Active radar homing head args. Homing heads Television homing head working principle
State Committee of the Russian Federation for Higher Education
BALTIC STATE TECHNICAL UNIVERSITY
_____________________________________________________________
Department of Radioelectronic Devices
RADAR HOMING HEAD
St. Petersburg
2. GENERAL INFORMATION ABOUT RLGS.
2.1 Purpose
The radar homing head is installed on the surface-to-air missile to ensure automatic target acquisition, its auto-tracking and the issuance of control signals to the autopilot (AP) and radio fuse (RB) at the final stage of the missile's flight.
2.2 Specifications
RLGS is characterized by the following basic performance data:
1. search area by direction:
Azimuth ± 10°
Elevation ± 9°
2. search area review time 1.8 - 2.0 sec.
3. target acquisition time by angle 1.5 sec (no more)
4. Maximum angles of deviation of the search area:
In azimuth ± 50° (not less than)
Elevation ± 25° (not less than)
5. Maximum deviation angles of the equisignal zone:
In azimuth ± 60° (not less than)
Elevation ± 35° (not less than)
6. target capture range of the IL-28 aircraft type with the issuance of control signals to (AP) with a probability of not less than 0.5 -19 km, and with a probability of not less than 0.95 -16 km.
7 search zone in range 10 - 25 km
8. operating frequency range f ± 2.5%
9. average transmitter power 68W
10. RF pulse duration 0.9 ± 0.1 µs
11. RF pulse repetition period T ± 5%
12. sensitivity of receiving channels - 98 dB (not less)
13.power consumption from power sources:
From the mains 115 V 400 Hz 3200 W
Mains 36V 400Hz 500W
From the network 27 600 W
14. station weight - 245 kg.
3. PRINCIPLES OF OPERATION AND CONSTRUCTION OF RLGS
3.1 The principle of operation of the radar
RLGS is a radar station of the 3-centimeter range, operating in the mode of pulsed radiation. At the most general consideration, the radar station can be divided into two parts: - the actual radar part and the automatic part, which provides target acquisition, its automatic tracking in angle and range, and the issuance of control signals to the autopilot and radio fuse.
The radar part of the station works in the usual way. High-frequency electromagnetic oscillations generated by the magnetron in the form of very short pulses are emitted using a highly directional antenna, received by the same antenna, converted and amplified in the receiving device, pass further to the automatic part of the station - the target angle tracking system and the rangefinder.
The automatic part of the station consists of the following three functional systems:
1. antenna control systems that provide antenna control in all modes of operation of the radar station (in the "pointing" mode, in the "search" mode and in the "homing" mode, which in turn is divided into "capture" and "autotracking" modes)
2. distance measuring device
3. a calculator for control signals supplied to the autopilot and radio fuse of the rocket.
The antenna control system in the "auto-tracking" mode works according to the so-called differential method, in connection with which a special antenna is used in the station, consisting of a spheroidal mirror and 4 emitters placed at some distance in front of the mirror.
When the radar station operates on radiation, a single-lobe radiation pattern is formed with a maμmum coinciding with the axis of the antenna system. This is achieved due to the different lengths of the waveguides of the emitters - there is a hard phase shift between the oscillations of different emitters.
When working at reception, the radiation patterns of the emitters are shifted relative to the optical axis of the mirror and intersect at a level of 0.4.
The connection of the emitters with the transceiver is carried out through a waveguide path, in which there are two ferrite switches connected in series:
· Axes commutator (FKO), operating at a frequency of 125 Hz.
· Receiver switch (FKP), operating at a frequency of 62.5 Hz.
Ferrite switches of the axes switch the waveguide path in such a way that first all 4 emitters are connected to the transmitter, forming a single-lobe directivity pattern, and then to a two-channel receiver, then emitters that create two directivity patterns located in a vertical plane, then emitters that create two patterns orientation in the horizontal plane. From the outputs of the receivers, the signals enter the subtraction circuit, where, depending on the position of the target relative to the equi-signal direction formed by the intersection of the radiation patterns of a given pair of emitters, a difference signal is generated, the amplitude and polarity of which is determined by the position of the target in space (Fig. 1.3).
Synchronously with the ferrite axis switch in the radar station, the antenna control signal extraction circuit operates, with the help of which the antenna control signal is generated in azimuth and elevation.
The receiver commutator switches the inputs of the receiving channels at a frequency of 62.5 Hz. The switching of receiving channels is associated with the need to average their characteristics, since the differential method of target direction finding requires the complete identity of the parameters of both receiving channels. The RLGS rangefinder is a system with two electronic integrators. From the output of the first integrator, a voltage proportional to the speed of approach to the target is removed, from the output of the second integrator - a voltage proportional to the distance to the target. The range finder captures the nearest target in the range of 10-25 km with its subsequent auto-tracking up to a range of 300 meters. At a distance of 500 meters, a signal is emitted from the rangefinder, which serves to cock the radio fuse (RV).
The RLGS calculator is a computing device and serves to generate control signals issued by the RLGS to the autopilot (AP) and RV. A signal is sent to the AP, representing the projection of the vector of the absolute angular velocity of the target sighting beam on the transverse axes of the missile. These signals are used to control the missile's heading and pitch. A signal representing the projection of the velocity vector of the target's approach to the missile onto the polar direction of the target's sighting beam arrives at the RV from the computer.
The distinctive features of the radar station in comparison with other stations similar to it in terms of their tactical and technical data are:
1. the use of a long-focus antenna in a radar station, characterized by the fact that the beam is formed and deflected in it using the deflection of one rather light mirror, the deflection angle of which is half that of the beam deflection angle. In addition, there are no rotating high-frequency transitions in such an antenna, which simplifies its design.
2. use of a receiver with a linear-logarithmic amplitude characteristic, which provides an expansion of the dynamic range of the channel up to 80 dB and, thereby, makes it possible to find the source of active interference.
3. building a system of angular tracking by the differential method, which provides high noise immunity.
4. application in the station of the original two-circuit closed yaw compensation circuit, which provides a high degree of compensation for the rocket oscillations relative to the antenna beam.
5. constructive implementation of the station according to the so-called container principle, which is characterized by a number of advantages in terms of reducing the total weight, using the allotted volume, reducing interconnections, the possibility of using a centralized cooling system, etc.
3.2 Separate functional radar systems
RLGS can be divided into a number of separate functional systems, each of which solves a well-defined particular problem (or several more or less closely related particular problems) and each of which is to some extent designed as a separate technological and structural unit. There are four such functional systems in the RLGS:
3.2.1 Radar part of the RLGS
The radar part of the RLGS consists of:
the transmitter.
receiver.
high voltage rectifier.
the high frequency part of the antenna.
The radar part of the RLGS is intended:
· to generate high-frequency electromagnetic energy of a given frequency (f ± 2.5%) and a power of 60 W, which is radiated into space in the form of short pulses (0.9 ± 0.1 μs).
· for the subsequent reception of signals reflected from the target, their conversion into intermediate frequency signals (Fpch = 30 MHz), amplification (via 2 identical channels), detection and delivery to other radar systems.
3.2.2. Synchronizer
Synchronizer consists of:
Receiving and Synchronization Manipulation Unit (MPS-2).
· receiver switching unit (KP-2).
· Control unit for ferrite switches (UF-2).
selection and integration node (SI).
