RU2814065C1 - Two-system guided missile in transport and launch container - Google Patents

Abstract

FIELD: military.

SUBSTANCE: two-system guided missile in a transport and launch container contains an on-board connector, a detachable starting engine, a warhead, a propulsion system with booster and sustainer engines, a propulsion engine launch unit, a gyroscopic roll sensor, an angular velocity sensor, a photoreceiving device, an on-board radio equipment unit, and an on-board power supply, digital processing unit, decoder, microcontroller, data bus and steering compartment connected in a certain way.

EFFECT: ensures an increase in the controlled flight range and a reduction in the near firing zone.

1 cl, 5 dwg

Description

A two-system guided missile in a transport and launch container belongs to the field of guided missile design and can be used in anti-tank missile systems.

There is an anti-tank guided missile "Sturm" and its modernization - "Attack" as part of the self-propelled and helicopter complexes "Sturm-S" and "Sturm-V", placed on the BM 9P149 combat vehicle and on the Mi-24 and Mi-28 / cm helicopters . “9M114 guided projectile. Technical description and operating instructions", Moscow, Voenizdat, 1982; Angelsky R.D. Domestic anti-tank systems. Illustrated reference book. - M.: ACT, Astrel, p. 80, 146/. The onboard control equipment of this missile has one control system, which is built on the principle of a semi-automatic radio command control system. The disadvantage of this missile is its insufficient controlled flight range.

There is the Khrizantema missile system, which provides almost simultaneous firing of two missiles at two targets from one carrier. The guidance method used in this complex / Angelsky R.D. Domestic anti-tank systems. Illustrated reference book. - M.: ACT, Astrel, p. 159, 160/, includes detection and recognition of two targets, automatic tracking of the first target, launch of the first missile controlled by a radio channel, semi-automatic tracking of the second target and launch of the second missile controlled by a laser beam.

Missile system implementing this method / Angelsky R.D. Domestic anti-tank systems. Illustrated reference book. - M.: ACT, Astrel, p. 159, 160/, includes two independent guidance systems via a radio channel and a laser beam, as well as a radar station for detecting and automatically tracking a target.

The disadvantage of this missile is the significant range of launching the missile onto the aiming line and the insufficient range of controlled flight.

The closest invention in terms of the totality of essential features is a guided missile (Russia, patent No. 2518126, September 25, 2012), which has the ability to be guided either by a laser beam or by radio command, and is accepted as a prototype.

The specified missile structurally consists of: a guided missile, a detachable booster engine and a transport and launch container (TPC).

The TPK serves to protect the rocket from mechanical damage during operation (storage, transportation) and is a guiding device during rocket launch.

A guided missile (UR) in a TPK contains a warhead, a steering compartment, a booster engine, a sustainer engine and an equipment compartment. To create the necessary lifting force, four arched wings are used, and the control aerodynamic force is created by deflecting the aerodynamic rudders. To ensure launch from the TPK, the guided missile has no protruding parts; the rudders and wings are folded and open after the missile has taken off from the transport and launch container.

The missile launcher is electrically connected to the TPK via a side connector, which is disconnected at the moment of firing.

For mechanical and electrical docking with the carrier, the TPK has two trunnions. In the front trunnion there is a pyrostop for holding the missile launcher, and in the rear trunnion there is a board for connecting the launch circuits of the rocket with the launch circuits of the carrier (combat vehicle, helicopter, etc.).

The sustainer dual-mode engine accelerates the rocket to supersonic speed, and then passive flight occurs with a decrease in speed. The range of a controlled flight is considered to be the range at which the available overload, depending on the speed, is sufficient to both compensate for gravity and select guidance errors, i.e. at the end of a controlled flight it is 1.2-1.4, while the speed reaches the minimum value Vmin.

Shooting such missiles into the beam with a low initial speed (less than 100 m/s) is characterized by their large deviations in the transition process due to the influence of initial disturbances in the angular velocity of the longitudinal axis of the rocket, wind, acceleration from gravity and other factors, as well as significant duration transition process (the rocket reaches the beam axis).

