proportional to the coordinates of the rocket relative to the axis of the information field of the laser beam, and converting them into control commands, opening the aerodynamic rudders and converting control commands into steering deflections, additionally, at the moment the rocket leaves the carrier, the angular position of the axis of the launching device is measured and stored relative to the coordinate system associated with the carrier, as well as the angular position of the carrier relative to the earth coordinate system, the disclosure is aerodynamically rudders carried out with a time delay on the opening of the radiation detector in the opening of the radiation receiver aligned axis of the information field of the laser beam stored in a time of de missile axis position of the starting device, and at time t _HB, corresponding to the rudders disclosure, begin combining information field axis of the laser a beam with a target line of sight according to the law:

Δφ _{woo, z} = 0.5 · Δφ ° _{wu, z} · [1 + cos (ω _{wu, z} · (tt _nv ))], for t _nv <t≤t _{kvu, z} ,

Δφ _{wu, z} = 0, for t> t _{kwu, z} ,

where Δφ _{wu, z} are the current angles between the axis of the information field and the line of sight of the target,

Δφ ° _{wu, z} are the flap angles of the axis of the information field from the line of sight of the target at the moment the radiation receiver was opened on the rocket,

Δφ _{vmu, z} - maximum angular velocity of rotation of the axis of the information field in the process of combining;

, при t=t_НВ,

, at t = t _HB ,

, при t=t_НВ,

, at t = t _HB ,

,

- угловые скорости вращения линии визирования цели по углам места и азимута соответственно,

,

- the angular velocity of rotation of the line of sight of the target at elevation and azimuth angles, respectively,

t _{kvu, z} is the end time of the combination,

t _{qy, z} = t _HB + Δt _{woo, z} ,

where Δt _{wu, z} = π / ω _{wu, z} .

The proposed method of guiding the missile is implemented in the anti-tank complex "Whirlwind", located on a helicopter. When working with an anti-tank guided armament "Whirlwind" from a helicopter, the pilot, when detecting a target using the helmet-mounted target designation system, orientes the line of sight of the Shkval-V complex in the direction of the target and, switching the field of view from wide to narrow (having a larger increase), recognizes the target and after determining the purpose of the target, the enemy decides to attack and selects a weapon ([1], p.171-179, Mazepov A. et al., “Ka-50”, M., Favorite Book, 1997, p. 84- 91). If a rocket is chosen as a weapon, the pilot with the joystick on the helicopter control knob adjusts the position and size of the reticle on the television indicator so that it lies on the target and the target occupies about Ud of its area, and then puts the equipment into auto tracking mode. In this case, the laser range finder measures the distance to the target. Teleautomaton (TA) remembers the image of the target and provides further support. The system provides automatic tracking of the target even after a brief disappearance of its image - for example, in cases where another object is between the helicopter and the target, for example, a tank. In the event that, nevertheless, an automatic tracking failure occurred, the pilot re-captures the target for auto tracking. The launcher fulfills the elevation angle of the target. Upon reaching the permitted range on the windshield indicator (ILS) and the television indicator, signals appear allowing the opening of fire.

When firing a rocket, it is necessary to ensure that the missile is shot in the control field of the laser beam. By piloting the helicopter, the pilot must combine on the HUD the symbols of the line of sight of the target and the launch zone. If this does not happen, then a lock is triggered and the missile launch enable signal is not displayed. If the TA for some reason has lost the target, then it is possible with the joystick to adjust the position of the aiming mark on the target and again switch to auto tracking.

Prior to the implementation of the invention, the introduction of the rocket into the information field was provided by the maneuver (aiming) of the helicopter, during which the pilot, having on the windshield indicator information about the angular position of the line of sight of the target (TA axis and the axis of the information field aligned with it) in the form " movable mark "and the angular position of the launch container in the form of an" aiming mark "with a ring of a shooting zone around it with a radius of, for example, 1 °, acting on the helicopter controls, combined the aiming mark with under moving mark in vertical and horizontal planes. When a movable mark got into the ring of the shooting zone, the pilot pressed the rocket launch button and, until it left, ensured that the moving mark was kept in the ring of the shooting zone (~ 0.8 ... 1.0 s after pressing the button).

