Homing heads of domestic long-range surface-to-surface missiles. Homing systems "Automatic target tracking" mode
FOREIGN MILITARY REVIEW No. 4/2009, pp. 64-68
Colonel R. SHCHERBININ
Currently, the leading countries of the world are conducting R&D aimed at improving the coordinators of optical, optoelectronic and radar homing heads (GOS) and correction devices for control systems of aircraft missiles, bombs and cassettes, as well as autonomous ammunition of various classes and purposes.
Coordinator - a device for measuring the position of the missile relative to the target. Tracking coordinators with gyroscopic or electronic stabilization (homing heads) are generally used to determine the angular velocity of the line of sight of the missile-moving target system, as well as the angle between the longitudinal axis of the missile and the line of sight and a number of other necessary parameters. Fixed coordinators (without moving parts), as a rule, are part of correlation-extreme guidance systems for stationary ground targets or are used as auxiliary channels of combined seekers.
In the course of ongoing research, a search for breakthrough technical and design solutions is carried out, the development of a new elemental and technological base, improvement of software, optimization of weight and size characteristics and cost indicators of on-board equipment of guidance systems.
At the same time, the main directions for improving tracking coordinators have been identified: the creation of thermal imaging seekers operating in several sections of the IR wavelength range, including with optical receivers that do not require deep cooling; practical application of active laser ranging devices; introduction of active-passive radar seekers with a flat or conformal antenna; creation of multi-channel combined seekers.
In the USA and a number of other leading countries, over the past 10 years, for the first time in world practice, thermal imaging coordinators of HTO guidance systems have been widely introduced.
Preparing for a combat mission of the A-10 attack aircraft (URAGM-6SD “Maverick” in the foreground)
American air-to-ground missile AGM-158A (JASSM program)
Promising air-to-ground guided missile AGM-169
IN In the infrared seeker, the optical receiver consisted of one or more sensitive elements, which did not allow obtaining a full signature of the target. Thermal imaging seekers operate at a qualitatively higher level. They use multi-element OPs, which are a matrix of sensitive elements placed in the focal plane of the optical system. To read information from such receivers, a special optical-electronic device is used, which determines the coordinates of the corresponding part of the target image projected onto the OP by the number of the exposed sensitive element, followed by amplification, modulation of the received input signals and their transmission to the computing unit. The most widespread are reading devices with digital image processing and the use of fiber optics.
The main advantages of thermal imaging seekers are a significant field of view in scanning mode, amounting to ± 90° (for infrared seekers with four to eight element OPs, no more than + 75°) and an increased maximum target acquisition range (5-7 and 10-15 km, respectively). In addition, it is possible to work in several sections of the infrared range, as well as implement automatic target recognition and aiming point selection modes, including in adverse weather conditions and at night. The use of a matrix OP reduces the likelihood of simultaneous damage to all sensitive elements by active countermeasure systems.
Thermal imaging coordinator for the Damascus target
Thermal imaging devices with uncooled receivers:
A - fixed coordinator for use in correlation systems
corrections; B - tracking coordinator; B - aerial reconnaissance system camera
Radar seeker With flat phased array antenna
For the first time, the American AGM-65D Maverick medium-range and AGM-158A JASSM long-range air-to-ground missiles are equipped with a fully automatic (not requiring operator correction commands) thermal imaging seeker. Thermal imaging target coordinators are also used as part of the UAB. For example, the GBU-15 UAB uses a semi-automatic thermal imaging guidance system.
In order to significantly reduce the cost of such devices in the interests of their mass use as part of mass-produced JDAM-type UAB, American specialists developed a thermal imaging coordinator for the Damascus target. It is designed to detect, recognize a target and correct the final section of the UAB trajectory. This device, made without a servo drive, is rigidly fixed in the nose of the bombs and uses the standard power source of the aerial bomb. The main elements of the TCC are the optical system, an uncooled array of sensitive elements and an electronic computing unit that ensures image formation and conversion.
The coordinator is activated after the UAB is reset at a distance to the target of about 2 km. Automatic analysis of incoming information is carried out within 1-2 s with an image changing speed of the target area of 30 frames/s. To recognize a target, correlation-extremal algorithms are used to compare the image obtained in the infrared range with photographs of specified objects converted into digital format. They can be obtained during the preliminary preparation of a flight mission from reconnaissance satellites or aircraft, as well as directly using on-board devices.
In the first case, target designation data is entered into the UAB during pre-flight preparation, in the second - from aircraft radars or infrared stations, information from which is sent to the tactical situation indicator in the cockpit. After detecting and identifying the target, the ISU data is corrected. Further control is carried out in the usual mode without using a coordinator. Moreover, the bombing accuracy (BAC) is no worse than 3 m.
Similar research with the aim of developing relatively cheap thermal imaging coordinators with uncooled OPs is being carried out by a number of other leading companies.
Such OPs are planned to be used in seekers, correlation correction systems and aerial reconnaissance. The sensitive elements of the OP matrix are made on the basis of intermetallic (cadmium, mercury and tellurium) and semiconductor (indium antimonide) compounds.
Promising optoelectronic homing systems also include an active laser seeker, developed by Lockheed-Martin to equip promising missile launchers and autonomous ammunition.
For example, as part of the seeker of the experimental autonomous aircraft munition LOCAAS, a laser location station was used, which provides detection and recognition of targets through three-dimensional high-precision surveying of terrain areas and objects located on them. To obtain a three-dimensional image of a target without scanning it, the principle of reflected signal interferometry is used. The LLS design uses a laser pulse generator (wavelength 1.54 microns, pulse repetition rate 10 Hz-2 kHz, duration 10-20 ns), and a charge-coupled device matrix of sensitive elements is used as a receiver. Unlike LLS prototypes, which had a raster scan of the scanning beam, this station has a larger (up to ± 20°) viewing angle, less image distortion and significant peak radiation power. It is interfaced with automatic target recognition equipment based on signatures of up to 50 thousand typical objects stored in the on-board computer.
During the flight of ammunition, the LLS can search for a target in a strip of the earth's surface 750 m wide along the flight path, and in recognition mode this zone will decrease to 100 m. When several targets are detected simultaneously, the image processing algorithm will ensure the possibility of attacking the highest priority of them.
According to American experts, equipping the US Air Force with aviation ammunition with active laser systems, providing automatic detection and recognition of targets with their subsequent high-precision destruction, will be a qualitatively new step in the field of automation and will help increase the effectiveness of air strikes during combat operations in theaters.
Radar seekers of modern missile defense systems are used, as a rule, in guidance systems for medium- and long-range aircraft weapons. Active and semi-active seekers are used in air-to-air missiles and anti-ship missiles, passive seekers are used in anti-ship missiles.
Promising missile launchers, including combined (universal) missiles designed to engage ground and air targets (air-air-ground class), are planned to be equipped with radar seekers with flat or conformal phased array antennas, made using visualization technologies and digital processing of inverse target signatures.
It is believed that the main advantages of seekers with flat and conformal antenna arrays compared to modern coordinators are: more effective adaptive rejection of natural and organized interference; electronic beam control with complete elimination of the use of moving parts with a significant reduction in weight and size characteristics and power consumption; more efficient use of polarimetric mode and Doppler beam narrowing; increase in carrier frequencies (up to 35 GHz) and resolution, aperture and field of view; reducing the influence of the properties of radar conductivity and thermal conductivity of the radome, which cause aberration and signal distortion. In such seekers it is also possible to use modes of adaptive adjustment of an equal-signal zone with automatic stabilization of the characteristics of the radiation pattern.
In addition, one of the directions for improving tracking coordinators is the creation of multi-channel active-passive seekers, for example, thermal imaging-radar or thermal imaging-laser-radar. In their design, to reduce weight, size and cost, the target tracking system (with gyroscopic or electronic stabilization of the coordinator) is planned to be used in only one channel. The remaining seekers will use a fixed emitter and receiver of energy, and to change the angle of sight it is planned to use alternative technical solutions, for example, in the thermal imaging channel - a micromechanical device for precise adjustment of lenses, and in the radar - electronic scanning of the beam of the directional pattern.
Prototypes of combined active-passive seekers:
on the left - radar-thermal imaging gyro-stabilized seeker for
promising air-to-ground and air-to-air missiles; on right -
active radar seeker with phased array antenna and
passive thermal imaging channel
Tests in the wind tunnel of the SMACM missile launcher being developed (in the picture on the right the missile seeker)
It is planned to equip the promising JCM missile launcher with a combined seeker with semi-active laser, thermal imaging and active radar channels. Structurally, the optoelectronic unit of the seeker receivers and the radar antenna are made in a single tracking system, which ensures their separate or joint operation during the guidance process. This seeker implements the principle of combined homing depending on the type of target (heat- or radio-contrast) and environmental conditions, according to which the optimal guidance method is automatically selected in one of the seeker operating modes, and the others are used in parallel to form a contrast image of the target when calculating the point aiming.
When creating guidance equipment for promising missile launchers, Lockheed Martin and Boeing plan to use existing technological and technical solutions obtained during work on the LOCAAS and JCM programs. In particular, as part of the developed SMACM and LCMCM missile launchers, it is proposed to use various versions of the modernized seeker installed on the AGM-169 air-to-ground missile launcher. These missiles are expected to enter service no earlier than 2012.