Thus, the disadvantage of this missile is the significant range of launching the missile onto the aiming line (near firing zone) and the insufficient range of controlled flight.

For a helicopter, the mass of ammunition is essential; the lighter it is, the more ammunition you can have on board. Therefore, when modernizing a missile, the task of increasing the tactical and technical characteristics while maintaining the mass and dimensions of the new product is solved.

The purpose of the proposal is to increase the combat capabilities of a helicopter missile system when firing guided missiles from a fixed launcher (PU) by expanding the target engagement area while maintaining the mass and dimensions of the new missile.

The technical result of the present invention is to increase the range of controlled flight and reduce the short range of fire.

To achieve the specified technical result in the proposed rocket in a transport and launch container containing a detachable starting engine, a propulsion system including booster and sustainer engines, a propulsion engine launch unit, a warhead, a steering compartment, an on-board radio equipment unit, a photoreceiving device, a digital processing unit with a decoder, the input of which is connected to the output of the photodetector, a gyroscopic roll sensor, an on-board power supply, an on-board connector for the electrical connection of the rocket circuits with the carrier, the total thrust impulse of the propulsion system, which includes the booster engine and the end-combustion propulsion engine, is distributed as follows: in the range (50-75)% falls on the accelerating engine and in the range of (25-50)% falls on the end-combustion propulsion engine. Also, an angular velocity sensor has been introduced, the digital processing unit is equipped with a microcontroller, the second contact of the onboard connector is connected to the first input of the microcontroller, the second input of the microcontroller is connected to the first output of the onboard radio equipment unit, the output of the gyroscopic roll sensor is connected to the third input of the microcontroller, the output of the angular velocity sensor connected to the fourth input of the microcontroller, the second output of the on-board radio equipment unit is connected to the fifth input of the microcontroller, the decoder output is connected to the microcontroller via a data bus, the microcontroller output is connected to the steering compartment. In this case, the propulsion engine starting unit contains an inertial contactor with two contacts connected under the action of starting acceleration, the first output of the onboard power supply is connected to the first contact of the inertial contactor, and the second input of the secondary power source is connected to the second output of the onboard power supply, the first input of the secondary the power source is connected to the second contact of the inertial contactor, and the first output of the secondary power source is connected to the first inputs of the time relay, protective relay and switching relay, the second output of the secondary power source is connected to the second input of the time relay, and the output of the time relay is connected to the second input of the switching relay , the output of the switching relay is connected to the second input of the protective relay, the output of the protective relay is connected to the electric igniter of the main engine.

To maintain the dimensions of the TPK and the same mass of the rocket, the fuel mass is reduced compared to the prototype, an accelerator engine is used, and an end combustion engine with optimal thrust is used as a propulsion engine, providing an increased maximum range of the controlled flight of the rocket. The amount of fuel and the total thrust impulse to maintain the total mass and dimensions of the rocket are reduced by 18%. This makes it possible to increase the maximum controlled flight range of the rocket by 40% compared to the prototype.

The main engine is started after the booster engine has finished working and the speed is reduced to 1.2-Vmin (this time is ~5.5 s), while the thrust force of the main engine is selected to maintain the flight speed at the level of 1.2-Vmin, for which a solid fuel rocket engine is used (solid propellant rocket engine) end combustion. During operation of the booster engine, 5% of the maximum range is covered, and during passive flight before the main engine is turned on, another 17% of the maximum range is covered.

To reduce the close firing zone, a DUS is used, which makes it possible to stabilize the longitudinal axis of the missile and thereby reduce the deflection of the missile in the area of impingement into the beam and reduce the duration of the transition process (the missile reaches the beam axis) (“Design of anti-aircraft guided missiles” edited by Golubev I.S. . and Svetlova V.G., M., publishing house "MAI", 2001, pp. 370, 371, 374, 375).

A guided missile has two control systems: a laser beam and a radio beam. The presence of two control systems allows simultaneous firing at two targets simultaneously. In this case, the carrier can be either dual-system, with the ability to point in laser and radio beams, or single-system.