The proposed guidance method allows you to launch a missile without precise aiming by helicopter maneuvering when implementing the software alignment of the axis of the control beam with respect to the TA axis in the vertical and horizontal planes. At the moment of rocket descent, the angular position of the helicopter ϑ _н0 , ϑ _н0 and the position of the launching device relative to the construction horizontal of the fuselage (GFS) of the helicopter are fixed. At the moment of time corresponding to the moment of opening the cover of the radiation receiver, two optical wedges are introduced into the optical scheme of the guidance device, which divert the axis of the transmitter to the position of the axis of the launching device fixed at the time of launch, thereby ensuring rocket “trapping” along the trajectory.

SP needs angles flap axis relative to axis TA to provide input to the rocket SP _count Δφ °, Δφ ° _BZ in the vertical and horizontal planes must be calculated by relationships:

Δφ ° _wu = ϑ _{n0 -ϑ n} + ϑ _pu -φ _y -ΔΘ,

Δφ ° _bz = ψ _n0 -ψ _n + ψ _pu -φ _z -Δφ,

where ϑ _н0 , ψ _н0 are the pitch and yaw angles of the helicopter at the time of rocket descent;

ϑ _n , ψ _n - the current pitch and yaw angles of the helicopter;

ϑ _pu , ψ _pu - angles of rotation of the axis of the launching device relative to the GFS helicopter in the vertical and horizontal planes, respectively;

φ _y , φ _z are the current angles of rotation of the TA axis relative to the GFS helicopter in the vertical and horizontal planes, respectively;

ΔΘ, Δφ - the required angles of rotation of the axis of the launching device relative to the axis of the FE in the vertical and horizontal planes, respectively, at the time of the descent of the rocket, calculated in the computing unit of the carrier.

At the time tнв corresponding to the moment of opening the rudders, the alignment of the IP axis with the line of sight of the target begins. In the process of combining, the angular mismatches Δφ _{wu, z} must vary according to the law:

Δφvu, z = 0.5 · Δφ ° wu, z · [1 + cos (ωvu, z · (t-tnv))], with tnv <t≤tquu, z,

Δφv, z = 0, for t> tqu, z,

Where

ω _{vmu, z} - maximum angular velocity of rotation of the axis of the PI in the process of combining;

, при t=t_НВ,

, at t = t _HB ,

, при t=t_НВ,

, at t = t _HB ,

,

- угловые скорости вращения линии визирования цели по углам места и азимута соответственно, рассчитываемые в вычислительном блоке носителя по алгоритмам математического обеспечения.

,

- the angular velocity of rotation of the line of sight of the target at elevation and azimuth angles, respectively, calculated in the computing unit of the medium according to mathematical support algorithms.

When ωvu, z> π, ωvu is taken, z = π, and the maximum angular velocity of alignment of the IP axis is determined by the dependence:

tqu, z - the end time of the combination:

tqua, z = tnv + Δtvu, z,

where Δtvu, z = π / ωvu, z

The invention is illustrated by drawings.

1, 2 illustrate the relative angular position of the helicopter, the axes of the teleautomaton, the pointing device, the launching device, the horizon and the target (for the vertical plane).