The onboard guidance system equipment equipped with these seekers must ensure the performance of such tasks as: patrolling in a designated area for an hour; reconnaissance, detection and destruction of designated targets. According to the developers, the main advantages of such seekers are: increased noise immunity, ensuring a high probability of the missile hitting the target, the possibility of use in difficult interference and weather conditions, optimized weight and size characteristics of the guidance equipment, and relatively low cost.
Thus, R&D carried out in foreign countries with the aim of creating highly effective and at the same time inexpensive aviation weapons with a significant increase in the reconnaissance and information capabilities of on-board complexes of both combat and support aviation. will significantly increase the performance of combat use.
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MOSCOW AVIATION INSTITUTE
(STATE TECHNICAL UNIVERSITY)
Guided air-to-surface missile
Compiled by:
Buzinov D.
Vankov K.
Kuzhelev I.
Levin K.
Sichkar M.
Sokolov Ya.
Moscow. 2009
Introduction.
The rocket is made according to a normal aerodynamic design with X-shaped wings and tail. The rocket body is welded and made of aluminum alloys without technological connectors.
The power plant consists of a sustainer turbojet engine and a starting solid fuel accelerator (not available on aircraft-based missiles). The air intake of the main engine is located in the lower part of the body.
The control system is combined, includes an inertial system and an active radar homing head ARGS-35 for the final section, capable of operating in radio countermeasures conditions. To ensure quick detection and acquisition of a target, the seeker antenna has a large rotation angle (45° in both directions). The seeker is covered with a fiberglass radio-transparent fairing.
The missile's penetrating high-explosive fragmentation-incendiary warhead allows it to reliably hit surface vessels with a displacement of up to 5,000 tons.
The missile's combat effectiveness is enhanced by flying at extremely low altitudes (5-10 m depending on wave height), which significantly complicates its interception by ship anti-missile systems, and by the fact that the missile is launched without the carrier entering the air defense zone of the attacked ships.
Specifications.
Rocket modifications:
Rice. 1. 3M24 "Uran" rocket.
3M24 "Uran" - a ship-based and land-based missile, used from missile boats with the Uran-E complex and coastal missile systems "Bal-E"
Rice. 2. ITs-35 rocket.
ITs-35 - target (target simulator). It is distinguished by the absence of warheads and seekers.
Rice. 3. Kh-35V rocket.
X-35B - helicopter. Features a shortened starting accelerator. Used on Ka-27, Ka-28, Ka-32A7 helicopters.
Rice. 4. Kh-35U rocket.
Kh-35U - aviation (aircraft) missile. It is distinguished by the absence of a starting accelerator; it is used from ejection launchers AKU-58, AKU-58M or APU-78 on the MiG-29K and Su-27K
Rice. 5. Kh-35E rocket.
Kh-35E - export.
Rocket glider.
2.1. General information.
The rocket airframe has the following main structural elements: body, wings, rudders and stabilizers. (Fig. 6).
The body serves to house the power plant, equipment and systems that ensure autonomous flight of the missile, targeting and hitting the target. It has a monocoque structure consisting of a load-bearing skin and frames, and is made of separate compartments assembled mainly using flanged connections. When connecting the radio transparent fairing to the body of compartment 1 and the starting engine (compartment 6) with adjacent compartments 5 and 7, wedge connections are used.
Fig.6. General form.
The wing is the main aerodynamic surface of the rocket, creating lift. The wing consists of a fixed part and foldable modules. The folding console is made according to a single-spar design with skin and ribs.
Rudders and stabilizers provide controllability and stability in the longitudinal and lateral movement of the rocket; like the wings, they have folding consoles.
2.2. Housing design
The body of compartment 1 (Fig. 7) is a frame structure consisting of power frames 1,3 and casing 2, connected by welding.
Fig.7. Compartment 1.
1. Front frame; 2. Sheathing; 3. Rear frame
The body of compartment 2 (Fig. 8) is a frame structure; consisting of frames 1,3,5,7 and skin 4. To install the warhead, a hatch is provided, reinforced with brackets 6 and frames 3.5. Hatch with edging 2 is intended for fastening the on-board tear-off connector block. There are brackets inside the compartment for placing equipment and routing harnesses.
Fig.8. Compartment 2
1. Front frame; 2. Edging; 3. Frame; 4. Sheathing;
5. Frame; 6. Bracket; 7. Rear frame
The body of compartment 3 (Fig. 9) is a welded frame structure made of frames 1,3,8,9,13,15,18 and skins 4,11,16. The components of the compartment body are the hardware frame 28, the fuel tank 12 and the air intake device (AUD) 27. Yokes 2.14 are installed on frames 1.3 and 13.15. On the frame 9 there is a rigging unit (bushing) 10.
The landing surfaces and places for attaching the wings are provided on frame 8. There are brackets 25 and 26 to accommodate the equipment. The approach to the electrical equipment and pneumatic system is through hatches closed with covers 5,6,7,17. To attach the fairing to the body, profiles 23 are welded. A pneumatic block is installed on brackets 21 and 22. Bracket 20 and cover 24 are designed to accommodate fuel system units. Ring 19 is necessary to ensure a tight connection of the VZU channel with the main engine.
Fig.9. Compartment 3.
1. Frame; 2. Yoke; 3. Frame; 4. Sheathing; 5. Cover;
6. Cover; 7. Cover; 8. Frame; 9. Frame; 10. Bushing;
11. Sheathing; 12. Fuel tank; 13. Frame; 14. Yoke;
15. Frame;16. Sheathing; 17. Cover; 18. Frame; 19. Ring; 20. Bracket; 21. Bracket;; 22. Bracket; 23. Profile;
24. Cover; 25. Bracket; 26. Bracket; 27. VZU;
28. Hardware part of the compartment
The body of compartment 4 (Fig. 10) is a welded frame structure consisting of frames 1,5,9 and skins 2,6. To install the engine in frames 1 and 5, there are mounting surfaces and holes.
Fig. 10. Compartment 4.
1. Frame; 2. Sheathing; 3. Edging; 4. Cover;
5. Frame; 6. Sheathing; 7. Edging; 8. Cover;
9. Frame; 10. Bracket; 11. Bracket.
For fastening the rudders, frame 5 has landing pads and holes. Brackets 10,11 are designed to accommodate equipment. Access to the equipment installed inside the compartment is provided through hatches with edgings 3.7, closed with covers 4.8.
The body of compartment 5 (Fig. 11) is a welded frame structure made of load-bearing frames 1,3 and casing 2.
To connect the starting engine harness connector, a hatch is provided, reinforced with edging 4, which is closed with a lid 5. To install 4 pneumatic bridges, holes are made in the housing.
Rice. 11. Compartment 5.
1. Frame. 2. Sheathing. 3. Frame. 4. Edging. 5. Cover.
The starting motor is located in the housing of compartment 6 (Fig. 12). The compartment housing is also the engine housing. The body is a welded structure consisting of a cylindrical shell 4, front 3 and rear 5 frames, bottom 2 and neck 1.
Fig. 12. Compartment 6.
1. Neck; 2. Bottom; 3. Front clip; 4. Shell;
5. Rear clip
Compartment 7 (Fig. 13) is a power ring on which there are seats for stabilizers and a yoke. The compartment is closed at the back with a lid. At the bottom of the compartment there is a hole used as a loading unit.
Rice. 13. Compartment 7.
Note. Compartments 5,6 and 7 are available only on missiles used in missile defense systems.
2.3. Wing.
The wing (Fig. 14) consists of a fixed part and a rotating part 3, connected by an axis 2. The fixed part includes a body 5, a front fairing 1 and 6, fixed to the body with screws 4. The body houses a pneumatic mechanism for folding the wing. The rotating part contains a mechanism for locking the wing in the unfolded position.
Folding of the wing is carried out as follows: under the action of air pressure supplied through passage 12, piston 7 with eye 8, using link 10, drives the rotating part. The link is connected to the eye and the rotating part of the wing with pins 9 and 11.
The wings are locked in the unfolded position using pins 14, buried in the conical holes of the bushings 13 under the action of springs 17. The influence of the springs is transmitted through pins 15, with which the pins are secured in the sleeves 16 from falling out.
The wing is spread out by lifting the pins from the holes of the bushings by winding ropes 18 onto the roller 19, the ends of which are secured in the pins. The roller rotates counterclockwise.
The wing is installed on the rocket along surfaces D and E and hole B. Four holes D for screws are used to attach the wing to the rocket.
Fig. 14. Wing
1. Front fairing; 2. Axle; 3. Rotating part; 4. Screw; 5. Body; 6. Rear fairing; 7. Piston; 8. Eyelet;
9. Pin; 10. Link; 11. Pin; 12. Passer; 13. Bushing;
14. Pin; 15. Pin;16. Sleeve;17. Spring;18. Rope;
2.4. Steering wheel.
The steering wheel (Fig. 15) is a mechanism consisting of a blade 4, movably connected to a tail 5, which is installed in the housing 1 on bearings 8. Reinforcement is transferred to the steering wheel through a lever 6 with a hinged bearing 7. The blade is a riveted structure consisting of casing and stiffening elements. The trailing edge of the blade is welded. The blade is riveted to the bracket 11, which is movably connected by an axis 10 to the tail.