Error signal selection unit (CO)
· ultrasonic delay line (ULZ).
generation of synchronization pulses for launching individual circuits in the radar station and control pulses for the receiver, SI unit and rangefinder (MPS-2 unit)
Formation of impulses for controlling the ferrite switch of axes, the ferrite switch of the receiving channels and the reference voltage (UV-2 node)
Integration and summation of received signals, voltage regulation for AGC control, conversion of target video pulses and AGC into radio frequency signals (10 MHz) for their delay in the ULZ (SI node)
· isolation of the error signal necessary for the operation of the angular tracking system (CO node).
3.2.3. Rangefinder
The rangefinder consists of:
Time modulator node (EM).
time discriminator node (VD)
two integrators.
The purpose of this part of the RLGS is:
search, capture and tracking of the target in range with the issuance of signals of the range to the target and the speed of approach to the target
issuance of signal D-500 m
Issuance of selection pulses for receiver gating
Issuance of pulses limiting the reception time.
3.2.4. Antenna Control System (AMS)
The antenna control system consists of:
Search and gyro stabilization unit (PGS).
Antenna head control unit (UGA).
· knot of the automatic capture (A3).
· storage unit (ZP).
· output nodes of the antenna control system (AC) (on the channel φ and channel ξ).
Electric spring assembly (SP).
The purpose of this part of the RLGS is:
control of the antenna during rocket takeoff in the modes of guidance, search and preparation for capture (assemblies of PGS, UGA, US and ZP)
Target capture by angle and its subsequent auto-tracking (nodes A3, ZP, US, and ZP)
4. OPERATING PRINCIPLE OF THE ANGLE TRACKING SYSTEM
In the functional diagram of the angular target tracking system, the reflected high-frequency pulse signals received by two vertical or horizontal antenna emitters are fed through the ferrite switch (FKO) and the ferrite switch of the receiving channels - (FKP) to the input flanges of the radio frequency receiving unit. To reduce reflections from the detector sections of the mixers (SM1 and SM2) and from the receiver protection arresters (RZP-1 and RZP-2) during the recovery time of the RZP, which worsen the decoupling between the receiving channels, resonant ferrite valves (FV- 1 and FV-2). The reflected pulses received at the inputs of the radio frequency receiving unit are fed through the resonant valves (F A-1 and F V-2) to the mixers (CM-1 and CM-2) of the corresponding channels, where, mixing with the oscillations of the klystron generator, they are converted into pulses of the intermediate frequencies. From the outputs of the mixers of the 1st and 2nd channels, the intermediate frequency pulses are fed to the intermediate frequency preamplifiers of the corresponding channels - (PUFC unit). From the output of the PUFC, the amplified intermediate frequency signals are fed to the input of a linear-logarithmic intermediate frequency amplifier (UPCL nodes). Linear-logarithmic intermediate frequency amplifiers amplify, detect and subsequently amplify the video frequency of the intermediate frequency pulses received from the PUFC.
Each linear-logarithmic amplifier consists of the following functional elements:
Logarithmic amplifier, which includes an IF (6 stages)
Transistors (TR) for decoupling the amplifier from the addition line
Signal addition lines (LS)
Linear detector (LD), which in the range of input signals of the order of 2-15 dB gives a linear dependence of the input signals on the output
The summing cascade (Σ), in which the linear and logarithmic components of the characteristic are added
Video amplifier (VU)
The linear-logarithmic characteristic of the receiver is necessary to expand the dynamic range of the receiving path up to 30 dB and eliminate overloads caused by interference. If we consider the amplitude characteristic, then in the initial section it is linear and the signal is proportional to the input, with an increase in the input signal, the increment of the output signal decreases.
To obtain a logarithmic dependence in UPCL, the method of sequential detection is used. The first six stages of the amplifier work as linear amplifiers at low input signal levels and as detectors at high signal levels. The video pulses generated during detection are fed from the emitters of the IF transistors to the bases of the decoupling transistors, on the common collector load of which they are added.
To obtain the initial linear section of the characteristic, the signal from the output of the IF is fed to a linear detector (LD). The overall linear-logarithmic dependence is obtained by adding the logarithmic and linear amplitude characteristics in the addition stage.
Due to the need to have a fairly stable noise level of the receiving channels. In each receiving channel, a system of inertial automatic noise gain control (AGC) is used. For this purpose, the output voltage from the UPCL node of each channel is fed to the PRU node. Through the preamplifier (PRU), the key (CL), this voltage is fed to the error generation circuit (CBO), into which the reference voltage "noise level" from resistors R4, R5 is also introduced, the value of which determines the noise level at the receiver output. The difference between the noise voltage and the reference voltage is the output signal of the video amplifier of the AGC unit. After appropriate amplification and detection, the error signal in the form of a constant voltage is applied to the last stage of the PUCH. To exclude the operation of the AGC node from various kinds of signals that may occur at the input of the receiving path (the AGC should work only on noise), switching of both the AGC system and the block klystron has been introduced. The AGC system is normally locked and opens only for the duration of the AGC strobe pulse, which is located outside the area of reflected signal reception (250 μs after the TX start pulse). In order to exclude the influence of various kinds of external interference on the noise level, the generation of the klystron is interrupted for the duration of the AGC, for which the strobe pulse is also fed to the klystron reflector (through the output stage of the AFC system). (Figure 2.4)
It should be noted that the disruption of klystron generation during AGC operation leads to the fact that the noise component that is created by the mixer is not taken into account by the AGC system, which leads to some instability in the overall noise level of the receiving channels.
Almost all control and switching voltages are connected to the PUCH nodes of both channels, which are the only linear elements of the receiving path (at the intermediate frequency):
· AGC regulating voltages;
The radio-frequency receiving unit of the radar station also contains a klystron automatic frequency control circuit (AFC), due to the fact that the tuning system uses a klystron with dual frequency control - electronic (in a small frequency range) and mechanical (in a large frequency range) AFC system also divided into electronic and electromechanical frequency control system. The voltage from the output of the electronic AFC is fed to the klystron reflector and performs electronic frequency adjustment. The same voltage is fed to the input of the electromechanical frequency control circuit, where it is converted into an alternating voltage, and then fed to the motor control winding, which performs mechanical frequency adjustment of the klystron. To find the correct setting of the local oscillator (klystron), corresponding to a difference frequency of about 30 MHz, the AFC provides for an electromechanical search and capture circuit. The search takes place over the entire frequency range of the klystron in the absence of a signal at the AFC input. The AFC system works only during the emission of a probing pulse. For this, the power supply of the 1st stage of the AFC node is carried out by a differentiated start pulse.
From the UPCL outputs, the video pulses of the target enter the synchronizer to the summation circuit (SH "+") in the SI node and to the subtraction circuit (SH "-") in the CO node. The target pulses from the outputs of the UPCL of the 1st and 2nd channels, modulated with a frequency of 123 Hz (with this frequency the axes are switched), through the emitter followers ZP1 and ZP2 enter the subtraction circuit (SH "-"). From the output of the subtraction circuit, the difference signal obtained as a result of subtracting the signals of the 1st channel from the signals of the 2nd channel of the receiver enters the key detectors (KD-1, KD-2), where it is selectively detected and the error signal is separated along the axes " ξ" and "φ". The enabling pulses necessary for the operation of the key detectors are generated in special circuits in the same node. One of the permissive pulse generation circuits (SFRI) receives integrated target pulses from the "SI" synchronizer node and a reference voltage of 125– (I) Hz, the other receives integrated target pulses and a reference voltage of 125 Hz – (II) in antiphase. Enable pulses are formed from the pulses of the integrated target at the time of the positive half-cycle of the reference voltage.
