RU2400690C1 - Aa missile guidance system - Google Patents

Abstract

FIELD: weapons and ammunition.

SUBSTANCE: proposed system compares signals from optical and IR digital photo cameras and radar signal to differentiate between true and false targets. System generates anticipation trajectory via feedback between rudders with movable homing head, that is, the latter turns in direction opposite the rudder deviation unless the rudders stay in neutral position. System can perform leading anticipation towards airframe via shifting rudders position transducer neutral toward the side coinciding with homing head deviation, or its additional deviation toward the same side.

EFFECT: high probability of heating maneuvering target.

4 cl, 3 dwg

Description

The invention relates to air-to-air and ground-to-air missiles with all types of homing heads (hereinafter referred to as GOS).

Missiles with thermal seekers are known (see “History of Aircraft Arms”, Minsk, 1999, p. 444) containing a fuselage, an engine, an infrared or radar target sensor, power steering amplifiers and drives, but they can be removed from the target by thermal traps or the sun . Missiles with trajectory correction in terms of gyroscope precession speed are known (see ibid., P. 417), but this system is complex and not accurate enough, which can lead to a miss with an energetic maneuver of the target aircraft.

The objective of the invention is to increase the likelihood of a missile entering a maneuvering target against a background of interference. This problem is solved jointly in two ways. Firstly, the implementation of electronic discrimination of false infrared targets. And secondly, a more accurate guidance of the missile along the intersecting trajectory, and even better - along a slightly ahead trajectory. At the same time, traps quickly leave the missile’s seeker’s field of view, and the rudders of the rocket are practically in a neutral position, which leads to an increased readiness of the rocket to perform maximum maneuver in any direction.

Invention 1. In addition to amplifiers and rudder drives, the proposed system contains two digital cameras as a target sensor, one of which operates in the optical range, and the other in the infrared (hereinafter, “optical camera” and “infrared camera”). The pixels of these cameras are connected by a threshold signal transmission unit (hereinafter PPS) of the optical camera (for example, using dynistors) and a block for switching off the corresponding infrared pixels (hereinafter VIP) of the infrared camera (for example, a two-transistor electronic key circuit).

That is, the signal from the pixels of the optical camera does not go further until its level reaches a certain brightness (brighter than the signal from the jet engine nozzle of the aircraft, sky, clouds). If the signal exceeds this brightness, for example, a signal from the sun, from a heat trap, then it passes through the PPS block almost without attenuation and enters the VIP block, which disables the image from the same section of the infrared camera, see Fig. 1.

That is, where there is bright illumination on the virtual image of the optical camera, a black spot is “cut out” on the same section of the infrared camera, and the rocket does not “see” the source of infrared radiation, if it is also a source of visible radiation. Thus, the rocket does not respond to the sun, traps and burning aircraft.

It is necessary to foresee the enemy’s countermeasures in advance: in order to pass the true target as false, it is enough to increase the luminosity of the aircraft nozzle, for which you can blow aluminum powder or just an additional amount of fuel into the nozzle. In this case, the system in the virtual infrared image "cuts out" the black spot at the nozzle of the aircraft and there will be no infrared signals.

If this happened close enough to the aircraft, then the rocket will not be fooled - with sufficient sensitivity, it will redirect to the front edges of the wings or blades, or to the air intakes. But if the target is still far away, and it is identified as a point object, this can deceive the rocket.

To avoid this, the guidance system has an electronic control key (hereinafter ESC), which, by a zero signal (no signal) from the infrared camera through the delay line (for example, a time relay for 0.001 s), turns off the optically visible channel (for example, a VIP unit), and the rocket sees all the infrared targets again. Then, the ECU again turns on the optical channel, and the infrared channel again “fades”. In such a pulsating mode, the rocket will nevertheless confidently aim at the most powerful source of infrared radiation until the infrared camera captures the input edges of the wings. Or the missile will be pointed to the most powerful heat source to the end.