The steering wheel is folded out as follows. Under the influence of air pressure supplied to the housing through fitting 2, piston 13 through earring 9 drives the blade, which rotates around axis 10 by 135 degrees and is fixed in the unfolded position by retainer 12, which fits into the conical socket of the shank and is held in this position by a spring.
Fig. 15. Steering wheel.
1. Body; 2. Fitting;3. Stopper; 4. Blade; 5. Shank; 6. Lever; 7. Bearing; 8. Bearing; 9. Earring; 10. Axle; 11. Bracket; 12. Lock; 13. Piston
The steering wheel is folded as follows: through hole B, the latch is removed from the conical hole using a special key and the steering wheel is folded. In the folded position, the steering wheel is held using a spring-loaded stopper 3.
To install the rudder on the rocket, the body has four holes B for bolts, a hole D and a groove D for pins, and there are also seats with threaded holes E for attaching fairings.
2.5. Stabilizer.
The stabilizer (Fig. 16) consists of platform 1, base 11 and console 6. The base has a hole for the axis around which the stabilizer rotates. The console is a riveted structure, consisting of a skin 10, a stringer 8 and an end 9. The console is connected to the base through a pin 5.
Fig. 16. Stabilizer.
1. Platform; 2. Axle; 3. Earring; 4. Spring; 5. Pin; 6. Console;
7. Loop; 8. Stringer; 9. Ending; 10. Sheathing; 11. Base
The stabilizers are hinged on the rocket and can be in two positions - folded and unfolded.
In the folded position, the stabilizers are located along the rocket body and are held by the hinges by 7 pneumatic stop rods installed on compartment 5. To bring the stabilizers from the folded position to the open position, spring 4 is used, which is connected at one end to the earring 3, hingedly mounted on the platform, and at the other to a pin 5.
When compressed air is supplied from the pneumatic system, the pneumatic stops release each stabilizer, and under the action of an extended spring it is installed in the open position.
Power point
3.1. Compound.
The rocket uses two engines as a power plant: a solid fuel starting engine (SD) and a sustaining turbojet bypass engine (MD).
SD - compartment 6 of the rocket, ensures launch and acceleration of the rocket to cruise flight speed. At the end of the work, the SD along with compartments 5 and 7 are fired.
The MD is located in compartment 4 and serves to ensure the autonomous flight of the rocket and to provide its systems with power supply and compressed air. The power plant also includes an air intake device and a fuel system.
The VSU is a tunnel type, semi-recessed with flat walls, located in compartment 3. The VSU is designed to organize the air flow entering the MD.
3.2. Starting motor.
The launch engine is designed to launch and accelerate a rocket at the initial level of the flight trajectory and is a single-mode solid fuel rocket engine.
Technical data
Length, mm_______________________________________________550
Diameter, mm________________________________________________420
Weight, kg______________________________________________________________103
Fuel mass, kg___________________________________________69±2
Maximum permissible pressure in the combustion chamber, MPa________11.5
Gas outflow velocity at the nozzle exit, m/s______________________2400
Temperature of gases at the nozzle exit, K______________________________2180
The SD consists of a body with a charge of solid rocket propellant (SRP) 15, a cover 4, a nozzle block, an igniter 1, and a squib 3.
Docking of the LED with adjacent compartments is carried out using wedges, for which purpose the holders have surfaces with annular grooves. For correct installation of the LED, longitudinal grooves are provided on the holders. On the inner surface of the rear cage there is an annular groove for keys 21 for fastening the nozzle block. The dowels are inserted through the windows, which are then covered with crackers 29 and linings 30, fastened with screws 31.
A nut 9 is screwed onto the neck 8; its correct installation is ensured by pin 7 pressed into the neck.
A heat-protective coating 11 and 17 is applied to the inner side of the housing surface, to which cuffs 13 and 18 are attached, which reduce the voltage in the TRT charge when its temperature changes.
Fig. 17. Starting motor.
1. Igniter; 2. Plug; 3. Squib; 4. Cover;
5. Heat-protective insert; 6. O-ring; 7. Pin;
8. Neck; 9. Nut; 10. Bottom; 11. Heat-protective coating;
12. Film; 13. Front cuff; 14. Front clip; 15. TRT charge; 16. Shell; 17. Heat-protective coating; 18. Back cuff; 19. Rear clip; 20. O-ring; 21. Key; 22. Cover; 23. Heat-protective disk; 24. Clip; 25. O-ring; 26. Bell; 27. Liner; 28. Membrane;
29. Rusk; 30. Overlay; 31. Screw.
The TRT charge is a monoblock firmly attached to the cuffs, made by pouring fuel mass into the housing. The charge has an internal channel of three different diameters, which ensures, when fuel burns through the channel and the rear open end, an approximately constant combustion surface and, therefore, an almost constant thrust. A film 12 separating them is laid between the front cuff and the heat-protective coating.
Cover 4 has: a thread for attaching the igniter, a threaded hole for the squib, a threaded hole for installing a pressure sensor in the combustion chamber when testing, an annular groove for sealing ring 6, a longitudinal groove for pin 7. During operation, the hole for the pressure sensor is closed. plug 2. A heat-protective insert 5 is fixed to the inner surface of the cover. The nozzle block consists of a cover 22, a cage 24, a bell 26, a liner 27 and a membrane 28.
On the outer cylindrical surface of the cover there are annular grooves for the sealing ring 20 and keys 21, on the inner cylindrical surface there is a thread for connection with the holder 24. A heat-protective disk 23 is attached to the front of the cover. The holder 24 has a thread and an annular groove for the sealing ring 25.
The LED starts to work when a DC voltage of 27 V is supplied to the squib. The squib fires and ignites the igniter. The igniter flame ignites the TRT charge. When the charge burns, gases are formed that break through the diaphragm and, leaving the nozzle at high speed, create a reactive force. Under the influence of the SD thrust, the rocket accelerates to the speed at which the MD begins to operate.
3.3. Main engine
A turbojet bypass engine is a short-life, disposable engine designed to create jet thrust during autonomous flight of a rocket and to provide its systems with power and compressed air.
Technical data.
Startup time, s, no more than:
At heights of 50m________________________________________________6
3500m_______________________________________________8
The MD double-circuit turbojet engine includes a compressor, combustion chamber, turbine, nozzle, fairy tale and venting system, starting system, fuel supply and regulation, and electrical equipment.
The first circuit (high pressure) is formed by the flow part of the compressor, the flame tube of the combustion chamber and the flow part of the turbine up to the cut off of the nozzle body.
The second circuit (low pressure) is limited on the outside by the middle housing and the outer wall of the MD, and on the inside by the flow separator, the combustion chamber body and the nozzle body.
The mixing of the air flows of the first and second circuits occurs behind the cut of the nozzle body.
Fig. 18. Main engine.
1. Oil tank; 2. Fan housing; 3. Fan;
4. 2nd stage straightening apparatus; 5. Turbogenerator;
6. 2nd circuit; 7. Compressor; 8. 1st circuit; 9. Pyro candle; 10. Combustion chamber; 11. Turbine; 12. Nozzle; 13. Gas generator.
The MD is secured to the rocket using a suspension bracket through threaded holes in the front and rear suspension struts. The suspension bracket is a power element on which the units and sensors of MD and communications connecting them are placed. In the front part of the bracket there are holes for attaching it to the MD and lugs for attaching the MD to the rocket.
On the outer wall of the MD there are two hatches for installing spark plugs and an air bleed flange for the steering actuators. On the housing there is an air bleeder for pressurizing the fuel tank.
3.3.1. Compressor.
The MD is equipped with a single-shaft, axial eight-stage compressor 7, consisting of a two-stage fan, a middle housing with a device for dividing the air flow into the first and second circuits, and a six-stage high-pressure compressor.
In fan 3, preliminary compression of the air entering the MD is carried out, and in the high-pressure compressor, only the air flow of the primary circuit is compressed to the calculated value.
The fan rotor is of a drum-disk design. The disks of the first and second stages are connected by a spacer and radial pins. The fan rotor and fairing are secured to the shaft with a bolt and nuts. Torque from the shaft to the fan rotor is transmitted using a splined connection. The working blades of the first and second stages are installed in dovetail grooves. The blades are secured against axial movements by a fairing, a spacer and a locking ring. The fan shaft has a gear that drives the gearbox of the pump unit. The oil cavity of the compressor is vented through the cavities of the MD transmission shafts.
Fan housing 2 is welded with cantilever blades of the first stage straightening apparatus soldered into it. The straightening apparatus of the second stage is made as a separate unit and consists of two rings, into the grooves of which the blades are soldered.
Oil tank 1 is located in the front upper part of the housing. The fan housing together with the oil tank is secured to the flange of the middle housing with studs.
The middle body is the main power element of the MD. In the middle case, the air flow leaving the fan is divided into contours.