The reference voltages of 125 Hz - (I), 125 Hz - (II), shifted relative to each other by 180, necessary for the operation of the permissive pulse generation circuits (SFRI) in the CO synchronizer node, as well as the reference voltage through the "φ" channel, are generated by sequential dividing by 2 the station repetition rate in the KP-2 node (switching receivers) of the synchronizer. Frequency division is performed using frequency dividers, which are RS flip-flops. The frequency divider start pulse generation circuit (ОΦЗ) is triggered by the trailing edge of a differentiated negative reception time limit pulse (T = 250 μs), which comes from the range finder. From the voltage output circuit of 125 Hz - (I), and 125 Hz - (II) (CB), a synchronization pulse with a frequency of 125 Hz is taken, which is fed to the frequency divider in the UV-2 (DCh) node. In addition, a voltage of 125 Hz is supplied to the circuit forming a shift by 90 relative to the reference voltage. The circuit for generating the reference voltage over the channel (TOH φ) is assembled on a trigger. A synchronization pulse of 125 Hz is fed to the divider circuit in the UV-2 node, the reference voltage "ξ" with a frequency of 62.5 Hz is removed from the output of this divider (DF), supplied to the US node and also to the KP-2 node to form a shifted by 90 degrees of reference voltage.
The UF-2 node also generates axes switching current pulses with a frequency of 125 Hz and receiver switching current pulses with a frequency of 62.5 Hz (Fig. 4.4).
The enabling pulse opens the transistors of the key detector and the capacitor, which is the load of the key detector, is charged to a voltage equal to the amplitude of the resulting pulse coming from the subtraction circuit. Depending on the polarity of the incoming pulse, the charge will be positive or negative. The amplitude of the resulting pulses is proportional to the angle of mismatch between the direction to the target and the direction of the equisignal zone, so the voltage to which the capacitor of the key detector is charged is the voltage of the error signal.
From the key detectors, an error signal with a frequency of 62.5 Hz and an amplitude proportional to the angle of mismatch between the direction to the target and the direction of the equisignal zone arrives through the RFP (ZPZ and ZPCH) and video amplifiers (VU-3 and VU-4) to the nodes US-φ and US-ξ of the antenna control system (Fig. 6.4).
The target pulses and UPCL noise of the 1st and 2nd channels are also fed to the CX+ addition circuit in the synchronizer node (SI), in which time selection and integration are carried out. Time selection of pulses by repetition frequency is used to combat non-synchronous impulse noise. Radar protection from non-synchronous impulse interference can be carried out by applying to the coincidence circuit non-delayed reflected signals and the same signals, but delayed for a time exactly equal to the repetition period of the emitted pulses. In this case, only those signals whose repetition period is exactly equal to the repetition period of the emitted pulses will pass through the coincidence circuit.
From the output of the addition circuit, the target pulse and noise through the phase inverter (Φ1) and the emitter follower (ZP1) are fed to the coincidence stage. The summation circuit and the coincidence cascade are elements of a closed-loop integration system with positive feedback. The integration scheme and the selector work as follows. The input of the circuit (Σ) receives the pulses of the summed target with noise and the pulses of the integrated target. Their sum goes to the modulator and generator (MiG) and to the ULZ. This selector uses an ultrasonic delay line. It consists of a sound duct with electromechanical energy converters (quartz plates). ULZ can be used to delay both RF pulses (up to 15 MHz) and video pulses. But when the video pulses are delayed, a significant distortion of the waveform occurs. Therefore, in the selector circuit, the signals to be delayed are first converted using a special generator and modulator into RF pulses with a duty cycle of 10 MHz. From the output of the ULZ, the target impulse delayed for the period of repetition of the radar is fed to the UPCH-10, from the output of the UPCH-10, the signal delayed and detected on the detector (D) through the key (CL) (UPC-10) is fed to the coincidence cascade (CS), to this the same cascade is supplied with the summed target impulse.
At the output of the coincidence stage, a signal is obtained that is proportional to the product of favorable voltages, therefore, the target pulses, synchronously arriving at both inputs of the COP, easily pass the coincidence stage, and noise and non-synchronous interference are strongly suppressed. From the output (CS), the target pulses through the phase inverter (Φ-2) and (ZP-2) again enter the circuit (Σ), thereby closing the feedback ring, in addition, the integrated target pulses enter the CO node, to the circuits for generating allowing key impulses, detectors (OFRI 1) and (OFRI 2).
The integrated pulses from the key output (CL), in addition to the coincidence cascade, are fed to the protection circuit against non-synchronous impulse noise (SZ), on the second arm of which the summed target pulses and noises from (3P 1) are received. The anti-synchronous interference protection circuit is a diode coincidence circuit that passes the smaller of the two voltages synchronously applied to its inputs. Since the integrated target pulses are always much larger than the summed ones, and the voltage of noise and interference is strongly suppressed in the integration circuit, then in the coincidence circuit (CZ), in essence, the summed target pulses are selected by the integrated target pulses. The resulting "direct target" pulse has the same amplitude and shape as the stacked target pulse, while noise and jitter are suppressed. The impulse of the direct target is supplied to the time discriminator of the rangefinder circuit and the node of the capture machine, the antenna control system. Obviously, when using this selection scheme, it is necessary to ensure a very accurate equality between the delay time in the CDL and the repetition period of the emitted pulses. This requirement can be met by using special schemes for the formation of synchronization pulses, in which the stabilization of the pulse repetition period is carried out by the LZ of the selection scheme. The synchronization pulse generator is located in the MPS - 2 node and is a blocking oscillator (ZVG) with its own self-oscillation period, slightly longer than the delay time in the LZ, i.e. more than 1000 µs. When the radar is turned on, the first ZVG pulse is differentiated and launches the BG-1, from the output of which several synchronization pulses are taken:
· Negative clock pulse T=11 µs is fed along with the rangefinder selection pulse to the circuit (CS), which generates the control pulses of the SI node for the duration of which the manipulation cascade (CM) opens in the node (SI) and the addition cascade (CX +) and all subsequent ones work. As a result, the BG1 synchronization pulse passes through (SH +), (Φ 1), (EP-1), (Σ), (MiG), (ULZ), (UPC-10), (D) and delayed by the radar repetition period (Tp=1000µs), triggers the ZBG with a rising edge.
· Negative locking pulse UPC-10 T = 12 μs locks the key (KL) in the SI node and thereby prevents the BG-1 synchronization pulse from entering the circuit (KS) and (SZ).
· Negative differentiated impulse synchronization triggers the rangefinder start pulse generation circuit (SΦZD), the rangefinder start pulse synchronizes the time modulator (TM), and also through the delay line (DL) is fed to the start pulse generation circuit of the transmitter SΦZP. In the circuit (VM) of the range finder, negative pulses of the reception time limit f = 1 kHz and T = 250 μs are formed along the front of the range finder start pulse. They are fed back to the MPS-2 node on the CBG to exclude the possibility of triggering the CBG from the target pulse, in addition, the trailing edge of the receive time limit pulse triggers the AGC strobe pulse generation circuit (SFSI), and the AGC strobe pulse triggers the manipulation pulse generation circuit (СΦМ ). These pulses are fed into the RF unit.