Figure 1 shows two positions of the rocket on the trajectory: the first is the position of the rocket at time t = 0.3 s, preceding the start of alignment of the IP and TA axes, and the second is the position of the rocket at the end of the alignment of IP and TA axes. Figure 2 marked:

ε _c - the angle between the target line and the horizon line,

φ _Y - between the TA axis (target line) and the GFS carrier,

φ _Y ^PM - the angle between the axis of the reticle and the GFS helicopter,

ε _BB - the estimated angle of deviation PU relative to the line of the target,

υ _n - the angle between the horizon line and the GFS helicopter,

υ _n0 - the angle between the horizon line and the GFS helicopter at the time of launch,

υ _PU - the angle between the GFS helicopter and the axis of the launch device,

Δφ ° _wu - the required angle of deviation of the axis of the PI relative to the axis of the TA to ensure the shooting of the rocket in the PI,

Δφ is the error of pointing the axis of the beam IP,

ΔΘ - the required angular position of the PU relative to the axis of the IP.

Figure 3 shows a block diagram of a guidance system that implements the proposed guidance method. Figure 3 introduced the following notation: 1 - on-board digital computer (BCM), which includes the following blocks: 2 - block algorithm shooting; 11, 16 - subtraction blocks (BV1, BV2); 17 - adder (C), 22 - command generator (FC) control the actuator of the starting device. The algorithm of the shooting algorithm can be made up of a set of subtracting blocks, adders, product blocks, and constant storage blocks. Adders and subtraction blocks are made according to the scheme of Fig. 11.1 (W. Titze, K. Schenk, Semiconductor circuitry, M., Mir, 1982, p.137, [3]). Blocks of the product can be performed according to the quadrant multiplication scheme ([3], p. 162, Fig. 11.41). Constant storage units can be made in the form of read-only memory ([3], p.125-127). Block 24 - actuator of the starting device, block 25 - sensor of the angular position of the axis of the starting device (DUP3).

The guidance system includes a teleautomaton unit 3, which is a television equipment (TAp.) 4, coupled with a laser rangefinder, target designator and a device for guiding missiles along the laser beam 5 ([1], pp. 84-85). The television equipment of the complex has a wide and narrow field of view, the deviation angles of the line of sight in azimuth ± 35 °, in elevation from + 15 ° to -80 °. The angular deviations of the axis of the machine through the angle sensor 8 are converted into electrical signals. The sensors of the angular position of the axis of the teleautomaton and the axis of the starting device can be made, for example, on the basis of a photoelectric angle to code converter (Fundamentals of Theory and Design of Computing Devices and Control Machines, edited by L.N. Presnukhin, M., Higher School, 1970 ., p. 199, 325).

A television can perform the functions of a search and search system. The line of sight to search for the target deviates automatically or manually from the joystick. After recognition of the target and its capture, the target is tracked automatically. The television can be made on the basis of a television sight, presented in the book of V.P. Demidov, N.Sh. Kutyev, Anti-aircraft missile control, M., Military Publishing House, 1989, p.184-186.

On the indicator of the windshield 12 displays information about the position of the target and the aiming mark of the starting device. Based on this information, the pilot (the pilot, as an element of the control system, is indicated on the diagram by block 18) must ensure that the missile is shot in the field of view of the search and search system. By piloting a helicopter 23, it must combine the symbols of the line of sight of the target and the launch zone on the ILS. After the launch of the ATGM, the pilot must pilot the helicopter within the permissible limits of the deviation angles of the target line of sight.

The axis of the teleautomaton is aligned with the axis of the IP of the laser beam generated by the guidance device 5, which includes two pulse sources (AI) 6, 10, a pulse code generator (FIC) 7, a scanning unit (BS) 13, an angular position sensor (DUP2) of the beam axis 14, the first synchronizer (C1) 15, the lettering unit (BL) 21, the optical panoramic system (OPS) 19, aligned with the optical sight 20. The missile is represented by block 26, which includes a radiation receiver (PI) 27, a signal gating block (BSS) 28, block of synchronizing pulses (BSI) 29, the second sync a reducer (C2) 30 and a pulse code analyzer (AIC) 31. The device and operation of the guidance device are disclosed in the description of the invention presented in RF patent 2382315 "Guided projectile guidance system".