The retail price of digital cameras fell to 2,000 rubles, and the size of cameras built into mobile phones with a resolution of 2 megapixels approached the size of a pea. Therefore, the proposed part of the guidance system will have a size of a thimble, a weight of several grams, and a cost of about 10,000 rubles.

If the GOS is combined and has, in addition to the optical and thermal channels, an active or semi-active radar station (hereinafter referred to as the radar), then the reliability and noise immunity of the guidance can be significantly increased. To do this, the selective optical-infrared signal about the target and the signal of the radar channel in the same format and scale are fed to the I-YES logical unit, the signal from which is then sent to the system for execution, to the steering amplifiers and drives.

That is, a missile is aimed only at that target that emits infrared radiation, does not have strong optical radiation and reflects an active or passive radar signal.

Such a combined scheme is especially useful in cloudy weather: if the plane, having detected the launch of a rocket, dives into the clouds, the capture of the thermal seeker can be disrupted. And the presence of a radar channel will continue the attack. Accordingly, the presence of a thermal channel allows the rocket to be insensitive to artificial and natural interference in the radio channel.

Invention 2. Guidance of the rocket on the gyroscope precession rate is not high enough. The proposed rocket has a simple and reliable, not afraid of an electron impulse, a system for obtaining intersecting trajectories. The system consists of a movable in two planes GOS of any type, an amplifier, rudder drives, rudder position sensor and GOS drives. For a missile with a cruciform wing, two such channels are needed - horizontally and vertically.

The algorithm of the system is as follows: after launch, the GOS controls the rocket, deflecting the rudders. But the GOS itself deviates in the direction opposite to the deviation of the rudders (with the aerodynamic design “weathervane duck”, and with the rear and gas rudders - vice versa), and with a speed proportional to the deviation of the rudders. That is, in conjunction with the GOS drive, accumulating the deviation, a proportional-integral (“PI-regulation”) course angle of the target relative to the rocket occurs. The deviation of the seeker will increase until the sensors deviation of the rudders from "zero" (neutral position) does not show "0", that is, the rudders will become in the neutral position. After that, the GOS will remain in the same position, and the rocket will fly along a direct path. In this case, the target angle with respect to the rocket will be constant. What, as you know, leads to hitting the target, see figure 2.

It is advisable that the rocket does not rotate, at least faster than 0.2 revolutions per second. Special measures for this can not be taken. It is enough to observe the accuracy of manufacturing and to carry out a control purge of the rocket in the wind tunnel. Although, of course, it is more reliable to have roll stabilization with the help of “scissors” and rudders.

Missile miss analysis showed that, as a rule, missiles pass behind targets. This is because signal processing by the guidance system takes time. There are guidance correction systems, for example, the shift of guidance from the nozzle to the fuselage, but they are quite complicated. The proposed rocket has a simple and reliable correction of the intersection trajectory by a small lead.

To do this, the described system additionally contains a mechanism or an electronic element (for example, a bridge electrical circuit) that biases the “0” of the rudder position sensor by a fixed or speed-dependent value (for example, by 0.1 degrees) in the same direction in which the GOS is turned relative to the longitudinal axis of the rocket (see figure 3 by a dotted line). Or after the rudders got to "0", additionally shifts the seeker in the same direction.

As a result, the rocket flies with a slightly greater lead than necessary and would fly ahead of the target if it were not for a constant flight in a very gentle arc. At the final stage of the flight, the missile “underregulates” and will fall 2-3 meters in front of the radiation source (in front of the nozzle, in front of the center of the effective area of radar scattering).

It should not be feared that the presence of a GOS rotation mechanism, the speed of which in order to avoid overshooting, should be less than the speed of the rudders, but more than the speed of the rocket's reaction to the rudders, will reduce the maneuverability of the rocket. This will not happen - the GOS will always track the target ahead of schedule, and rudder performance will remain at the same level.