Error signals from the output of the node (CO) of the synchronizer are fed to the nodes of the angular tracking (US φ, US ξ) of the antenna control system to the error signal amplifiers (USO and USO). From the output of the error signal amplifiers, the error signals are fed to the paraphase amplifiers (PFC), from the outputs of which the error signals in opposite phases are fed to the inputs of the phase detector - (PD 1). Reference voltages are also supplied to the phase detectors from the outputs of PD 2 of reference voltage multivibrators (MVON), the inputs of which are supplied with reference voltages from the UV-2 unit (φ channel) or the KP-2 unit (ξ channel) of the synchronizer. From the outputs of phase signal voltage detectors, errors are fed to the contacts of the capture preparation relay (RPZ). Further operation of the node depends on the mode of operation of the antenna control system.
5. RANGEFINDER
The RLGS 5G11 rangefinder uses an electrical range measurement circuit with two integrators. This scheme allows you to get a high speed of capturing and tracking the target, as well as giving the range to the target and the speed of approach in the form of a constant voltage. The system with two integrators memorizes the last rate of approach in case of a short-term loss of the target.
The operation of the rangefinder can be described as follows. In the time discriminator (TD), the time delay of the pulse reflected from the target is compared with the time delay of the tracking pulses ("Gate"), created by the electrical time modulator (TM), which includes a linear delay circuit. The circuit automatically provides equality between gate delay and target pulse delay. Since the target pulse delay is proportional to the distance to the target, and the gate delay is proportional to the voltage at the output of the second integrator, in the case of a linear relationship between the gate delay and this voltage, the latter will be proportional to the distance to the target.
The time modulator (TM), in addition to the “gate” pulses, generates a reception time limit pulse and a range selection pulse, and, depending on whether the radar station is in the search or target acquisition mode, its duration changes. In the "search" mode T = 100 μs, and in the "capture" mode T = 1.5 μs.
6. ANTENNA CONTROL SYSTEM
In accordance with the tasks performed by the SUA, the latter can be conditionally divided into three separate systems, each of which performs a well-defined functional task.
1. Antenna head control system. It includes:
UGA node
Scheme of storing on the channel "ξ" in the node ZP
· drive - an electric motor of the SD-10a type, controlled by an electric machine amplifier of the UDM-3A type.
2. Search and gyro stabilization system. It includes:
PGS node
output cascades of US nodes
Scheme of storing on the channel "φ" in the node ZP
· a drive on electromagnetic piston couplings with an angular velocity sensor (DSUs) in the feedback circuit and the ZP unit.
3. Angular target tracking system. It includes:
nodes: US φ, US ξ, A3
Scheme for highlighting the error signal in the CO synchronizer node
· drive on electromagnetic powder clutches with CRS in feedback and SP unit.
It is advisable to consider the operation of the control system sequentially, in the order in which the rocket performs the following evolutions:
1. "take off",
2. "guidance" on commands from the ground
3. "search for the target"
4. "pre-capture"
5. "ultimate capture"
6. "automatic tracking of a captured target"
With the help of a special kinematic scheme of the block, the necessary law of motion of the antenna mirror is provided, and, consequently, the movement of the directivity characteristics in azimuth (φ axis) and inclination (ξ axis) (fig.8.4).
The trajectory of the antenna mirror depends on the operating mode of the system. In mode "escort" the mirror can perform only simple movements along the φ axis - through an angle of 30 °, and along the ξ axis - through an angle of 20 °. When operating in "Search", the mirror performs a sinusoidal oscillation about the φ n axis (from the drive of the φ axis) with a frequency of 0.5 Hz and an amplitude of ± 4°, and a sinusoidal oscillation about the ξ axis (from the cam profile) with a frequency f = 3 Hz and an amplitude of ± 4°.
Thus, viewing of the 16"x16" zone is provided. the angle of deviation of the directivity characteristic is 2 times the angle of rotation of the antenna mirror.
In addition, the viewing area is moved along the axes (by the drives of the corresponding axes) by commands from the ground.
7. MODE "TAKEOFF"
When the rocket takes off, the radar antenna mirror must be in the zero position "top-left", which is provided by the PGS system (along the φ axis and along the ξ axis).
8. POINT MODE
In the guidance mode, the position of the antenna beam (ξ = 0 and φ = 0) in space is set using control voltages, which are taken from the potentiometers and the search area gyro stabilization unit (GS) and are brought into the channels of the OGM unit, respectively.
After launching the missile into level flight, a one-time "guidance" command is sent to the RLGS through the onboard command station (SPC). On this command, the PGS node keeps the antenna beam in a horizontal position, turning it in azimuth in the direction specified by the commands from the ground "turn the zone along" φ ".
The UGA system in this mode keeps the antenna head in the zero position relative to the "ξ" axis.
9. MODE "SEARCH".
When the missile approaches the target to a distance of approximately 20-40 km, a one-time "search" command is sent to the station through the SPC. This command arrives at the node (UGA), and the node switches to the high-speed servo system mode. In this mode, the sum of a fixed frequency signal of 400 Hz (36V) and the high-speed feedback voltage from the TG-5A current generator are supplied to the input of the AC amplifier (AC) of the node (UGA). In this case, the shaft of the executive motor SD-10A begins to rotate at a fixed speed, and through the cam mechanism causes the antenna mirror to swing relative to the rod (i.e., relative to the "ξ" axis) with a frequency of 3 Hz and an amplitude of ± 4°. At the same time, the engine rotates a sinus potentiometer - a sensor (SPD), which outputs a "winding" voltage with a frequency of 0.5 Hz to the azimuth channel of the OPO system. This voltage is applied to the summing amplifier (US) of the node (CS φ) and then to the antenna drive along the axis. As a result, the antenna mirror begins to oscillate in azimuth with a frequency of 0.5 Hz and an amplitude of ± 4°.
Synchronous swinging of the antenna mirror by the UGA and OPO systems, respectively in elevation and azimuth, creates a search beam movement shown in Fig. 3.4.
In the "search" mode, the outputs of the phase detectors of the nodes (US - φ and US - ξ) are disconnected from the input of the summing amplifiers (SU) by the contacts of a de-energized relay (RPZ).
In the "search" mode, the processing voltage "φ n" and the voltage from the gyroazimuth "φ g" are supplied to the input of the node (ZP) via the "φ" channel, and the processing voltage "ξ p" via the "ξ" channel.
10. "CAPTURE PREPARATION" MODE.
To reduce the review time, the search for a target in the radar station is carried out at high speed. In this regard, the station uses a two-stage target acquisition system, with storing the position of the target at the first detection, followed by returning the antenna to the memorized position and the secondary final target acquisition, after which its auto-tracking follows. Both preliminary and final target acquisition are carried out by the A3 node scheme.
When a target appears in the station search area, video pulses of the "direct target" from the asynchronous interference protection circuit of the synchronizer node (SI) begin to flow through the error signal amplifier (USO) of the node (AZ) to the detectors (D-1 and D-2) of the node (A3 ). When the missile reaches a range at which the signal-to-noise ratio is sufficient to trigger the cascade of the capture preparation relay (CRPC), the latter triggers the capture preparation relay (RPR) in the nodes (CS φ and DC ξ). The capture automaton (A3) cannot work in this case, because. it is unlocked by voltage from the circuit (APZ), which is applied only 0.3 sec after operation (APZ) (0.3 sec is the time required for the antenna to return to the point where the target was originally detected).
