RU2192605C2 - Method of guidance of remote-controlled rocket and guidance system for its realization - Google Patents

Abstract

FIELD: rocket engineering, weapon complexes of remote-controlled rockets. SUBSTANCE: method of guidance of remote-controlled rocket includes formation of wide and narrow control fields, launch of rocket at angle to target sight line, rocket boost with the aid of engine in the course of time interval t_b, rocket guidance in wide control field in the course of time interval t_b with formation of control command proportional to angular mismatch between flame of rocket engine and target sight line, rocket guidance in narrow control field with formation of control command proportional to angular mismatch between radiation source on rocket and target sight line. With rocket guidance in wide control field reference signal is formed, present angular velocity of rocket relative to target sight line is measured, signal proportional to angular position of longitudinal axis of plum of flame of rocket engine with respect to its sight line is formed on its basis, generated signal is compared with reference signal. If signal proportional to angular position of longitudinal axis of plume of flame of rocket engine relative to its sight line is less than reference one then correction control command proportional to present angular velocity of rocket with reference to target sight line is formed and rocket is guided with allowance for this correction command. Guidance system of remote-controlled rocket includes target direction finder and rocket control circuit each pitching and course channel of which has rocket direction finder with narrow control field, rocket direction finder with wide control field, controlled switch which second input is connected to first output and which information input is connected to second output of rocket direction finder with narrow control field and former of control command proportional to angular mismatch between rocket and target sight line which second input is connected to corresponding output of target direction finder connected in series. Circuit is inserted with former of reference signal proportional to angular position of longitudinal axis of plume of flame of rocket engine relative to its sight line which input is connected to second output of rocket direction finder with wide control field, with comparator which second input is connected to output of former of reference signal, with controlled key which input is connected to second output of rocket direction finder with wide control field and which information input is connected to output of comparator, with former of correction control command and with adder which second input is connected to output of former of control command proportional to angular mismatch between rocket and target sight line and which output is connected to input of transmitter of control commands. EFFECT: increased noise immunity of optical communication line CARRIER-ROCKET and efficiency of weapon complexes of remote-controlled rockets. 2 cl, 2 dwg

Description

 The invention relates to rocket technology and can be used in weapon systems of remote-controlled missiles.

 Guidance of guided missiles is accompanied by smoke formation of the engines in the accelerating section, which in the case of using the command guidance system ([1], pp. 29-31) with optical sighting of the target and the missile at the first guidance stage associated with the launch of the missile on a kinematic trajectory ([2 ], p. 329), makes it difficult to track the target and reduces the noise immunity of the optical communication line (OLS) of the carrier - rocket.

 There are various methods of guiding missiles in the conditions of smoke generation of their own engines, which make it possible to increase the noise immunity of a carrier-rocket OLS. One of them is shooting at an angle to the line of sight of the target (LCV) with the formation of a control program command on a portion of the flight path with a working engine and transmitting a command through a wireline to send the missile to the LCV and further correct the missile’s flight using remote control when the missile is located on the LCF, implemented, for example, in the anti-tank complex "SWINGFIRE" ([3], p. 509-513, [4], p. 28-29).

 A known guidance system for a telecontrolled missile that implements the above method and consisting of a target finder and a missile control loop, including a program command generation unit in each pitch and course channel, rocket direction finder, a control command generation unit, the second input of which is connected to the corresponding output of the target finder , a managed switch, the second input of which is connected to the output of the block forming the program command, and the information input to the second output engatora missiles transmission device control commands to the missile and rocket.

 The known method of guidance of a remote-controlled missile and guidance system that implements it, due to the use of a wired communication line and program control command on a portion of the flight path with a working engine, which must take into account the variation in the operating time of the rocket engine in the temperature range of its combat use, the departure of parameters during the storage of missiles in conditions of military operation, the speed and maneuverability of the target are limited to use in weapon systems.

Closest to the proposed one is a remote-control missile guidance method implemented in the ADATS anti-aircraft complex ([1], pp. 62-64), which includes the formation of a wide and narrow control field, launching the rocket at an angle to the LCF, and accelerating the rocket using the engine in the passage of time t _p , guidance of the rocket in a wide control field over time t _p with the formation of a control command proportional to the angular mismatch between the torch of the rocket engine and the LEC, guidance of the rocket in a narrow control field with the formation of a control command, p proportional to the angular mismatch between the radiation source on the rocket and the LEC.