Attached to the middle body:
MD suspension bracket to the rocket
Pump block
Middle support (ball bearing) cover
Turbogenerator stator
Combustion chamber housing.
A fuel-oil heat exchanger, an oil filter, a pump-out valve and a P-102 sensor for measuring the air temperature behind the fan are installed on the outer wall of the middle housing. The walls of the housing are connected by four power racks, inside of which there are channels for placing fuel, oil and electrical communications.
The middle housing houses a high-pressure compressor housing with 3-7 stage straighteners. The high-pressure compressor housing has holes for unregulated air bypass from the first to the second circuit, which increases the reserves of gas-dynamic stability at low and medium speeds of rotation of the MD rotor.
The rotor of the high-pressure compressor is of a drum-disc design, double-porous. The high-pressure compressor rotor has splined connections with the fan shaft and turbine shaft. The working blades are installed in the annular T-shaped grooves of the rotor disks.
3.3.2. The combustion chamber.
In the combustion chamber, the chemical energy of the fuel is converted into thermal energy and the temperature of the gas flow increases. The MD is equipped with an annular combustion chamber 10, which consists of the following main components:
Flame tube
Main fuel manifold
Additional fuel manifold
Two fire candles with electric igniters
Pyro candles.
The combustion chamber housing is of brazed-welded design. Two rows of straightening blades of the eighth stage of the compressor are soldered into its front part. In addition, the oil system switches are soldered to the body. On the outer wall of the housing there are fourteen flanges for fastening the injectors of the main manifold, flanges for two spark plugs, a fitting for measuring air pressure behind the compressor, and a flange for attaching the adapter to the spark plug.
The flame tube is a circular welded structure. Fourteen cast “snail” swirlers are welded on the front wall. The main fuel manifold is made of two halves. Each has eight nozzles.
To improve the quality of the mixture and increase the reliability of starting the MD, especially at negative ambient temperatures, an additional fuel manifold with fourteen centrifugal nozzles is installed in the flame tube.
3.3.3. Turbine
The turbine is designed to convert the thermal energy of the gas flow of the primary circuit into mechanical energy of rotation and drive of the compressor and units installed on the MD.
Axial two-stage turbine 11 consists of:
First stage nozzle apparatus
Second stage nozzle apparatus
The turbine rotor consists of two wheels (first and second stages), a connecting inter-disk spacer, a starting turbine wheel and a turbine shaft.
The wheels of the stages and the starting turbine are cast together with the rims of the working blades. The first stage nozzle apparatus has 38 hollow blades and is attached to the combustion chamber body. The second stage nozzle apparatus has 36 blades. The first stage wheel is cooled by air taken from the combustion chamber housing. The internal cavity of the turbine rotor and its second stage are cooled by air taken from the fifth stage of the compressor.
The turbine rotor is supported by a roller bearing without an internal race. The outer race has holes to reduce oil pressure under the rollers.
3.3.4. Nozzle.
In the jet nozzle 12, the air flows of the first and second circuits are mixed. On the inner ring of the nozzle body there are 24 blades for spinning up the flow of gases coming out of the starting turbine during startup, and four bosses with pins for fastening the gas generator 13. The tapering nozzle is formed by the profile of the outer wall of the MD and the surface of the gas generator body.
3.3.5. Starting system.
The starting, fuel supply and control system spins up the rotor, supplies metered fuel at start-up, “counter-start” and in the “maximum” mode during start-up, oxygen is supplied to the combustion chamber from an oxygen accumulator through spark plugs.
The system consists of the following main components:
Solid fuel gas generator
Pyro candles with electric igniters
Oxygen battery
Low pressure fuel system
High pressure fuel system
Integrated engine regulator (IEC)
The oxygen battery is a 115 cc cylinder. The mass of oxygen being filled is 9.3 - 10.1 g.
A disposable solid fuel gas generator (SFG) is designed to spin up the MD rotor when it is started. The gas turbine engine consists of an uncharged gas generator and equipment elements: a solid fuel charge 7, an igniter 9 and an electric igniter (EI)
An unloaded gas generator consists of a cylindrical body 10 that turns into a truncated cone, a cover 4 and fasteners.
The housing is provided with a threaded hole for installing a fitting for measuring pressure in the combustion chamber of the gas turbine engine during testing. During operation, the hole is closed with a plug 11 and a gasket 12. On the outside of the body there is an annular groove for the sealing ring 5.
The cover has eight supersonic nozzles 1, which are located tangentially to the longitudinal axis of the gas turbine engine. The nozzles are closed with glued-in plugs, ensuring the tightness of the gas turbine generator and the initial pressure in the combustion chamber of the gas turbine generator necessary for ignition of the solid fuel charge. The cover is connected to the body using a nut 6. The internal cavity of the body is the combustion chamber of the solid fuel charge and igniter placed in it.
Fig. 19. Solid fuel gas generator.
1. Nozzle; 2. Gasket; 3. Electric igniter; 4. Cover;
5. O-ring; 6. Nut; 7. CT charge; 8. Nut;
9. Igniter; 10. Body; 11. Plug; 12. Gasket.
The igniter is installed in a nut 8 screwed into the bottom of the housing. The solid fuel charge is placed in the combustion chamber between the seal and the stop, which protects it from mechanical damage when triggered.
The gas turbine generator is triggered when an electrical impulse is applied to the contacts of the electric igniter. The electric current heats up the filaments of the electric igniter bridges and ignites the igniter compositions. The force of the flame pierces the igniter case and ignites the black powder placed in it. The flame from the igniter ignites the solid fuel charge. The combustion products of the charge and igniter destroy the nozzle plugs and flow out of the combustion chamber through the nozzle holes. Combustion products, falling on the blades of the MD rotor, spin it.
3.3.6. Electrical equipment.
The electrical equipment is designed to control the launch of the MD and power the rocket units with direct current during its autonomous flight.
The electrical equipment includes a turbogenerator, sensors and automation units, starting units, a thermocouple collector and electrical communications. The sensors and units automatically include air temperature sensors behind the fan, an air pressure sensor behind the compressor and a metering needle position sensor installed in the fuel dispenser, a solenoid of the dispenser control valve, and a stop valve solenoid.
Launching units include devices that provide preparation for the launch and launch of the MD, as well as “counter” launch of the MD when it stalls or surges.
Active radar homing head ARGS
4.1. Purpose
The active radar homing head (ARGS) is designed to accurately guide the Kh-35 missile to a surface target in the final part of the trajectory.
To ensure the solution of this problem, the ARGS is activated upon command from the inertial control system (ICS) when the missile reaches the final section of the trajectory, detects incoming targets, selects a target to be hit, determines the position of this target in azimuth and elevation, angular velocities of the line of sight (LOS) ) targets in azimuth and elevation, range to the target and speed of approach to the target and displays these values in the IMS. Based on signals coming from the ARGS, the ISU guides the missile to the target at the final part of the trajectory.
The target can be a reflector target (CR) or an active interference source target (CIAP).
ARGS can be used both for single and salvo missile launches. The maximum number of missiles in a salvo is 100.
ARGS ensures operation at ambient temperatures from minus 50˚С to 50˚С, in the presence of precipitation and with sea waves up to 5-6 points and at any time of the day.
The ARGS provides data to the ISU for guiding the missile to the target when the range to the target decreases to 150 m;
ARGS ensures missile guidance to the target when exposed to active and passive interference created from target ships, naval and air cover forces.
4.2. Compound.
ARGS is located in compartment 1 of the rocket.
According to functional characteristics, ARGS can be divided into:
Transceiver/receiver device (RTD);
Computing complex (VC);
Secondary power supply unit (SPS).
The PPU includes:
Antenna;
Power amplifier (PA);
Intermediate frequency amplifier (IFA);
Signal conditioner (FS);
Reference and reference oscillator modules;
Phase shifters (FV1 and FV2);
Microwave modules.
The VK includes:
Digital computing device (DCU);
Synchronizer;
Information processing unit (IPU);
Control unit;
Converter SKT-code.
4.3. Operating principle.
Depending on the designated operating mode, the PPU generates and emits four types of microwave radio pulses into space:
a) pulses with linear frequency modulation (chirp) and average frequency f0;
b) pulses with highly stable frequency and phase (coherent) microwave oscillations;
c) pulses consisting of a coherent probing part and a distracting part, in which the oscillation frequency of microwave radiation changes according to a random or linear law from pulse to pulse;
d) pulses consisting of a probing part, in which the frequency of microwave oscillations changes according to a random or linear law from pulse to pulse, and a coherent distracting part.
The phase of coherent oscillations of microwave radiation, when the corresponding command is turned on, can change according to a random law from pulse to pulse.
The PPU generates probing pulses and carries out the conversion and preliminary amplification of the reflected pulses. ARGS can generate probing pulses at a technological frequency (peacetime frequency - fmv) or at combat frequencies (flit).
To exclude the possibility of generating pulses at combat frequencies during testing, experimental and training work, the ARGS is equipped with a “MODE B” toggle switch.
When the “MODE B” toggle switch is set to ON, probing pulses are generated only at frequency flit, and when the toggle switch is set to OFF, only at frequency fmv.