Simultaneously with the operation of the relay (RPZ):
· from node of storage (ZP) input signals "ξ p" and "φ n" are disconnected
The voltages that control the search are removed from the inputs of the nodes (PGS) and (UGA)
· the storage node (ZP) begins to issue stored signals to the inputs of the nodes (PGS) and (UGA).
To compensate for the error of the storage and gyro stabilization circuits, the swing voltage (f = 1.5 Hz) is applied to the inputs of the nodes (OSG) and (UGA) simultaneously with the stored voltages from the node (ZP), as a result of which, when the antenna returns to the memorized point, the beam swings with a frequency of 1.5 Hz and an amplitude of ± 3°.
As a result of the operation of the relay (RPZ) in the channels of the nodes (RS) and (RS), the outputs of the nodes (RS) are connected to the input of the antenna drives via the channels "φ" and "ξ" simultaneously with the signals from the OGM, as a result of which the drives begin to be controlled also an error signal of the angle tracking system. Due to this, when the target re-enters the antenna pattern, the tracking system retracts the antenna into the equisignal zone, facilitating the return to the memorized point, thus increasing the capture reliability.
11. CAPTURE MODE
After 0.4 seconds after the capture preparation relay is triggered, the blocking is released. As a result of this, when the target re-enters the antenna pattern, the capture relay cascade (CRC) is triggered, which causes:
· actuation of the capture relay (RC) in the nodes (US "φ" and US "ξ") that turn off the signals coming from the node (SGM). Antenna control system switches to automatic target tracking mode
actuation of the relay (RZ) in the UGA node. In the latter, the signal coming from the node (ZP) is turned off and the ground potential is connected. Under the influence of the appeared signal, the UGA system returns the antenna mirror to the zero position along the "ξ p" axis. Arising in this case, due to the withdrawal of the equisignal zone of the antenna from the target, the error signal is worked out by the SUD system, according to the main drives "φ" and "ξ". In order to avoid tracking failure, the return of the antenna to zero along the axis "ξ p" is carried out at a reduced speed. When the antenna mirror reaches the zero position along the axis "ξ p ". the mirror locking system is activated.
12. MODE "AUTOMATIC TRACKING"
From the output of the CO node from the video amplifier circuits (VUZ and VU4), the error signal with a frequency of 62.5 Hz, divided along the "φ" and "ξ" axes, enters through the nodes US "φ" and US "ξ" to phase detectors. The reference voltage "φ" and "ξ" are also fed to the phase detectors, which comes from the reference voltage trigger circuit (RTS "φ") of the KP-2 unit and the switching pulse shaping circuit (SΦPCM "P") of the UV-2 unit. From the phase detectors, the error signals are fed to the amplifiers (CS "φ" and CS "ξ") and further to the antenna drives. Under the influence of the incoming signal, the drive turns the antenna mirror in the direction of decreasing the error signal, thereby tracking the target.
The figure is located at the end of the entire text. The scheme is divided into three parts. Transitions of conclusions from one part to another are indicated by numbers.
BALTIC STATE TECHNICAL UNIVERSITY
_____________________________________________________________
Department of Radioelectronic Devices
RADAR HOMING HEAD
St. Petersburg
2. GENERAL INFORMATION ABOUT RLGS.
2.1 Purpose
The radar homing head is installed on the surface-to-air missile to ensure automatic target acquisition, its auto-tracking and the issuance of control signals to the autopilot (AP) and radio fuse (RB) at the final stage of the missile's flight.
2.2 Specifications
RLGS is characterized by the following basic performance data:
1. search area by direction:
Elevation ± 9°
2. search area review time 1.8 - 2.0 sec.
3. target acquisition time by angle 1.5 sec (no more)
4. Maximum angles of deviation of the search area:
In azimuth ± 50° (not less than)
Elevation ± 25° (not less than)
5. Maximum deviation angles of the equisignal zone:
In azimuth ± 60° (not less than)
Elevation ± 35° (not less than)
6. target capture range of the IL-28 aircraft type with the issuance of control signals to (AP) with a probability of not less than 0.5 -19 km, and with a probability of not less than 0.95 -16 km.
7 search zone in range 10 - 25 km
8. operating frequency range f ± 2.5%
9. average transmitter power 68W
10. RF pulse duration 0.9 ± 0.1 µs
11. RF pulse repetition period T ± 5%
12. sensitivity of receiving channels - 98 dB (not less)
13.power consumption from power sources:
From the mains 115 V 400 Hz 3200 W
Mains 36V 400Hz 500W
From the network 27 600 W
14. station weight - 245 kg.
3. PRINCIPLES OF OPERATION AND CONSTRUCTION OF RLGS
3.1 The principle of operation of the radar
RLGS is a radar station of the 3-centimeter range, operating in the mode of pulsed radiation. At the most general consideration, the radar station can be divided into two parts: - the actual radar part and the automatic part, which provides target acquisition, its automatic tracking in angle and range, and the issuance of control signals to the autopilot and radio fuse.
The radar part of the station works in the usual way. High-frequency electromagnetic oscillations generated by the magnetron in the form of very short pulses are emitted using a highly directional antenna, received by the same antenna, converted and amplified in the receiving device, pass further to the automatic part of the station - the target angle tracking system and the rangefinder.
The automatic part of the station consists of the following three functional systems:
1. antenna control systems that provide antenna control in all modes of operation of the radar station (in the "pointing" mode, in the "search" mode and in the "homing" mode, which in turn is divided into "capture" and "autotracking" modes)
2. distance measuring device
3. a calculator for control signals supplied to the autopilot and radio fuse of the rocket.
The antenna control system in the "auto-tracking" mode works according to the so-called differential method, in connection with which a special antenna is used in the station, consisting of a spheroidal mirror and 4 emitters placed at some distance in front of the mirror.
When the radar station operates on radiation, a single-lobe radiation pattern is formed with a maμmum coinciding with the axis of the antenna system. This is achieved due to the different lengths of the waveguides of the emitters - there is a hard phase shift between the oscillations of different emitters.
When working at reception, the radiation patterns of the emitters are shifted relative to the optical axis of the mirror and intersect at a level of 0.4.
The connection of the emitters with the transceiver is carried out through a waveguide path, in which there are two ferrite switches connected in series:
· Axes commutator (FKO), operating at a frequency of 125 Hz.
· Receiver switch (FKP), operating at a frequency of 62.5 Hz.
Ferrite switches of the axes switch the waveguide path in such a way that first all 4 emitters are connected to the transmitter, forming a single-lobe directivity pattern, and then to a two-channel receiver, then emitters that create two directivity patterns located in a vertical plane, then emitters that create two patterns orientation in the horizontal plane. From the outputs of the receivers, the signals enter the subtraction circuit, where, depending on the position of the target relative to the equi-signal direction formed by the intersection of the radiation patterns of a given pair of emitters, a difference signal is generated, the amplitude and polarity of which is determined by the position of the target in space (Fig. 1.3).
Synchronously with the ferrite axis switch in the radar station, the antenna control signal extraction circuit operates, with the help of which the antenna control signal is generated in azimuth and elevation.
The receiver commutator switches the inputs of the receiving channels at a frequency of 62.5 Hz. The switching of receiving channels is associated with the need to average their characteristics, since the differential method of target direction finding requires the complete identity of the parameters of both receiving channels. The RLGS rangefinder is a system with two electronic integrators. From the output of the first integrator, a voltage proportional to the speed of approach to the target is removed, from the output of the second integrator - a voltage proportional to the distance to the target. The range finder captures the nearest target in the range of 10-25 km with its subsequent auto-tracking up to a range of 300 meters. At a distance of 500 meters, a signal is emitted from the rangefinder, which serves to cock the radio fuse (RV).