 In this method, in the guidance section with the engine running, a command remote control system is implemented. The direction finder of the rocket determines the angular deviation of the rocket relative to the LCF from the torch of the working engine, and based on this mismatch, the ground equipment of the complex forms a command transmitted to the rocket using temporary beam modulation. When a command is transmitted, the energy of the laser beam is concentrated in a narrow beam, which ensures the transmission of information through the engine plume and the reception of radiation by detectors located at the ends of the wings of the rocket.

 After the acceleration engine finishes, a system of teleorienting the rocket in the laser beam is implemented. Two detectors located in the tail of the rocket receive laser radiation. The rocket's on-board equipment converts these signals into rudder control commands that hold the rocket in the center of the beam until they meet the target.

 Closest to the proposed is a remote-control missile guidance system ([1], pp. 62-64), which implements the known guidance method and consists of a target finder and a missile control loop, containing missile direction finder with a narrow control field in each channel pitch and course, connected rocket direction finder with a wide control field, a controlled switch, the second input of which is connected to the first output, and the information input - with the second output of the rocket direction finder with a narrow control field, Control technological command proportional to the angular mismatch between the missile and the target line of sight, a second input coupled to a corresponding output of the direction finder objective transmission device control commands and rocket.

 The known method of pointing a remote-controlled missile and the guidance system that implements it, due to the exclusion of a wired communication line, can increase the range of the target’s destruction, taking into account the presence of a crosswind, speed and maneuverability of the target — to increase the accuracy of pointing the missile at the target. At the same time, the known missile guidance method and the system that implements it have the disadvantage of disrupting missile guidance in the flight section with the engine running due to the possible carrier-rocket overlap in a real-world flight condition, the missile geometrically coinciding with the missile’s line of sight, a smoke plume torch own rocket engine.

A diagram explaining the condition that the carrier-rocket OLS is blocked by the smoke plume of the rocket engine torch is shown in FIG. 1, where the angular size of the smoke plume of the flare of the rocket engine relative to its longitudinal axis is denoted by χ, and the angle between the longitudinal axis of the smoke plume and the line of sight of the rocket is denoted by ζ. From figure 1 it is seen that the absence of overlapping OLS carrier - rocket smoke plume torch own rocket engine occurs if
χ <ζ. (1)
The angle of launch of the rocket relative to the LCF is chosen so that in the process of launching the rocket to the LCF, the angle ζ is higher than the angular size of the smoke plume of the engine χ. Shooting with guided missiles is accompanied by dispersion of trajectories associated with the action of random disturbances on the missile and with errors in pointing it at the target ([1], pp. 137-138), and during the guidance process it may turn out that condition (1) is not fulfilled. In this case, the carrier-rocket OLS will be blocked by the smoke plume of the engine torch, and missile guidance will be disrupted.

 The objective of the invention is to prevent the overlapping of the OLS carrier - rocket with a smoke plume of the torch of the rocket engine and disruption of the guidance of the rocket in a wide control field.

The problem is achieved due to the fact that in the method of pointing a remote-controlled missile, including the formation of a wide and narrow control field, launching the rocket at an angle to the LC, accelerating the rocket with the engine for time t _p , guiding the rocket in a wide control field for time t _p with the formation of a control command proportional to the angular mismatch between the flare of the rocket engine and the LEC, guiding the rocket in a narrow control field with the formation of a control command proportional to the angular mismatch between the radiation source of the rocket and the LEC, when the missile is guided in a wide control field, a reference signal is formed, the current angular velocity of the rocket relative to the LEC is measured, a signal proportional to the angular position of the longitudinal axis of the smoke plume of the rocket engine’s torch relative to its line of sight is formed, the received signal is compared with a reference signal, and in the event that a signal proportional to the angular position of the longitudinal axis of the plume of the plume of the flame of the rocket engine relative to its line of sight is less than porn, form the corrective control command proportional to the current angular velocity of the rocket relative to the LCF, and then the missile is guided by this corrective command.