In addition to probing pulses, the PPU generates a special pilot signal, which is used to adjust the PPU receiving signal and organize built-in control.
The VK converts into digital form and processes radar information (RL) according to algorithms corresponding to the modes and tasks of the ARGS. The main functions of information processing are distributed between the control unit and the digital control unit.
The synchronizer generates synchronizing signals and commands to control the blocks and nodes of the control panel and issues service signals to the CU that ensure recording of information.
BOI is a high-speed computing device that processes radar images in accordance with the modes listed in Table. 4.1, under the control of the Central Control Unit.
Fighting carries out:
Analog-to-digital conversion of radar images coming from the PPU;
Processing of digital radar images;
Issuance of processing results to the digital control unit and reception of control information from the central digital control unit;
PPU synchronization.
The central digital control unit is intended for secondary processing of radar images and control of ARGS blocks and nodes in all operating modes of ARGS. TsVU solves the following tasks:
Execution of algorithms for switching on operating and control modes of ARGS;
Reception of initial and current information from the ISU and processing of received information;
Reception of information from the control unit, its processing, as well as transmission of control information to the control unit;
Formation of calculated angles for antenna control;
Solving AGC problems;
Formation and transmission of the necessary information to the IMS and automated testing equipment (ATE).
The control unit and the SKT-code converter provide the generation of control signals for the antenna drive motors and the reception of angular channel information from the central digital device and transmission to the digital digital device. From the central control unit the control unit receives:
Calculated antenna position angles in azimuth and elevation (11-bit binary code);
Synchronization signals and control commands.
From the SKT-code converter, the control unit receives the values of the antenna position angles in azimuth and elevation (11-bit binary code).
VIPs are designed to power the ARGS units and units and convert the 27 V BS voltage into constant voltages
4.4. External Relations.
The ARGS is connected to the rocket's electrical circuit by two connectors U1 and U2.
Through connector U1, the ARGS receives power supply voltages of 27 V BS and 36 V 400 Hz.
Through connector U2, control commands are supplied to the ARGS in the form of a voltage of 27 V and digital information is exchanged using a bipolar serial code.
Connector U3 is intended for control. Through it, the “Control” command is sent to the ARGS, and from the ARGS an integrated analog signal “Service” is issued, information about the operability of the ARGS units and devices in the form of a bipolar serial code and the voltage of the ARGS secondary power source.
4.5. Power supply
To power the ARGS, the following is supplied from the rocket's electrical circuit:
Constant voltage BS 27 ± 2.7
Three-phase alternating voltage 36 ± 3.6 V with a frequency of 400 ± 20 Hz.
Consumption currents from the power supply system:
For a 27 V circuit - no more than 24.5 A;
For a 36 V 400 Hz circuit - no more than 0.6 A for each phase.
4.6. Design.
The monoblock is made of a cast magnesium case on which blocks and components are installed, and a cover that is attached to the rear wall of the case. On the cover there are connectors U1 – U3, a technological connector “CONTROL”, not used in operation, a toggle switch “MODE B” is fixed in a certain position with a protective cap (bushing). There is an antenna in the front of the monoblock. The elements of the high-frequency path and their control devices are located directly on the waveguide-slot array of the antenna. The body of compartment 1 is made in the form of a welded titanium structure with frames.
The cone is made of ceramic radiotransparent fiberglass and ends with a titanium ring, which secures the cone to the body of compartment 1 using a wedge connection.
Rubber gaskets are installed around the perimeter of the lid and cone to ensure sealing of the ARGS.
After final adjustment at the factory, before installing the monoblock into the housing, all external metal parts that do not have a paint coating are degreased and coated with lubricant.
Automatic devices installed on carriers of combat charges (CBC) - missiles, torpedoes, bombs, etc. to ensure a direct hit on the target of attack or approach at a distance less than the radius of destruction of the charges. Homing heads perceive the energy emitted or reflected by the target, determine the position and nature of the target’s movement and generate appropriate signals to control the movement of the NBZ. According to the principle of operation, homing heads are divided into passive (perceive energy emitted by the target), semi-active (perceive energy reflected from the target, the source is located outside the homing head) and active (perceive energy reflected from the target, the source is located in the head itself) homing); by type of perceived energy - radar, optical (infrared or thermal, laser, television), acoustic, etc.; according to the nature of the perceived energy signal - pulsed, continuous, quasi-continuous, etc.
The main components of the homing heads are coordinator and electronic computing device. The coordinator provides search, acquisition and tracking of the target by angular coordinates, range, speed and spectral characteristics of the perceived energy. The electronic computing device processes the information received from the coordinator and generates control signals for the coordinator and the movement of the NBZ, depending on the adopted guidance method. This ensures automatic tracking of the target and guidance of the NBZ on it. Receivers of energy emitted by the target (photoresistors, television tubes, horn antennas, etc.) are installed in the coordinators of passive homing heads; Target selection, as a rule, is made according to angular coordinates and the spectrum of energy emitted by it. A receiver of energy reflected from the target is installed in the coordinators of semi-active homing heads; Target selection can be made based on angular coordinates, range, speed and characteristics of the received signal, which increases the information content and noise immunity of homing heads. An energy transmitter and its receiver are installed in the coordinators of active homing heads; target selection can be carried out similarly to the previous case; active homing heads are fully autonomous automatic devices. Passive homing heads are considered the simplest in design, while active ones are the most complex. To increase information content and noise immunity, there can be combined homing heads, in which various combinations of operating principles, types of perceived energy, methods of modulation and signal processing are used. An indicator of the noise immunity of homing heads is the probability of capturing and tracking a target in conditions of interference.
Lit.: Lazarev L.P. Infrared and light devices for homing and guidance of aircraft. Ed. 2nd. M., 1970; Design of missile and barrel systems. M., 1974.
VC. Baklitsky.
The OGS is designed to capture and automatically track a target by its thermal radiation, measure the angular velocity of the missile-target line of sight and generate a control signal proportional to the angular velocity of the line of sight, including under the influence of a false thermal target (FTC).
Structurally, the OGS consists of a coordinator 2 (Fig. 63) and an electronic unit 3. An additional element that makes up the OGS is the body 4. The aerodynamic nozzle 1 serves to reduce the aerodynamic drag of the rocket in flight.
The OGS uses a cooled photodetector, to ensure the required sensitivity of which a cooling system 5 is used. The coolant is liquefied gas obtained in the cooling system from nitrogen gas by throttling.
The block diagram of the optical homing head (Fig. 28) consists of tracking coordinator and autopilot circuits.
The tracking coordinator (SC) carries out continuous automatic tracking of the target, generates a correction signal to align the optical axis of the coordinator with the line of sight and provides a control signal proportional to the angular velocity of the line of sight to the autopilot (AP).
The tracking coordinator consists of a coordinator, an electronic unit, a gyroscope correction system and a gyroscope.
The coordinator consists of a lens, two photodetectors (FPok and FPvk) and two preamplifiers of electrical signals (PUok and PUvk). In the focal planes of the main and auxiliary spectral ranges of the coordinator lens there are, respectively, photodetectors FPok and FPvk with rasters of a certain configuration radially located relative to the optical axis.
The lens, photodetectors, and preamplifiers are mounted on the gyroscope rotor and rotate with it, and the optical axis of the lens coincides with the axis of the gyroscope rotor’s own rotation. The gyroscope rotor, the main mass of which is a permanent magnet, is installed in a gimbal suspension, allowing it to deviate from the longitudinal axis of the GGS by a bearing angle in any direction relative to two mutually perpendicular axes. When the gyroscope rotor rotates, the space within the field of view of the lens is viewed in both spectral ranges using photoresistors.
Images of a remote radiation source are located in the focal planes of both spectra of the optical system in the form of scattering spots. If the direction towards the target coincides with the optical axis of the lens, the image is focused to the center of the OGS field of view. When an angular mismatch appears between the lens axis and the direction towards the target, the scattering spot shifts. When the gyroscope rotor rotates, the photoresistors are illuminated while the scattering spot passes over the photosensitive layer. Such pulsed illumination is converted by photoresistors into electrical pulses, the duration of which depends on the magnitude of the angular mismatch, and with increasing mismatch for the selected raster shape, their duration decreases. The pulse repetition rate is equal to the rotation frequency of the photoresistor.
Rice. 28. Block diagram of the optical homing head
The signals from the outputs of the photodetectors FPok and FPvk are respectively supplied to the preamplifiers PUok and PUvk, which are connected by a common automatic gain control system AGC1, operating on a signal from the PUok. This ensures the constancy of the ratio of values and the preservation of the shape of the output signals of the pre-amplifiers in the required range of changes in the power of the received OGS radiation. The signal from the PUok is supplied to a switching circuit (SC), designed to protect against LTC and background interference. Protection against LTC is based on different temperatures of radiation from the real target and LTC, which determine the difference in the position of the maxima of their spectral characteristics.
The SP also receives a signal from the PUVK, containing information about interference. The ratio of the amount of radiation from the target received by the auxiliary channel to the amount of radiation from the target received by the main channel will be less than unity, and the signal from the LTC does not pass to the output of the SP.