,

.Figures 4.1 and 4.2 show the trajectories of the Whirlwind rocket in the vertical and horizontal planes, respectively, obtained as a result of mathematical modeling of the guidance process with pilot aiming errors in the vertical plane of -2.5 °, in the horizontal plane of ± 1.0 ° under the following conditions: charge temperature, air Тз, в = 20 ° С; crosswind Wz = ± 10 m / s. Initial disturbances during the descent of the rocket:

,

.

,

, Wz=0.Trajectory 1 was obtained for the conditions: Θ ₀ = 1.5 °, φ ₀ = -0.7 °,

,

, Wz = 0.

,

, Wz=10 м/с.Trajectory 2 was obtained for the conditions: Θ ₀ = 1.5 ° + 0.33 °, φ ₀ = -0.7 ° + 1.0 °,

,

, Wz = 10 m / s.

,

, Wz=-10 м/с.Trajectory 3 was obtained for the conditions: Θ ₀ = 1.5 ° -0.33 °, φ ₀ = -0.7 ° -1.0 °,

,

, Wz = -10 m / s.

Figures 5.1 and 5.2 show the trajectories of the Whirlwind rocket, respectively, in the vertical and horizontal planes, obtained with errors in aiming the pilot in the vertical plane + 2.5 °, in the horizontal plane ± 1.0 ° for three trajectories under the same conditions.

In the simulation, the most severe case of firing in the helicopter “hover” mode, which is characterized by the greatest dispersion of missiles in an uncontrolled flight section, was considered in terms of the most difficult missile entry.

Analysis of the results of mathematical modeling shows that when implementing the proposed method for guiding the Vikhr missile, for example, from a Ka-52 helicopter and pilot aiming errors of up to ± 2.5 °, the missile is entered into the information field of the control beam and then guided to the target. The proposed method allows to reduce the requirements for the dispersion of missiles at the initial guidance site, as well as reduce the requirements for precision aiming by aircraft carriers, i.e. the opportunity is realized to aim and launch missiles into the information field of the control beam without carrier maneuver.

Claims (2)

1. The optical beam guidance method of a rocket launching from a movable carrier, comprising forming a laser beam with a control information field on the carrier, aiming an optical sight on the target whose axis is aligned with the axis of the laser beam information field, orienting the launch device axis in the direction of the information field axis a laser beam, launching a rocket with folded aerodynamic rudders and entering the rocket into the information field of the laser beam, opening the radiation receiver on the rocket and generating signals, proportional to the coordinates of the rocket relative to the axis of the information field of the laser beam, and converting them into control commands, opening the aerodynamic rudders, converting control commands to steering deviations, characterized in that at the time of the missile’s departure on the carrier, the angular position of the axis of the launching device relative to the carrier is measured and stored coordinate systems, as well as the angular position of the carrier relative to the earth coordinate system, the disclosure of the rudders is carried out with a time delay relative to the open When the radiation receiver is opened, the axis of the information field of the laser beam is combined with the position of the axis of the launcher memorized at the time of the launch, and at time t _HB , corresponding to the moment the rudders open, the axis of the information field of the laser beam begins to align with the line of sight of the target. 2. The method according to claim 1, characterized in that the combination of the axis of the information field of the laser beam with the line of sight of the target is carried out according to the law:

at t _nv <t≤t _{quu, z} , Δφ _{wu, z} = 0, at t> t _{quu, z} ,
where Δφ _{wu, z} are the current angles between the axis of the information field and the line of sight of the target,

- the angle of the lapel axis of the information field from the line of sight of the target at the time of opening of the radiation receiver on the rocket,

ω _{bmu, z} are the maximum angular rotational speeds of the axis of the information field in the process of combining;

at t = t _HB ,

, at t = t _HB ,

- the angular velocity of rotation of the line of sight of the target at elevation and azimuth angles, respectively,
t _{kvu, z} is the end time of the combination,
t _{qy, z} = t _HB + Δt _{woo, z} ,
where Δt _{wu, z} = π / ω _{wu, z} .

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