 The problem is achieved due to the fact that in the guidance system of a telecontrolled missile, consisting of a direction finder and a missile control circuit, containing in each channel of the pitch and course the direction finder of the rocket with a narrow control field, series-connected direction finder of the rocket with a wide control field, a controlled switch, the second the input of which is connected to the first output, and the information input to the second output of the direction finder of the rocket with a narrow control field, and a control command generation unit proportional to the angular a mismatch between the rocket and the LCF, the second input of which is connected to the corresponding output of the target finder, the control command transmission device and the rocket are connected in series, the reference signal generating unit is introduced, the signal generating unit is connected in series, which is proportional to the angular position of the longitudinal axis of the smoke plume of the rocket engine’s flame relative to its line a sight, the input of which is connected to the second output of the direction finder of the rocket with a wide control field, a comparison unit, the second input of which of the second one is connected to the output of the reference signal generating unit, a controlled key is connected in series, the input of which is connected to the second output of the direction finder of the rocket with a wide control field, and the information input is connected to the output of the comparison unit, the corrective control command formation unit and the adder, the second input of which is connected to the output a control command formation unit proportional to the angular mismatch between the rocket and the LCV, and the output is connected to the input of the control command transmission device. In this guidance method and a system for its implementation, the solution to the problem is based on a combination of missile television guidance operations and additional missile control operations on the accelerating section of its movement with the formation of a corrective command proportional to the measured current angular velocity of the missile relative to the LC, which takes into account the angular orientation of the missile and its smoke plume in space. The presence of a corrective command makes it possible to ensure the fulfillment of condition (1) and to prevent overlapping of the carrier-rocket OLS by a smoke plume of the torch of the rocket’s own engine.

 A comparison of the claimed technical solutions with the well-known allowed us to establish compliance with their criterion of "novelty." When studying other well-known technical solutions in the art, the features that distinguish the claimed invention from the prototype were not identified and therefore they provide the claimed technical solutions with the criterion of "inventive step".

где r - дальность до ракеты;
V - скорость ракеты;
φ - угол между ОЛС носитель - ракета и ЛВЦ;

угловая скорость ракеты относительно ЛВЦ;
η - угол разворота вектора скорости ракеты относительно ее линии визирования;
α - угол атаки.In accordance with figure 1, the kinematic equations characterizing the geometry of the motion of the rocket relative to the LCF have the following form ([2], p. 336):

where r is the range to the rocket;
V - rocket speed;
φ is the angle between the OLS carrier - rocket and LCV;

the angular velocity of the rocket relative to the LC;
η is the angle of rotation of the rocket velocity vector relative to its line of sight;
α is the angle of attack.

Таким образом, измеряя текущую угловую скорость ракеты относительно ЛВЦ

при известных скорости ракеты V и дальности до ракеты r (например, программных), в соответствии с (3) формируется сигнал ζ, пропорциональный угловому положению продольной оси дымового шлейфа факела двигателя ракеты относительно ее линии визирования, полученный сигнал ζ сравнивается с опорным сигналом ζ_оп, величина которого выбирается с учетом возможного угла атаки α и обеспечивает для конкретных летно-баллистических характеристик ракеты выполнение условия (1) и если
ζ<ζ_оп, (4)
то формируется корректирующая команда управления, пропорциональная текущей угловой скорости ракеты относительно ЛВЦ ([2], стр. 394)

где F(t) - известная функция, определяемая летно-баллистическими характеристиками ракеты;
K₀ - передаточный коэффициент (выбирается при проектировании из условия обеспечения устойчивости контура управления ракетой).From system (2), the angle ζ at small angles of attack α (which is ensured by the chosen guidance method) is

Thus, by measuring the current angular velocity of the rocket relative to the LCF

for known rocket speed V and range r (for example, software), in accordance with (3), a signal ζ is generated proportional to the angular position of the longitudinal axis of the plume of the plume of the rocket engine’s flame relative to its line of sight, the received signal ζ is compared with the reference signal ζ _op , the value of which is selected taking into account the possible angle of attack α and ensures the fulfillment of condition (1) for specific flight-ballistic characteristics of the rocket and if
ζ <ζ _op , (4)
then a corrective control command is formed proportional to the current angular velocity of the rocket relative to the LCF ([2], p. 394)

where F (t) is a known function determined by the flight-ballistic characteristics of the rocket;
K ₀ - gear ratio (selected during design from the condition of ensuring stability of the missile control loop).

 Next, the guidance of the rocket is carried out taking into account the corrective command, which allows to ensure the fulfillment of condition (1), which eliminates the overlapping of the carrier OLS - the rocket with the smoke plume of the torch of the rocket’s own engine and the failure of the rocket guidance.

 A functional diagram of a guidance system that implements a guidance method for a remote-controlled rocket is shown in FIG. 2.