In the SP, a throughput strobe is formed for the target; The target signal allocated to the SP is sent to a selective amplifier and amplitude detector. The amplitude detector (AD) produces a signal whose amplitude of the first harmonic depends on the angular mismatch between the optical axis of the lens and the direction to the target. Next, the signal passes through a phase shifter, which compensates for the signal delay in the electronic unit, and enters the input of a correction amplifier, which amplifies the signal in power, which is necessary to correct the gyroscope and feed the signal to the AP. The load of the correction amplifier (CA) is the correction windings and active resistances connected in series with them, the signals from which enter the AM.
The electromagnetic field induced in the correction coils interacts with the magnetic field of the gyroscope rotor magnet, forcing it to precess in the direction of decreasing the mismatch between the optical axis of the lens and the direction towards the target. Thus, the OGS tracks the target.
At small distances to the target, the size of radiation from the target perceived by the OGS increases, which leads to a change in the characteristics of pulse signals from the output of photodetectors, which deteriorates the ability of the OGS to track the target. To eliminate this phenomenon, the electronic unit of the SC is equipped with a near-zone circuit that provides monitoring of the energy center of the jet stream and nozzle.
The autopilot performs the following functions:
Filtering the signal from the SC to improve the quality of the missile control signal;
Generating a signal to turn the missile at the initial part of the trajectory to automatically ensure the required elevation and lead angles;
Converting the correction signal into a control signal at the missile control frequency;
Formation of a control command on a steering drive operating in relay mode.
The input signals of the autopilot are the signals of the correction amplifier, the near-field circuit and the bearing winding, and the output signal is a signal from a push-pull power amplifier, the load of which is the windings of the electromagnets of the steering valve spool valve.
The correction amplifier signal passes through a series-connected synchronous filter and dynamic limiter and enters the input of the adder ∑І. The signal from the bearing winding is supplied to the FSUR circuit according to the bearing. It is necessary at the initial section of the trajectory to reduce the time it takes to reach the guidance method and set the guidance plane. The output signal from the FSUR goes to the adder ∑І.
The signal from the output of the adder ∑І, the frequency of which is equal to the rotation frequency of the gyroscope rotor, is supplied to the phase detector. The reference signal of the phase detonator is the signal from the GON winding. The GON winding is installed in the OGS in such a way that its longitudinal axis lies in a plane perpendicular to the longitudinal axis of the OGS. The frequency of the signal induced in the GON winding is equal to the sum of the rotation frequencies of the gyroscope and the rocket. Therefore, one of the components of the output signal of the phase detector is a signal at the rocket rotation frequency.
The output signal of the phase detector is fed to a filter, at the input of which it is summed with the signal of the linearization generator in the adder ∑II. The filter suppresses high-frequency components of the signal from the phase detector and reduces nonlinear distortions of the linearization generator signal. The output signal from the filter is fed to a limiting amplifier with a high gain, the second input of which receives a signal from the rocket angular velocity sensor. From the limiting amplifier, the signal goes to the power amplifier, the load of which is the windings of the electromagnets of the steering valve spool valve.
The gyroscope locking system is designed to coordinate the optical axis of the coordinator with the sighting axis of the sighting device, which makes a given angle with the longitudinal axis of the missile. In this regard, when aiming, the target will be in the field of view of the OGS.
The sensor for deviation of the gyroscope axis from the longitudinal axis of the rocket is a bearing winding, the longitudinal axis of which coincides with the longitudinal axis of the rocket. If the gyroscope axis deviates from the longitudinal axis of the bearing winding, the amplitude and phase of the EMF induced in it unambiguously characterize the magnitude and direction of the mismatch angle. Opposite to the bearing winding, the tilt winding is connected, located in the launch tube sensor block. The EMF induced in the tilt winding is proportional in magnitude to the angle between the sighting axis of the sighting device and the longitudinal axis of the missile.
The difference signal from the tilt winding and bearing winding, amplified in voltage and power in the tracking coordinator, enters the gyroscope correction windings. Under the influence of a moment from the correction system, the gyroscope precesses in the direction of decreasing the misalignment angle with the sighting axis of the sighting device and is locked in this position. The gyroscope is released by the ARP when the GGS is switched to tracking mode.
To maintain the rotation speed of the gyroscope rotor within the required limits, a speed stabilization system is used.
Steering compartment
The steering compartment includes the missile flight control equipment. The steering compartment housing contains a steering engine 2 (Fig. 29) with rudders 8, an on-board power supply consisting of a turbogenerator 6 and a stabilizer-rectifier 5, an angular velocity sensor 10, an amplifier /, a powder pressure accumulator 4, a powder control motor 3, a socket 7 (with cocking unit) and destabilizer
Rice. 29. Steering compartment: 1 - amplifier; 2 - steering gear; 3 - control motor; 4 - pressure accumulator; 5 - stabilizer-rectifier; 6 - turbogenerator; 7 - socket; 8 - rudders (plates); 9 - destabilizer; 10 - angular velocity sensor
Rice. 30. Steering gear:
1 - output ends of the coils; 2 - body; 3 - clamp; 4 - clip; 5 - filter; 6 - rudders; 7 - stopper; 8 - stand; 9 - bearing; 10 and 11 - springs; 12 - leash; 13 - nozzle; 14 - gas distribution sleeve; 15 - spool; 16 - bushing; 17 - right coil; 18 - anchor; 19 - piston; 20 - left coil; B and C - channels
Steering gear designed for aerodynamic control of a rocket in flight. At the same time, the PM serves as a distribution device in the gas-dynamic control system of the rocket at the initial part of the trajectory, when the aerodynamic control surfaces are ineffective. It is a gas amplifier of control electrical signals generated by the OGS.
The steering gear consists of a holder 4 (Fig. 30), in the lugs of which a working cylinder with a piston 19 and a fine filter 5 are located. A housing 2 with a spool valve is pressed into the holder, consisting of a four-edge spool 15, two bushings 16 and armatures 18. The housing contains two coils 17 and 20 of electromagnets. The cage has two eyes, in which a strut 8 with springs (spring) and a driver 12 pressed onto it is located on bearings 9. In the grooves of the driver and the strut there are rudders 6, which are held in the open position in flight by stoppers 7 and springs 10 and 11. In the boss of the cage, between the eyes, there is a gas distribution sleeve 14, rigidly fixed to the rack using a clamp 3. The sleeve has a groove with cut-off edges for supplying gas coming from the control unit to channels B, C and nozzles 13.
The PM operates from PAD gases, which flow through a pipe through a fine filter to the spool and from it through channels in the rings, housing and holder under the piston. Command signals from the OGS are sent one by one to the coils of the PM electromagnets. When current passes through the right coil 17 of the electromagnet, the armature 18 with the spool is attracted towards this electromagnet and opens the passage of gas into the left cavity of the working cylinder under the piston. Under gas pressure, the piston moves to the extreme right position until it touches the lid. Moving, the piston carries the protrusion of the leash with it and turns the leash and the stand, and with them the rudders, to their extreme position. At the same time, the gas distribution sleeve also rotates, while the cut-off edge allows access of gas from the PUD through the channel to the corresponding nozzle.
When current passes through the left coil 20 of the electromagnet, the piston moves to another extreme position.
At the moment of switching the current in the coils, when the force created by the powder gases exceeds the force of attraction of the electromagnet, the spool moves under the influence of the force from the powder gases, and the movement of the spool begins earlier than the current increases in the other coil, which increases the speed of the RM.
Onboard power supply designed to power the rocket equipment in flight. The source of energy for it is the gases formed during the combustion of the PAD charge.
The BIP consists of a turbogenerator and a stabilizer-rectifier. The turbogenerator consists of a stator 7 (Fig. 31), a rotor 4, on the axis of which a turbine 3 is mounted, which is its drive.
The stabilizer-rectifier performs two functions:
Converts the alternating current voltage of the turbogenerator into the required values of constant voltages and maintains their stability when the speed of rotation of the turbogenerator rotor and load current changes;
It regulates the rotation speed of the turbogenerator rotor when the gas pressure at the nozzle inlet changes by creating an additional electromagnetic load on the turbine shaft.
Rice. 31. Turbogenerator:
1 - stator; 2 - nozzle; 3 - turbine; 4 – rotor
BIP works as follows. Powder gases from the combustion of the PAD charge are supplied through nozzle 2 to the blades of the turbine 3 and cause it to rotate together with the rotor. In this case, an alternating EMF is induced in the stator winding, which is supplied to the input of the stabilizer-rectifier. From the output of the stabilizer-rectifier, constant voltage is supplied to the OGS and the DUS amplifier. The electric igniters VZ and PUD receive voltage from the BIP after the rocket exits the tube and the RM control surfaces are deployed.
Angular velocity sensor designed to generate an electrical signal proportional to the angular velocity of the rocket's oscillations relative to its transverse axes. This signal is used to dampen the angular vibrations of the rocket in flight. The DUS is a frame 1 consisting of two windings (Fig. 32), which is suspended on the axle shafts 2 in the center screws 3 with corundum bearings 4 and can be pumped in the working gaps of the magnetic circuit, consisting of base 5, permanent magnet 6 and shoes 7. The signal is collected from the DUS sensitive element (frame) through flexible torque-free stretchers 8, soldered to contacts 10 of the frame and contacts 9, electrically isolated from the body.