The laser beam control system is a system in which the equipment that creates the information field by scanning space with a laser beam is located on the carrier, and on the UR there is a photoreceiving device that perceives laser radiation. Pulses from the illumination of the photodetector are sent to a decoder, which determines the coordinates of the rocket in the information field. These coordinates are sent to the rocket's onboard control system.

The control system in a radio beam represents a system in which a radar station that creates a radio beam aimed at a target is located at the control point of the missile launcher (Yu.P. Dobrolensky, V.I. Ivanova, G.S. Pospelov, Automation of Guided Projectiles, M., Oborongiz, 1963, pp. 139-148), and on the UR there is a radio receiver (unit of on-board radio equipment) (Fundamentals of radio control, edited by Veitzel V.A. and Tipugin V.N., M., Soviet radio, 1973 g., pp. 283, 294), receiving signals from the radar transmitter of the control point. This block is a measuring device that determines the magnitude and direction of the deviation of the missile launcher from the axis of the equal-signal direction in the coordinate system associated with this direction. From the output of the unit, the control signal enters the on-board control system of the missile launcher. When the steering wheels are turned, a control force is created that returns the missile to the axis of the equal-signal direction of the radio beam. As a result, the missile launcher will move along the radio beam.

The main advantages of beam control systems are their long range and comparative simplicity (less complexity of on-board equipment for creating control signals).

The essence of the claimed invention is illustrated by a drawing, where Fig. 1 shows a block diagram of a two-system guided missile.

In Fig.1 the following notations are used:

1 - onboard connector for connecting the electrical circuits of the rocket with the carrier;

2 - starting motor;

3 - warhead;

4 - accelerating engine;

5 - main engine;

6 - main engine starting unit;

7 - gyroscopic roll sensor (GDS);

8 - angular velocity sensor (ARS);

9 - photodetector device (PDU);

10 - on-board radio equipment unit (OBRA);

11 - digital processing unit;

12 - microcontroller (MK);

13 - decoder;

14 - data bus;

15 - steering compartment;

16 - first input of the MK;

17 - second input of the MK;

18 - third input of the MK;

19 - fourth input of the MK;

20 - fifth input of the MK;

21 - MK output;

22 - first contact of the on-board connector;

23 - second contact of the on-board connector.

The proposed two-system guided missile in the TPK contains an on-board connector 1, a starting engine 2, a warhead 3, an accelerating engine 4, a sustainer engine 5, a propulsion engine launch unit 6, a steering compartment 15, a gyroscopic roll sensor 7, DUS 8, an onboard radio equipment unit 10, photodetector 9, digital processing unit 11 with decoder 13 of the laser-beam system, and the output of the photodetector 9 is connected to the input of the decoder 13 of the laser-beam system. The digital processing unit 11 is equipped with a microcontroller 12. The output of the decoder 13 is connected to the microcontroller 12 via the data bus 14. The first input 16 of the microcontroller 12 is connected to the second contact 23 of on-board connector 1. To match voltage levels, communication can be carried out, for example, through a signal conversion block " Gathering" into digital form. The second input 17 of the microcontroller 12 is connected to the first output of the BBRA 10 (signal “System type”). To coordinate voltage levels, communication can be carried out, for example, through a block for converting the “System Type” signal into digital form. The third input 18 of the microcontroller 12 is connected to the gyroscopic roll sensor 7.

The fourth input 19 of the microcontroller 12 is connected to the DUS 8 via an ADC. The fifth input 20 of the microcontroller 12 is connected to the second output of the BBRA 10.

Output 21 of microcontroller 12 is connected to the steering compartment 15. To match voltage levels, communication can be carried out, for example, through a microcontroller output signal generation circuit.

The key for selecting a control system is the “System Type” signal, supplied from the first output of the BBRA 10 to the second input 17 of the microcontroller 12.

Warhead 3 is located in front of the missile launcher and is an autonomous unit.

The rocket propulsion system includes booster and sustainer engines.

Main engine 5 is a single-chamber end-combustion engine with two lateral inclined bells.

Behind the main engine there is an accelerating 4 single-chamber single-mode solid fuel engine with two lateral inclined bells.