 The guidance system for a remote-controlled missile consists of a target direction finder 1 and a missile control loop, containing a rocket direction finder with a narrow control field 2 in each channel of the pitch and course, a rocket direction finder with a wide control field 3, a controllable switch 4, the second input of which is connected to the first output and the information input - with the second output of the direction finder of the rocket with a narrow control field 2, and a control command generation unit 5 proportional to the angular mismatch between the rocket and the LCF, the second input for which it is connected to the corresponding output of the target finder 1, the control command transmission device 6 and the rocket 7, the reference signal generating unit 8, the signal generating unit proportional to the angular position of the longitudinal axis of the smoke plume of the rocket engine relative to its line of sight 9, connected in series the input of which is connected to the second output of the direction finder of the rocket with a wide control field 3, the comparison unit 10, the second input of which is connected to the output of the forming unit ia reference signal 8, connected in series with a controlled key 11, the input of which is connected to the second output of the direction finder of the rocket with a wide control field 3, and the information input is the output of the comparison unit 10, the block for generating the corrective control command 12 and the adder 13, the second input of which is connected to the output of the control command generation unit 5, and the output is connected to the input of the control command transmission device 6.

 The constituent elements of the system — the direction finder of target 1, the direction finders of the rocket 2 and 3, the transmission device for command control 6 — are known standard elements of missile guidance systems ([5], p. 335).

 The control command formation unit, proportional to the angular mismatch between the rocket and the LCF 5, is also a known device for the remote-control missile guidance system and can be performed on analogue counting-solving elements ([2], p. 371, 394).

 Elements - the comparison unit 10 and the adder 13 can be performed, for example, based on operational amplifiers ([5], p. 42, 232).

 Managed switch 4 and managed key 11 are implemented, for example, on the basis of electronic keys ([7], p. 378).

 The signal generating unit proportional to the angular position of the longitudinal axis of the plume of the rocket engine’s torch relative to its line of sight 9, and the unit for generating the corrective control command 12 can be performed, for example, in the form of decision circuits [6] based on operational amplifiers that implement, respectively, the relation (3 ) and (5).

 The signals from the output of the block forming the reference signal 8 can be set, for example, by the voltage from the power supply, which is scaled by an operational amplifier.

The missile guidance system works as follows (the operation of one guidance channel is considered, for example, in an elevation plane). The direction finder of target 1 carries out its support and measures the angular coordinate of the target φ _c . After the launch of the rocket 7, first, the direction finder of the rocket with a wide control field 3 (for a time t _p ), and then with a narrow control field 2, it is followed and the angular coordinates of the flight of the rocket are measured, respectively, φ ₁ and φ ₂ . The managed switch 4 transmits the angular coordinate φ of the rocket equal to φ ₁ (direction finding of the rocket in a wide control field) or φ ₂ (direction finding of the rocket in a narrow control field). Switch 4 is switched according to the signal from the direction finder of the rocket with a narrow control field 2, indicating the shooting of the rocket into a narrow control field. The measured angular coordinates of the rocket φ and the target φ _c are received, respectively, at the first and second inputs of the control command formation block 5, where the control command U _{h is formed} , which is proportional to the linear mismatch between the LC and the rocket ([2], p. 369-371).

The generated missile control command U _h from the output of block 5 goes to the first input of the adder 13, where it is summed up with the corrective command U _k , which arrives at its second input from the output of the block for generating the corrective control command 12, and then the control command transmission device 6 is transmitted to the rocket 7 Rocket 7 under the action of the total control command carries out movement relative to the LCV.

которая поступает на вход блока формирования сигнала, пропорционального угловому положению продольной оси дымового шлейфа факела двигателя ракеты относительно ее линии визирования 9, где по зависимости (3) формируется сигнал ζ. Далее этот сигнал сравнивается в блоке сравнения 10 с опорным сигналом ζ_оп, формируемым в блоке формирования опорного сигнала 8, и в случае выполнения условия (4) управляемый ключ 11 открывается, и на блок формирования корректирующей команды управления 12 с соответствующего выхода пеленгатора ракеты с широким полем управления 3 поступает текущая угловая скорость ракеты относительно ЛВЦ