The RLGS calculator is a computing device and serves to generate control signals issued by the RLGS to the autopilot (AP) and RV. A signal is sent to the AP, representing the projection of the vector of the absolute angular velocity of the target sighting beam on the transverse axes of the missile. These signals are used to control the missile's heading and pitch. A signal representing the projection of the velocity vector of the target's approach to the missile onto the polar direction of the target's sighting beam arrives at the RV from the computer.
The distinctive features of the radar station in comparison with other stations similar to it in terms of their tactical and technical data are:
1. the use of a long-focus antenna in a radar station, characterized by the fact that the beam is formed and deflected in it using the deflection of one rather light mirror, the deflection angle of which is half that of the beam deflection angle. In addition, there are no rotating high-frequency transitions in such an antenna, which simplifies its design.
2. use of a receiver with a linear-logarithmic amplitude characteristic, which provides an expansion of the dynamic range of the channel up to 80 dB and, thereby, makes it possible to find the source of active interference.
3. building a system of angular tracking by the differential method, which provides high noise immunity.
4. application in the station of the original two-circuit closed yaw compensation circuit, which provides a high degree of compensation for the rocket oscillations relative to the antenna beam.
5. constructive implementation of the station according to the so-called container principle, which is characterized by a number of advantages in terms of reducing the total weight, using the allotted volume, reducing interconnections, the possibility of using a centralized cooling system, etc.
3.2 Separate functional radar systems
RLGS can be divided into a number of separate functional systems, each of which solves a well-defined particular problem (or several more or less closely related particular problems) and each of which is to some extent designed as a separate technological and structural unit. There are four such functional systems in the RLGS:
3.2.1 Radar part of the RLGS
The radar part of the RLGS consists of:
the transmitter.
receiver.
high voltage rectifier.
the high frequency part of the antenna.
The radar part of the RLGS is intended:
· to generate high-frequency electromagnetic energy of a given frequency (f ± 2.5%) and a power of 60 W, which is radiated into space in the form of short pulses (0.9 ± 0.1 μs).
· for the subsequent reception of signals reflected from the target, their conversion into intermediate frequency signals (Fpch = 30 MHz), amplification (via 2 identical channels), detection and delivery to other radar systems.
3.2.2. Synchronizer
Synchronizer consists of:
Receiving and Synchronization Manipulation Unit (MPS-2).
· receiver switching unit (KP-2).
· Control unit for ferrite switches (UF-2).
selection and integration node (SI).
Error signal selection unit (CO)
· ultrasonic delay line (ULZ).
generation of synchronization pulses for launching individual circuits in the radar station and control pulses for the receiver, SI unit and rangefinder (MPS-2 unit)
Formation of impulses for controlling the ferrite switch of axes, the ferrite switch of the receiving channels and the reference voltage (UV-2 node)
Integration and summation of received signals, voltage regulation for AGC control, conversion of target video pulses and AGC into radio frequency signals (10 MHz) for their delay in the ULZ (SI node)
· isolation of the error signal necessary for the operation of the angular tracking system (CO node).
3.2.3. Rangefinder
The rangefinder consists of:
Time modulator node (EM).
time discriminator node (VD)
two integrators.
The purpose of this part of the RLGS is:
search, capture and tracking of the target in range with the issuance of signals of the range to the target and the speed of approach to the target
issuance of signal D-500 m
OGS is designed to capture and automatically track the target by its thermal radiation, measure the angular velocity of the line of sight of the missile - target and generate a control signal proportional to the angular velocity of the line of sight, including under the influence of a false thermal target (LTTs).
Structurally, the OGS consists of a coordinator 2 (Fig. 63) and an electronic unit 3. An additional element that formalizes the OGS is 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 is the cooling system 5. The refrigerant is liquefied gas obtained in the cooling system from gaseous nitrogen by throttling.
The block diagram of the optical homing head (Fig. 28) consists of the following coordinator and autopilot circuits.
The tracking coordinator (SC) performs 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 photodetectors FPok and FPvk, respectively, with rasters of a certain configuration radially located relative to the optical axis.
The lens, photodetectors, preamplifiers are fixed on the gyroscope rotor and rotate with it, and the optical axis of the lens coincides with the axis of proper rotation of the gyroscope rotor. The gyroscope rotor, the main mass of which is a permanent magnet, is installed in a gimbals, allowing it to deviate from the longitudinal axis of the OGS by a bearing angle in any direction relative to two mutually perpendicular axes. When the gyroscope rotor rotates, the space is surveyed within the field of view of the lens 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 to 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 to the target, the scattering spot shifts. When the gyroscope rotor rotates, the photoresistors are illuminated for the duration of the passage of the scattering spot 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 an increase in the mismatch for the selected raster shape, their duration decreases. The pulse repetition rate is equal to the rotation frequency of the photoresistor.
Rice. 28. Structural diagram of the optical homing head
The signals from the outputs of the photodetectors FPok and FPvk, respectively, arrive at the preamplifiers PUok and PUvk, which are connected by a common automatic gain control system AGC1, operating on a signal from 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 goes to the switching circuit (SP), designed to protect against LTC and background noise. LTC protection is based on different temperatures of radiation from a 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 one, and the signal from the LTC to the output of the SP does not pass.
In the SP, a throughput strobe is formed for the target; the signal selected for the SP from the target is fed to the selective amplifier and the amplitude detector. The amplitude detector (AD) selects a signal, the amplitude of the first harmonic of which depends on the angular mismatch between the optical axis of the lens and the direction to the target. Further, 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 that 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 (UC) is the correction windings and active resistances connected in series with them, the signals from which are fed to the AP.
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 to the target. Thus, the OGS is tracking the target.
At small distances to the target, the dimensions of the radiation from the target perceived by the OGS increase, which leads to a change in the characteristics of the pulse signals from the output of the photodetectors, which worsens the ability of the OGS to track the target. To exclude this phenomenon, the near-field circuit is provided in the electronic unit of the SC, which provides tracking of the energy center of the jet and nozzle.
The autopilot performs the following functions:
Filtering the signal from the SC to improve the quality of the missile control signal;
Formation of a signal to turn the missile at the initial section of the trajectory to automatically provide the necessary elevation and lead angles;
Converting the correction signal into a control signal at the missile's control frequency;
Formation of a control command on a steering drive operating in a 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 the signal from the push-pull power amplifier, the load of which is the windings of the electromagnets of the spool valve of the steering machine.
The signal of the correction amplifier passes through a synchronous filter and a dynamic limiter connected in series and is fed to the input of the adder ∑І. The signal from the bearing winding is fed to the FSUR circuit along the bearing. It is necessary at the initial section of the trajectory to reduce the time 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, whose frequency is equal to the rotational speed of the gyroscope rotor, is fed 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 rotational frequencies of the gyroscope and the rocket. Therefore, one of the components of the output signal of the phase detector is the signal at the rocket rotation frequency.
The output signal of the phase detector is fed to the filter, at the input of which it is added to the signal of the linearization generator in the adder ∑II. The filter suppresses the high-frequency components of the signal from the phase detector and reduces the non-linear distortion of the linearization generator signal. The output signal from the filter will be 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 is fed to the power amplifier, the load of which is the windings of the electromagnets of the spool valve of the steering machine.