Behind the accelerating engine 4 there is an equipment compartment where the digital processing unit 11, the gyroscopic roll sensor 7, the DUS 8, the power supply batteries (on-board power supply), the photo receiving device 9 of laser radiation, the on-board radio equipment unit 10 and the steering compartment 15 are located.

Starting engine 2 is a single-chamber, single-mode solid fuel engine.

The photodetector device 9 can be made, for example, based on the FD342-03 photodiode.

The on-board radio equipment unit 10 can be made in accordance with the design (Fundamentals of radio control, edited by V.A. Veitsel and V.N. Tipugin, M., Soviet Radio, 1973, pp. 283, 294).

The digital processing unit 11 can be made on the basis of the 1986 BE92U microcontroller (microcontroller 12) and the 5503BTs7U microcircuit (decoder 13).

The block for converting the “Generation” signal into digital form, which converts the voltage at the contact and its absence into digital values (1, 0), can be made on a 2T3129A9 transistor and a 2D707AS9 protective diode (Fig. 2).

The block for converting the “System Type” signal into digital form can be made according to the circuit shown in Fig.2.

To protect input signals from the FPU and GDK, blocks with the same type of circuit shown in Fig. 3 are used. The 2D707AS9 diode provides voltage limitation of the input signal, and the resistor serves to limit the current.

The circuit for generating the output signal 21 of microcontroller 12 can be implemented on a 5559IN14AU microcircuit.

To match analog signals coming from BBRA 10 and DUS 8, ADC amplifiers are used, which can be made on operational amplifiers 544UD16U3 (Fig. 4).

The two-system guided missile in the TPK functions as follows.

The operator on the carrier selects a specific missile launcher and the type of guidance system, aims the sight at the target and presses the “Start” button.

The “Start” signal received from the carrier to the first contact 22 of on-board connector 1 starts the on-board power supply 24, which leads to the initialization of the digital processing unit 11.

When the operator selects a guidance system in the radio beam, BBRA 10 receives clock pulses and generates a “System Type” signal equal to 1; in the absence of clock pulses, the “System Type” signal is equal to 0. The “System Type” signal is sent to the second input 17 of microcontroller 12. Microcontroller 12 V Depending on the value of the signal, “System Type” switches to either receiving signals from the BBRA 10 or receiving signals from the decoder 13. Power is supplied to the warhead 3, which is necessary for its cocking. The launch engine 2 of the rocket is started, which is designed to eject the rocket from the TPK.

When the rocket is separated from the TPK, the “Gathering” circuit breaks, thereby generating the “Getting” signal on the second contact 23 of on-board connector 1. The “Getting” signal is sent to the first input 16 of microcontroller 12 and the time countdown starts, and steering drive control signals are generated on the flight path in accordance with the laid down program.

The booster engine 4 is turned on after the rocket exits the TPK in the same way as the prototype (in the prototype this engine is called the main engine). After turning on the accelerating motor, the starting motor is separated.

At the first stage of guidance, the longitudinal axis of the missile is stabilized according to information from DUS 8. At the second stage of guidance, the missile is controlled by command-beam telecontrol signals.

When controlled from a signal for angular stabilization of the longitudinal axis of the rocket, the disturbing moment acting on the rocket in flight and caused by wind speed, starting disturbances, etc., is compensated by the control moment.

When located in the information field created by the scanning laser beam of the carrier guidance equipment, periodic illumination of the photodetector device 9 occurs. From the output of the photodetector 9, the pulses generated by it arrive at the input of the decoder 13, in which the deviation of the rocket from the axis of the information field is calculated from the time intervals between the pulses. For communication between the microcontroller 12 and the decoder 13, there is a data bus 14, implemented using digital ports, the exchange is implemented at the software level.

If you select guidance in the radio beam, BBRA 10 is synchronized with the transmitting antenna over the air. When located in the radio beam, deviations from the equal-signal direction arriving at the fifth input 20 of the microcontroller 12 are determined in the BBRA 10.