где по зависимости (5) формируется корректирующая команда управления U_к. Далее корректирующая команда управления U_к поступает на соответствующий вход сумматора 13.The corrective command U _to is formed as follows. The missile direction finder with a wide control field 3 also measures the angular velocity of the missile relative to the LCF

which is fed to the input of the signal generating unit, which is proportional to the angular position of the longitudinal axis of the smoke plume of the rocket engine’s torch relative to its line of sight 9, where, according to dependence (3), the signal ζ is generated. Further, this signal is compared in comparison block 10 with the reference signal ζ _op generated in the block for generating the reference signal 8, and if condition (4) is fulfilled, the controlled key 11 is opened, and to the block for generating the corrective control command 12 from the corresponding output of the rocket direction finder with wide control field 3 receives the current angular velocity of the rocket relative to the LC

where according to dependence (5) a corrective control command U _{k is formed} . Next, the corrective control command U _to goes to the corresponding input of the adder 13.

 Thus, the solution of the problem in the proposed method of guidance of a remote-controlled missile and guidance system for its implementation allows you to prevent overlapping OLS carrier - missile smoke plume of the torch’s own rocket engine and disruption of missile guidance in a wide control field.

 The proposed method for guidance of a remote-controlled missile and guidance system for its implementation can improve the noise immunity of the carrier-rocket OLS and the effectiveness of weapon systems of remote-controlled missiles, which distinguishes them from the known ones.

Sources of information
1. F.K. Neupokoev. Shooting anti-aircraft missiles. - M.: Military Publishing House, 1991.

 2. A.A. Lebedev, V.A. Karabanov. The dynamics of control systems for unmanned aerial vehicles. - M.: Mechanical Engineering, 1965.

 3. International Defense Review, 1972, v. 5,015.

 4. Aircraft, 1972, v. 5, N 12.

 5. Ed. V.V. Grigorina-Ryabova. Radar devices. - M.: Soviet Radio, 1970.

 6. I.M. Tetelbaum, Yu.R. Schneider The practice of analog modeling of dynamic systems. - M .: Energoatomizdat, 1987.

 7. S.V. Yakubovsky and others. Analog and digital integrated circuits. - M .: Radio and communications, 1985.

Claims (2)

1. A method of pointing a remote-controlled missile, including the formation of a wide and narrow control field, launching the rocket at an angle to the line of sight of the target, accelerating the rocket using the engine for a time t _p , guiding the rocket in a wide control field for a time t _p with the formation of a control command proportional to the angular mismatch between the flare of the rocket engine and the line of sight of the target, pointing the rocket in a narrow control field with the formation of a control command proportional to the angular mismatch between the source radiation source on the rocket and the line of sight of the target, characterized in that when pointing the rocket in a wide control field, a reference signal is generated proportional to the expected angular size of the plume of the plume of the rocket engine’s flame, the current angular velocity of the rocket relative to the line of sight of the target is measured, and a signal is generated on its basis, proportional to the angular position of the longitudinal axis of the smoke plume of the rocket engine’s torch relative to its line of sight, the received signal is compared with the reference signal if cash proportional to the angular position of the longitudinal axis of the smoke plume of the rocket engine’s flame relative to its line of sight is less than the reference one, a corrective control command is formed proportional to the current angular velocity of the rocket relative to the line of sight of the target and the missile is guided by this corrective command.  2. A guidance system for a telecontrolled missile, consisting of a direction finder and a missile control loop, containing a rocket direction finder in each channel of the pitch and course with a narrow control field, series-connected rocket direction finder with a wide control field, a controllable switch, the second input of which is connected to the first output, and the information input - with the second output of the direction finder of the rocket with a narrow control field, a control command formation unit proportional to the angular mismatch between the rocket and the line of sight I am the target, the second input of which is connected to the corresponding output of the target finder, and a control command transmission device, characterized in that a reference signal generating unit is introduced into it, proportional to the expected angular size of the smoke plume of the rocket engine torch, and a signal generating unit proportional to the angular position is connected in series the longitudinal axis of the plume of the flame of the rocket engine relative to its line of sight, the input of which is connected to the second output of the direction finder of the rocket with a wide a control terminal, a comparison unit, the second input of which is connected to the output of the reference signal generating unit, a controllable key, the input of which is connected to the second output of the direction finder of the rocket with a wide control field, and the information input - with the output of the comparison unit, the unit for generating the corrective control command and an adder, the second input of which is connected to the output of the control command formation unit, which is proportional to the angular mismatch between the missile and the target line of sight, and the output is connected to the input of the control command transmission device.

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