The gyroscope caging system is designed to match 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 the deviation of the gyroscope axis from the longitudinal axis of the missile is a bearing winding, the longitudinal axis of which coincides with the longitudinal axis of the missile. In the case of deviation of the gyroscope axis 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 direction finding winding, the tilt winding located in the launch tube sensor unit is turned on. The EMF induced in the slope winding is proportional in magnitude to the angle between the sighting axis of the aiming device and the longitudinal axis of the rocket.
The difference signal from the slope winding and the direction finding winding, amplified in voltage and power in the tracking coordinator, enters the gyroscope correction windings. Under the influence of a moment from the side of the correction system, the gyroscope precesses in the direction of decreasing the angle of mismatch with the sighting axis of the sighting device and is locked in this position. The gyroscope is de-caged by the ARP when the OGS is switched to the tracking mode.
To maintain the speed of rotation of the gyroscope rotor within the required limits, a speed stabilization system is used.
Steering compartment
The steering compartment includes the rocket flight control equipment. In the body of the steering compartment there is a steering machine 2 (Fig. 29) with rudders 8, an onboard power source 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
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Rice. 29. Steering compartment: 1 - amplifier; 2 - steering machine; 3 - control engine; 4 - pressure accumulator; 5 - stabilizer-rectifier; 6 - turbogenerator; 7 - socket; 8 - rudders (plates); 9 - destabilizer; 10 - angular velocity sensor
Rice. 30. Steering machine:
1 - output ends of the coils; 2 - body; 3 - latch; 4 - clip; 5 - filter; 6 - rudders; 7 - stopper; 8 - rack; 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
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Steering machine designed for aerodynamic control of the rocket in flight. At the same time, the RM serves as a switchgear in the gas-dynamic control system of the rocket in the initial section of the trajectory, when the aerodynamic rudders are ineffective. It is a gas amplifier for control electrical signals generated by the OGS.
The steering machine consists of a holder 4 (Fig. 30), in the tides of which there is a working cylinder with a piston 19 and a fine filter 5. The housing 2 is pressed into the holder with a spool valve, consisting of a four-edged spool 15, two bushings 16 and anchors 18. Two coils 17 and 20 of electromagnets are placed in the housing. The holder has two eyes, in which on the bearings 9 there is a rack 8 with springs (spring) and with a leash 12 pressed onto it. In the tide of the cage between the lugs, a gas distribution sleeve 14 is placed, rigidly fixed with a latch 3 on the rack. The sleeve has a groove with cut-off edges for supplying gas coming from the PUD to channels B, C and nozzles 13.
The RM is powered by PAD gases, which are supplied through a pipe through a fine filter to the spool and from it through channels in the rings, the housing and the holder under the piston. Command signals from the OGS are fed in turn to the coils of the electromagnets RM. 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 stops against the cover. Moving, the piston drags the protrusion of the leash behind it and turns the leash and the rack, and with them the rudders, to the extreme position. At the same time, the gas distribution sleeve also rotates, while the cut-off edge opens the gas access 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 action of the force from the powder gases, and the movement of the spool begins earlier than the current rises 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 are 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 an impeller 3 is mounted, which is its drive.
The stabilizer-rectifier performs two functions:
Converts the alternating current voltage of the turbogenerator to the required values of constant voltages and maintains their stability with changes in the speed of rotation of the rotor of the turbogenerator and load current;
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.
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Rice. 31. Turbogenerator:
1 - stator; 2 - nozzle; 3 - impeller; 4 - rotor
BIP works as follows. Powder gases from the combustion of the PAD charge through the nozzle 2 are fed to the blades of the turbine 3 and cause it to rotate together with the rotor. In this case, a variable EMF is induced in the stator winding, which is fed to the input of the stabilizer-rectifier. From the output of the stabilizer-rectifier, a constant voltage is supplied to the OGS and the DUS amplifier. The voltage from the BIP is supplied to the electric igniters of the VZ and PUD after the rocket exits the tube and the RM rudders are opened.
Angular velocity sensor is designed to generate an electrical signal proportional to the angular velocity of the missile's oscillations relative to its transverse axes. This signal is used to dampen the angular oscillations of the rocket in flight, the CRS is a frame 1 consisting of two windings (Fig. 32), which is suspended on the semiaxes 2 in the center screws 3 with corundum thrust 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 picked up from the sensitive element of the CRS (frame) through flexible momentless extensions 8, soldered to the contacts 10 of the frame and contacts 9, electrically isolated from the housing.
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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 in its windings is not induced. In the presence of rocket oscillations about transverse axes, the frame moves in a magnetic field. In this case, the EMF induced in the windings of the frame is proportional to the angular velocity of the rocket 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 vector of the absolute angular velocity of the rocket.
Powder pressure accumulator it is intended for feeding with powder gases RM and BIP. PAD consists of housing 1 (Fig. 33), which is a combustion chamber, and filter 3, in which gas is cleaned from solid particles. The gas flow rate and the parameters of the internal ballistics are determined by the throttle opening 2. Inside the body are placed a powder charge 4 and an igniter 7, consisting of an electric igniter 8, a sample of 5 gunpowder and a pyrotechnic firecracker 6.
Rice. 34. Powder control engine:
7 - adapter; 3 - body; 3 - powder charge; 4 - weight of gunpowder; 5 - pyrotechnic firecracker; 6 - electric igniter; 7 - igniter
PAD works as follows. An electrical impulse from the electronic unit of the trigger mechanism is fed to an electric igniter that 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 cleaned in the filter, after which they enter the RM and the BIP turbogenerator.
Powder control engine designed for gas-dynamic control of the rocket in the initial part of the flight path. The PUD consists of a body 2 (Fig. 34), which is a combustion chamber, and an adapter 1. Inside the body are a powder charge 3 and an igniter 7, consisting of an electric igniter 6, a sample of 4 gunpowder and a pyrotechnic firecracker 5. Gas consumption and parameters of the internal ballistics are determined by the orifice in the adapter.
PUD works as follows. After the rocket leaves the launch tube and the RM rudders open, an electrical impulse from the cocking capacitor is fed to an electric igniter, which ignites a sample of gunpowder and a 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 rudders of the RM, create a control force that ensures the turn of the rocket.
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 cocking unit to the electric igniters VZ (EV1) and PUD, as well as for switching the positive output of the BIP to the VZ after the rocket leaves the tube and the RM rudders open.
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Rice. 35. Scheme of the cocking block:
1 - circuit breaker
The cocking unit located in the socket housing consists of capacitors C1 and C2 (Fig. 35), resistors R3 and R4 to remove residual voltage from the capacitors after checks or a failed start, resistors R1 and R2 to limit the current in the capacitor circuit and diode D1, designed for electrical decoupling of BIP and VZ circuits. Voltage is applied to the cocking unit after the PM trigger is moved to the position until it stops.
Destabilizer is designed to provide overloads, the required stability and create additional torque, in connection with which its plates are installed at an angle to the longitudinal axis of the rocket.
Warhead
The warhead is designed to destroy an air target or cause damage to it, leading to the impossibility of performing a combat mission.
The damaging factor of the warhead is the high-explosive action of the shock wave of the explosive products of the warhead and the remnants of the propellant fuel, as well as the fragmentation action of the 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 carrier compartment of the rocket and is made in the form of an integral connection.