The generation of control signals for the steering compartment 15 is carried out programmatically in the microcontroller 12 using a signal from the gyroscopic roll sensor 7 supplied to the third input 18 of the microcontroller 12. The control signal is transmitted to the steering compartment 15 through output 21 of the microcontroller 12.

The gyroscopic roll sensor 7 is necessary so that the vector of control forces is formed in the required direction in space, at an arbitrary roll angle of the rocket body to which the rudders are attached.

In the steering compartment 15, the rudders are deflected, a control force arises, which leads to a decrease in the misalignment of the missile with the aiming line and the missile moves along the aiming line, which ensures hitting the target. When the target is hit, warhead 3 is triggered and the target is hit.

After a predetermined time after the descent, the starting unit 6 of the main engine 5 supplies power to the electric igniter of the main engine 5, and it begins its work.

The block diagram of the launch unit 6 of the main engine 5 is shown in Fig.5.

In Fig.5

24 - onboard power supply;

25 - inertial contactor;

26 - secondary power source;

27 - time relay;

28 - switching relay;

29 - protective relay;

30 - electric igniter of main engine 5.

The launch unit 6 of the main engine 5 contains an inertial contactor

25, secondary power supply 26, time relay 27, switching relay 28, protective relay 29.

In the initial state, the starting unit 6 of the main engine 5 is disconnected from the on-board power source 24, and the electric igniter 30 of the main engine 5 is bypassed by a protective relay 29, which ensures the safety of the product during assembly, storage and from exposure to electromagnetic fields during operation.

When the starting motor 2 is triggered (reset of the UR), the inertial contactor 25 supplies voltage from the on-board power supply to the secondary power source 26, made on a DC/DC converter of the MDM6-1 V type.

In this case, the secondary power source is activated

26, which supplies voltage to the time relay 27, the switching relay 28 and the protective relay 29.

Time relay 27, made on a relay of the RDV 11 type, provides a delay in the supply of power to the electric igniter of the main engine 5. After the time relay 27 is activated, the switching relay 28 is activated and the protective relay 29 is unlocked.

Switching relay 28 can be made on a 2M419A1 microassembly.

Protective relay 29 can be made on microassembly 249 KP14AR.

The signal from the output of the protective relay 29 is sent to the electric igniter 30 of the main engine 5 and it is activated.

The given design of the starting unit 6 of the main engine 5 ensures its reliable operation at a given point in time.

Claims (1)

A two-system guided missile in a transport and launch container, containing a detachable starting engine, a propulsion system including booster and sustainer engines, a propulsion engine launch unit, a warhead, a steering compartment, an on-board radio equipment unit, a photodetector, a digital processing unit with a decoder, the input of which is connected with the output of a photodetector, a gyroscopic roll sensor, an onboard power supply, an onboard connector for the electrical connection of the rocket circuits with the carrier, characterized in that the total thrust impulse of the propulsion system is distributed as follows: in the range of 50-75% for the booster engine and in the range of 25- 50% for the propulsion engine of the end combustion, and also an angular velocity sensor has been introduced, the digital processing unit is equipped with a microcontroller, the second contact of the onboard connector is connected to the first input of the microcontroller, the second input of the microcontroller is connected to the first output of the onboard radio equipment unit, the output of the gyroscopic roll sensor is connected to the third input microcontroller, the output of the angular velocity sensor is connected to the fourth input of the microcontroller, the second output of the on-board radio equipment unit is connected to the fifth input of the microcontroller, the decoder output is connected to the microcontroller via the data bus, the microcontroller output is connected to the steering compartment, while the propulsion engine starting unit contains an inertial contactor with two contacts connected under the action of starting acceleration, the first output of the on-board power source is connected to the first contact of the inertial contactor, and the second input of the secondary power source is connected to the second output of the on-board power source, the first input of the secondary power source is connected to the second contact of the inertial contactor, and the first output the secondary power supply is connected to the first inputs of the time relay, the protective relay and the switching relay, the second output of the secondary power supply is connected to the second input of the time relay, and the output of the time relay is connected to the second input of the switching relay, the output of the switching relay is connected to the second input of the protective relay, the output The protective relay is connected to the electric igniter of the main engine.

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