For a missile with a flat wing, the system will have a slightly different look. GOS should be controlled in two planes and roll, that is, the roll of the rocket should lead to the same roll in the same direction of the GOOS relative to its axis. GOS roll can be performed not mechanically, but virtually - by shifting the image scan orientation. The rocket should still have two control channels, but not horizontally and vertically, but in pitch and roll. For this, it should have only two separately controlled (left and right) horizontal aerodynamic and / or gas steering wheels. That is, the whole difference is that the missile is controlled by yaw not by deflecting the vertical rudders, but by a proportional roll (up to 90 degrees) and a corresponding increase in pitch. Otherwise, the system is identical to the above with the difference that the correction of the advance path is made by a small offset “0” of the roll sensor towards the deviation of the seeker. Or, as in the version with a cruciform wing, an additional shift of the seeker toward the target.

Figure 1 shows a block diagram of the guidance (fragment), consisting of optical and infrared cameras OFK and IFC, a block of threshold transmission of signals PPS, block off infrared pixels VIP, electronic control key ESC, delay line LZ, and may additionally have a radar station Radar and logical unit "I-YES."

Figure 2 shows the process of guiding a rocket to a lead point, where: 1 - rocket, 2 - GOS, 3 - rudders, 4 - target.

Figure 3 shows a block diagram of a guidance system (fragment — only lead system) in one direction, where: GOS — homing head, P — head drive, US — amplifier, CH — zero-shift unit for the rudder position sensor DR.

The system of FIG. 1 operates as follows: the signal from the OFK optical camera through the threshold transmittance block of the PPS signals is fed to the VIP infrared pixel shutdown unit, which “cuts out” the place corresponding to the optical signal in the image of the IFC infrared camera. In the absence of a signal from the IFC, the electronic control key of the ESC through the delay line LZ periodically disconnects the VIP unit, and the signal from the IFC becomes pulsating, which does not interfere with aiming at the target.

Additionally, the system may have a radar, the signal from which is fed to the I-YES unit, from where, if there is a signal from the IFC, the logical signal is then sent to the system for execution.

After the launch of rocket 1 in figure 2, 3 on target 4, flying to the left, the seeker 2 gives a signal, and the steering wheels 3 turn to the left. In this case, the rudder position sensor DR generates a signal to the power amplifier, and drive P turns the seeker to the right. But the GOS seeks to keep the target in the center of its field of vision and therefore commands the rocket to turn left in the direction of lead until the rudders take a neutral position. A rocket flies along the intersecting straight path "p". It is also useful to direct the rocket into a intersecting trajectory and turn the GOS to the target even before launch.

The system may additionally have a zero rudder sensor SN, which shifts the neutral position of the rudder sensor (for example, electrically using a controlled bridge circuit) to the right. In this case, the rocket flies along the leading gentle arc “o” and enters the fuselage slightly in front of the aiming point.

Claims (4)

1. The guidance system of anti-aircraft missiles, containing rudder drives and amplifiers, characterized in that it is equipped with a threshold signal transmission unit, a digital optical camera and a digital infrared camera, a pixel off unit for a digital infrared camera, an electronic key, a delay line, while the optical camera is connected through the threshold signal transmission unit with the infrared camera pixel off unit, and the infrared camera is connected via an electronic key and a delay line and with an infrared camera pixel off unit to block the signal from the optical camera. 2. The system according to claim 1, characterized in that it contains an active or semi-active radar station and an I-YES logical unit, the inputs of which are connected to the radar station and to the infrared camera, and the output to the guidance system. 3. The guidance system of anti-aircraft missiles, containing rudder drives and amplifiers, characterized in that it is equipped with a movable homing head and rudder position sensors, the homing head being capable of deflecting the rudder position by a signal from the rudder position sensor. 4. The system according to claim 3, characterized in that it is equipped with a mechanism or an electric circuit configured to shift the neutral position of the rudder position sensor in the same direction as the deviation of the homing head from the longitudinal axis of the rocket or additional displacement of the homing head in the same side

Images