Rice. 32. Angular velocity sensor:
1 - frame; 2 - axle shaft; 3 - center screw; 4 - thrust bearing; 5 - base; 6 - magnet;
7 - shoe; 8 - stretching; 9 and 10 - contacts; 11 - casing
The CRS is installed so that its X-X axis coincides with the longitudinal axis of the rocket. When the rocket rotates only around the longitudinal axis, the frame, under the action of centrifugal forces, is installed in a plane perpendicular to the axis of rotation of the rocket.
The frame does not move in a magnetic field. EMF is not induced in its windings. When the rocket oscillates relative to the transverse axes, the frame moves in a magnetic field. The EMF induced in the windings of the frame is proportional to the angular velocity of the rocket's oscillations. The frequency of the EMF corresponds to the frequency of rotation around the longitudinal axis, and the phase of the signal corresponds to the direction of the absolute angular velocity vector of the rocket.
Powder pressure accumulator Designed to be supplied with powder gases RM and BIP. The PAD consists of housing 1 (Fig. 33), which is a combustion chamber, and filter 3, in which the gas is purified from solid particles. Gas flow and internal ballistics parameters are determined by the throttle opening 2. A powder charge 4 and an igniter 7, consisting of an electric igniter 8, a gunpowder sample 5 and a pyrotechnic firecracker 6, are placed inside the housing.
Rice. 34. Powder control motor:
7 - adapter; 3 - body; 3 - powder charge; 4 - weight of gunpowder; 5 - pyrotechnic firecracker; 6 - electric igniter; 7 - igniter
The PAD works as follows. An electrical impulse from the electronic unit of the trigger mechanism is sent to an electric igniter, which ignites a sample of gunpowder and a pyrotechnic firecracker, from the force of the flame of which the powder charge is ignited. The resulting powder gases are purified in a filter, after which they enter the PM and the BIP turbogenerator.
Powder control motor designed for gas-dynamic control of the rocket at the initial stage of the flight path. The PUD consists of a housing 2 (Fig. 34), which is a combustion chamber, and an adapter 1. Inside the housing there is a powder charge 3 and an igniter 7, consisting of an electric igniter 6, a sample of gunpowder 4 and a pyrotechnic firecracker 5. Gas flow and internal parameters ballistics are determined by the throttle hole in the adapter.
The PUD works as follows. After the rocket takes off from the launch tube and the rudders of the PM are opened, an electrical impulse from the charging capacitor is supplied to the electric igniter, which ignites the powder charge and the firecracker, from the force of the flame of which the powder charge ignites. Powder gases, passing through the distribution sleeve and two nozzles located perpendicular to the plane of the PM rudders, create a control force that ensures the missile turns.
Socket provides electrical connection between the rocket and the launch tube. It has main and control contacts, a circuit breaker for connecting capacitors C1 and C2 of the charging unit to the electric igniters VZ (EV1) and PUD, as well as for switching the positive terminal of the BIP to the VZ after the rocket has taken off from the tube and the RM control surfaces have been deployed.
Rice. 35. Charging block diagram:
1 - circuit breaker
The charging unit located in the socket body consists of capacitors C1 and C2 (Fig. 35), resistors R3 and R4 for removing residual voltage from the capacitors after checks or a failed start, resistors R1 and R2 for limiting the current in the capacitor circuit and diode D1, designed for electrical isolation of BIP and VZ circuits. Voltage is supplied to the cocking unit after the PM trigger is moved to the full stop position.
Destabilizer designed to provide overloads, the required stability and create additional torque, and therefore its plates are installed at an angle to the longitudinal axis of the rocket.
Warhead
The warhead is designed to engage an air target or cause damage to it, making it impossible to complete a combat mission.
The damaging factor of warheads is the high-explosive effect of the shock wave of warhead explosive products and residual fuel from the combustion chamber, as well as the fragmentation effect of elements formed during the explosion and crushing of the hull.
The warhead consists of the warhead itself, a contact fuse and an explosive generator. The warhead is the load-bearing compartment of the rocket and is made in the form of an integral connection.
The actual warhead (high-explosive fragmentation action) is designed to create a specified destruction field that affects the target after receiving an initiating impulse from the high explosive. It consists of housing 1 (Fig. 36), combat charge 2, detonator 4, cuff 5 and tube 3, through which wires pass from the air intake to the steering compartment of the rocket. On the body there is a yoke L, into the hole of which there is a pipe stopper intended for fixing the rocket in it.
Rice. 36. Warhead:
Warhead - the actual warhead; VZ - fuse; VG - explosive generator: 1- building;
2 - combat charge; 3 - tube; 4 - detonator; 5 - cuff; A - yoke
The fuse is designed to issue a detonation pulse to detonate the warhead charge when the missile hits the target or after the self-destruction time has elapsed, as well as to transfer the detonation pulse from the warhead charge to the explosive generator charge.
The electromechanical type fuse has two safety stages, which are removed in flight, which ensures the safe operation of the complex (start-up, maintenance, transportation and storage).
The fuse consists of a safety detonating device (PDD) (Fig. 37), a self-destruction mechanism, a tube, capacitors C1 and C2, the main target sensor GMD1 (pulse vortex magnetoelectric generator), backup target sensor GMD2 (pulse wave magnetoelectric generator), starting electric igniter EV1, two combat electric igniters EV2 and EVZ, a pyrotechnic moderator, an initiating charge, a blasting cap and a fuse detonator.
The remote control serves to ensure safety in handling the fuse until it is cocked after the missile is launched. It includes a pyrotechnic fuse, a rotating sleeve and a locking stop.
The fuse detonator is used to detonate the warhead. Target sensors GMD 1 and GMD2 ensure that the detonator capsule is triggered when the missile hits the target, and the self-destruction mechanism ensures that the detonator capsule is triggered after the self-destruction time has elapsed in case of a miss. The tube ensures the transfer of impulse from the warhead charge to the explosive generator charge.
Explosive generator - designed to detonate the unburned part of the propulsion charge and create an additional destruction field. It is a cup located in the fuse body with an explosive composition pressed into it.
When launching a rocket, the fuse and warhead operate as follows. When the rocket takes off from the tube, the rudders of the PM are opened, while the contacts of the socket breaker are closed and the voltage from the capacitor C1 of the charging unit is supplied to the electric igniter EV1 of the fuse, from which the pyrotechnic fuse of the remote control and the pyrotechnic pressing of the self-destruction mechanism are simultaneously ignited.
Rice. 37. Block diagram of the fuse
In flight, under the influence of axial acceleration from a running main engine, the blocking stopper of the remote control settles and does not prevent the rotation of the rotary sleeve (the first stage of protection has been removed). 1-1.9 seconds after the rocket launch, the pyrotechnic fuse burns out, and the spring rotates the rotating sleeve into the firing position. In this case, the axis of the detonator capsule is aligned with the axis of the fuse detonator, the contacts of the rotary sleeve are closed, the fuse is connected to the missile's BIP (the second stage of protection has been removed) and is ready for action. At the same time, the pyrotechnic pressing of the self-destruction mechanism continues to burn, and the BIP feeds capacitors C1 and C2 of the fuse on everything. throughout the flight.
When a missile hits a target at the moment the fuse passes through a metal barrier (when it breaks through) or along it (when it ricochets), an electric pulse occurs in the winding of the main target sensor GMD1 under the influence of eddy currents induced in the metal barrier when moving the permanent magnet of the target sensor GMD1 current This impulse is supplied to the electric igniter EVZ, from the beam of which the detonator capsule is triggered, causing the action of the fuse detonator. The fuse detonator initiates the warhead detonator, the operation of which causes the rupture of the warhead charge and the explosive in the fuse tube, which transmits the detonation to the explosive generator. In this case, the explosive generator is triggered and the remaining fuel of the remote control (if any) is detonated.
When a missile hits a target, the backup target sensor GMD2 is also triggered. Under the influence of the will of elastic deformations that occur when a missile meets an obstacle, the armature of the target sensor GMD2 comes off, the magnetic circuit breaks, as a result of which an electric current pulse is induced in the winding, which is supplied to the electric igniter EV2. The fire beam of electric igniter EV2 ignites a pyrotechnic retarder, the burning time of which exceeds the time required for the main target sensor GMD1 to approach the obstacle. After the moderator burns out, the initiating charge is triggered, causing the detonator cap and warhead detonator to fire, detonating the warhead and remaining fuel from the remote control (if any).
If the missile misses the target after the pyrotechnic pressing of the self-destruction mechanism burns out, the detonator cap is triggered by the fire beam, causing the detonator to act and detonate the warhead with an explosive generator to self-destruct the missile.
Propulsion system
The solid propellant propulsion system is designed to ensure that the rocket takes off from the tube, gives it the necessary angular rotation speed, accelerates to cruising speed and maintains this speed in flight.
The remote control consists of a starting engine, a dual-mode single-chamber propulsion engine and a delayed-action beam igniter.
The launch engine is designed to ensure the rocket takes off from the tube and gives it the required angular rotation speed. The starting engine consists of chamber 8 (Fig. 38), starting charge 6, starting charge igniter 7, diaphragm 5, disk 2, gas supply tube 1 and nozzle block 4. The starting charge consists of tubular powder blocks (or monolith) freely installed in the annular volume of the chamber. The igniter of the starting charge consists of a housing in which the electric igniter and a sample of gunpowder are located. The disk and diaphragm ensure the charge is secured during operation and transportation.
The starting engine is connected to the nozzle part of the main engine. When docking the engines, the gas supply tube is put on the body of the slow-action beam igniter 7 (Fig. 39), located in the pre-nozzle volume of the main engine. This connection ensures the transfer of the fire impulse to the beam igniter. The electrical connection of the starting engine igniter with the starting tube is carried out through contact connection 9 (Fig. 38).
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Rice. 38. Starting motor:
1 - gas supply tube; 2 - disk; 3 - plug; 4 - nozzle block; 5 - diaphragm; 6 - starting charge; 7 - starting charge igniter; 8 - camera; 9 - contact connection
The nozzle block has seven (or six) nozzles located at an angle to the longitudinal axis of the rocket, ensuring rotation of the rocket in the area where the starting engine is operating. To ensure the tightness of the remote control chamber during operation and to create the necessary pressure when the starting charge is ignited, plugs 3 are installed in the nozzles.
Dual-mode single-chamber propulsion engine designed to ensure acceleration of the rocket to cruising speed in the first mode and maintaining this speed in flight in the second mode.
The main engine consists of chamber 3 (Fig. 39), main charge 4, main charge igniter 5, nozzle block 6 and delayed-action beam igniter 7. The bottom 1 with seats for docking the remote control and warhead is screwed into the front part of the chamber. To obtain the required combustion modes, the charge is partially armored and reinforced with six wires 2.
1 – bottom; 2 – wires; 3 – camera; 4 – sustaining charge; 5 – sustaining charge igniter; 6 – nozzle block; 7 – delayed-action beam igniter; 8 – plug; A – threaded hole
Rice. 40. Delayed beam igniter: 1 - pyrotechnic retarder; 2 - body; 3 - bushing; 4 - transfer charge; 5 - deton. charge
Rice. 41. Wing block:
1 - plate; 2 - front liner; 3 - body; 4 - axis; 5 - spring; 6 - stopper; 7 - screw; 8 - rear liner; B - protrusion
To ensure the tightness of the chamber during operation and to create the necessary pressure when the propulsion charge is ignited, a plug 8 is installed on the nozzle block, which is destroyed and burned by the propellant gases of the propulsion engine. On the outer part of the nozzle block there are threaded holes A for attaching the wing block to the remote control.
The delayed-action beam igniter is designed to ensure the operation of the main engine at a safe distance for the anti-aircraft gunner. During its combustion time, equal to 0.33 - 0.5 s, the rocket moves away from the anti-aircraft gunner to a distance of at least 5.5 m. This protects the anti-aircraft gunner from the effects of the jet of powder gases from the main engine.
A delayed-action beam igniter consists of a housing 2 (Fig. 40), which houses a pyrotechnic moderator 1, a transfer charge 4 in a sleeve 3. On the other side, a detonating charge 5 is pressed into the sleeve. From the powder gases formed in the starting engine chamber when the charge burns , the detonating charge ignites. The shock wave generated during detonation is transmitted through the wall of the bushing and ignites the transfer charge, which ignites the pyrotechnic moderator. After a delay time, the igniter of the sustaining charge ignites from the pyrotechnic moderator, which ignites the sustaining charge.
The remote control works as follows. When an electrical impulse is applied to the electric igniter of the starting charge, the igniter is fired, and then the starting charge. Under the influence of the reactive force created by the starting engine, the rocket flies out of the tube with the required angular velocity of rotation. The starting motor finishes working in the pipe and lingers there. From the powder gases formed in the starting engine chamber, a delayed-action beam igniter is triggered, igniting the igniter of the sustaining charge, from which the sustaining charge is fired at a safe distance for the anti-aircraft gunner. The reaction force created by the propulsion engine accelerates the rocket to cruising speed and maintains this speed in flight.
Wing block
The wing unit is designed to aerodynamically stabilize the rocket in flight, create lift at angles of attack and maintain the required rotation speed of the rocket along the trajectory.
The wing block consists of body 3 (Fig. 41), four folding wings and a mechanism for locking them.
The folding wing consists of a plate 7, which is attached with two screws 7 to liners 2 and 8, put on an axis 4 located in the hole in the housing.
The locking mechanism consists of two stoppers 6 and a spring 5, with the help of which the stoppers unclench and lock the wing when opened. After the rotating rocket leaves the tube, the wings open under the influence of centrifugal forces. To maintain the required rotation speed of the rocket in flight, the wings are rotated relative to the longitudinal axis of the wing unit at a certain angle.
The wing block is screwed onto the nozzle block of the main engine. There are four protrusions B on the body of the wing block for connecting it to the starting engine using an expanding connecting ring.
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Rice. 42. Pipe 9P39(9P39-1*)
1 - front cover; 2 and 11 - locks; 3 - sensor block; 4 - antenna; 5 - clips; 6 and 17 – covers; 7 – diaphragm; 8 – shoulder strap; 9 – clip; 10 – pipe; 12 - back cover; 13 - lamp; 14 - screw; 15 - block; 16 - glow mechanism lever; 18. 31 and 32 – springs; 19 38 – clamps; 20 – connector; 21 – rear pillar; 22 - side connector mechanism; 23 - handle; 24 - front pillar; 25 - fairing; 26 - nozzles; 27 – board; 28 – pin contacts; 29 – guide pins; 30 - stopper; 33 - traction; 34 - fork; 35 - body; 36 - button; 37 - eye; A and E - marks; B and M – holes; B – front sight; G – rear sight; D – triangular mark; F – cutout; I – guides; K - bevel; L and U - surfaces; D - groove; P and C – diameters; F – nests; Ш – board; Shch and E – gasket; Yu – overlay; I am a shock absorber;
*) Note:
1. Two types of pipes can be used: 9P39 (with antenna 4) and 9P39-1 (without antenna 4)
2. There are 3 variants of mechanical sights with a light information lamp in operation.
Homing head
A homing head is an automatic device that is installed on a controlled weapon in order to ensure high accuracy of targeting the target.
The main parts of the homing head are: a coordinator with a receiver (and sometimes with an energy emitter) and an electronic computing device. The coordinator searches, captures and tracks the target. The electronic computing device processes the information received from the coordinator and transmits signals that control the coordinator and the movement of the controlled weapon.
Based on the principle of operation, the following homing heads are distinguished:
1) passive – receiving energy emitted by the target;
2) semi-active - reacting to the energy reflected by the target, which is emitted by some external source;
3) active - receiving energy reflected from the target, which is emitted by the homing head itself.
Based on the type of energy received, homing heads are divided into radar, optical, and acoustic.
The acoustic homing head operates using audible sound and ultrasound. Its most effective use is in water, where sound waves attenuate more slowly than electromagnetic waves. Heads of this type are installed on controlled means of destroying sea targets (for example, acoustic torpedoes).
The optical homing head operates using electromagnetic waves in the optical range. Installed on controlled means of destruction of ground, air and sea targets. Aiming is carried out by a source of infrared radiation or by the reflected energy of a laser beam. On controlled means of destroying ground targets, which are classified as non-contrast, passive optical homing heads are used, which operate according to the optical image of the terrain.
Radar homing heads operate using electromagnetic radio waves. Active, semi-active and passive radar heads are used on controlled means of destruction of ground, air and sea targets. On controlled means of destruction of non-contrast ground targets, active homing heads are used, which operate using radio signals reflected from the terrain, or passive ones, which operate based on radio-thermal radiation from the area.
This text is an introductory fragment. From the book Locksmith's Guide to Locks by Phillips Bill From the book Locksmith's Guide to Locks by Phillips Bill author Team of authorsDividing head Dividing head is a device used for setting, securing and periodically turning or continuously rotating small workpieces processed on milling machines. In tool shops of machine-building enterprises
From the book Great Encyclopedia of Technology author Team of authorsTurret head The turret head is a special device in which various cutting tools are installed: drills, countersinks, reamers, taps, etc. The turret head is an important component of turret lathes (automatic and
From the book Great Encyclopedia of Technology author Team of authorsHoming head The homing head is an automatic device that is installed on a controlled weapon in order to ensure high accuracy of aiming at the target. The main parts of the homing head are: a coordinator with
From the book Great Soviet Encyclopedia (DE) by the author TSB From the book Great Soviet Encyclopedia (VI) by the author TSB From the book Great Soviet Encyclopedia (GO) by the author TSB From the book Great Soviet Encyclopedia (MA) by the author TSB From the book Great Soviet Encyclopedia (RA) by the author TSB From the book The Big Book of the Amateur Angler [with color insert] author Goryainov Alexey GeorgievichSinker head Today this device is more often called a jig head. It resembles a large jig with a fastening ring and a stopper for the bait. Spinning sinker heads are used mainly for horizontal guiding of soft baits and can vary in weight and
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