The warhead itself (high-explosive fragmentation) is designed to create a given defeat field that acts on the target after receiving an initiating pulse from the EO. It consists of body 1 (Fig. 36), warhead 2, detonator 4, cuff 5 and tube 3, through which the wires from the air intake to the steering compartment of the rocket pass. There is a yoke L on the body, the hole of which includes a pipe stopper designed to fix the rocket in it.
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Rice. 36. Warhead:
Warhead - the warhead itself; VZ - fuse; VG - explosive generator: 1- case;
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-liquidation time has elapsed, as well as to transfer the detonation pulse from the charge of the warhead to the charge of the explosive generator.
The fuse of the electromechanical type has two stages of protection, which are removed in flight, which ensures the safety of the operation of the complex (start-up, maintenance, transportation and storage).
The fuse consists of a safety detonating device (PDU) (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 retarder, an initiating charge, a detonator cap and a fuse detonator.
The remote control serves to ensure safety in handling the fuse until it is cocked after the rocket is launched. It includes a pyrotechnic fuse, a swivel sleeve and a blocking stop.
The fuse detonator is used to detonate warheads. Target sensors GMD 1 and GMD2 provide triggering of the detonator cap when the missile hits the target, and the self-destruct mechanism - triggering of the detonator cap after the self-detonation time has elapsed in case of a miss. The tube ensures the transfer of impulse from the charge of the warhead to the charge of the explosive generator.
Explosive generator - designed to undermine the unburned part of the marching charge of remote control and create an additional field of destruction. It is a cup located in the body of the fuse with an explosive composition pressed into it.
The fuse and warhead when launching a rocket work as follows. When the rocket leaves the pipe, the rudders of the RM open, while the contacts of the socket breaker close and the voltage from the capacitor C1 of the cocking unit is supplied to the electric igniter EV1 of the fuse, from which the pyrotechnic fuse of the remote control and the pyrotechnic press fitting of the self-destruct mechanism are simultaneously ignited.
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Rice. 37. Structural diagram of the fuse
In flight, under the influence of axial acceleration from a running main engine, the blocking stopper of the remote control unit settles and does not prevent the turning of the rotary sleeve (the first stage of protection is removed). After 1-1.9 seconds after the launch of the rocket, the pyrotechnic fuse burns out, the spring turns the rotary sleeve into the firing position. In this case, the axis of the detonator cap 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 fitting of the self-destruction mechanism continues to burn, and the BIP feeds the capacitors C1 and C2 of the fuse on everything. throughout the flight.
When a missile hits the target at the moment the fuse passes through a metal barrier (when it breaks through) or along it (when it ricochets) in the winding of the main target sensor GMD1, under the influence of eddy currents induced in the metal barrier when the permanent magnet of the target sensor GMD1 moves, an electric pulse occurs. current. This pulse is applied to the EVZ electric igniter, from the beam of which the detonator cap is triggered, causing the fuse detonator to act. The fuse detonator initiates the warhead detonator, the operation of which causes the warhead and explosive in the fuse tube to rupture, which transmits the detonation to the explosive generator. In this case, the explosive generator is triggered and the residual fuel of the remote control (if any) is detonated.
When the missile hits the target, the backup target sensor GMD2 is also activated. Under the influence of the will of elastic deformations that occur when a missile meets an obstacle, the armature of the GMD2 target sensor breaks off, the magnetic circuit breaks, as a result of which an electric current pulse is induced in the winding, which is supplied to the EV2 electric igniter. From the beam of fire of the electric igniter EV2, a pyrotechnic retarder is ignited, the burning time of which exceeds the time required for the main target sensor GMD1 to approach the barrier. After the moderator burns out, the initiating charge is triggered, causing the detonator cap and warhead detonator to fire, the warhead and residual propellant fuel (if any) are detonated.
In the event of a missile miss on a target, after the pyrotechnic press-fitting of the self-destruction mechanism burns out, a detonator cap is triggered by a beam of fire, causing the detonator to act and detonate the warhead warhead with an explosive generator to self-destruct the missile.
Propulsion system
Solid propellant control is designed to ensure that the rocket leaves the tube, gives it the necessary angular velocity of rotation, accelerates to cruising speed and maintains this speed in flight.
The remote control consists of a starting engine, a dual-mode single-chamber sustainer engine and a delayed-action beam igniter.
The starting engine is designed to ensure the launch of the rocket from the tube and give it the required angular velocity of rotation. 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 cartridges (or monolith) freely installed in annular volume of the chamber. The starting charge igniter consists of a housing in which an electric igniter and a sample of gunpowder are placed. The disk and the diaphragm secure the charge during operation and transportation.
The starting engine is connected to the nozzle part of the propulsion engine. When docking the engines, the gas supply tube is put on the body of the beam igniter 7 (Fig. 39) of delayed action, located in the pre-nozzle volume of the propulsion engine. This connection ensures the transmission of the fire pulse to the beam igniter. The electrical connection of the igniter of the starting engine with the launch tube is carried out through the contact connection 9 (Fig. 38).
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Rice. 38. Starting engine:
1 - gas supply tube; 2 - disk; 3 - plug; 4 - nozzle block; 5 - diaphragm; 6 - starting charge; 7 - starting charge igniter; 8 - camera; 9 - contact
The nozzle block has seven (or six) nozzles located at an angle to the longitudinal axis of the rocket, which ensure the rotation of the rocket in the area of operation of the starting engine. 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 the acceleration of the rocket to cruising speed in the first mode and maintain this speed in flight in the second mode.
The sustainer engine consists of a chamber 3 (Fig. 39), a sustainer charge 4, a sustainer charge igniter 5, a nozzle block 6 and a delayed-action beam igniter 7. Bottom 1 is screwed into the front part of the chamber with seats for docking remote control and warhead. To obtain the required combustion modes, the charge is partially booked and reinforced with six wires 2.
1 - bottom; 2 - wires; 3 - camera; 4 - marching charge; 5 – marching charge igniter; 6 - nozzle block; 7 - beam delayed igniter; 8 - plug; A - threaded hole
Rice. 40. Delayed beam igniter: 1 - pyrotechnic moderator; 2 - body; 3 - bushing; 4 - transfer charge; 5 - deton. charge
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Rice. 41. Wing block:
1 - plate; 2 - front insert; 3 - body; 4 - axis; 5 - spring; 6 - stopper; 7 - screw; 8 - rear insert; B - ledge
To ensure the tightness of the chamber during operation and create the necessary pressure when the main charge is ignited, a plug 8 is installed on the nozzle block, which collapses and burns out from the propellant gases of the main engine. On the outer part of the nozzle block there are threaded holes A for attaching the wing block to the PS.
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 at a distance of at least 5.5 m. This protects the anti-aircraft gunner from exposure to the jet of propellant gases of the sustainer engine.
A delayed-action beam igniter consists of a body 2 (Fig. 40), in which a pyrotechnic retarder 1 is placed, a transfer charge 4 in a sleeve 3. On the other hand, a detonating charge 5 is pressed into the sleeve. , the detonating charge is ignited. The shock wave generated during detonation is transmitted through the wall of the sleeve and ignites the transfer charge, from which the pyrotechnic retarder is ignited. After a delay time from the pyrotechnic retarder, the main charge igniter ignites, which ignites the main charge.
DU works as follows. When an electrical impulse is applied to the electric igniter of the starting charge, the igniter is activated, 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 engine finishes its work in the pipe and lingers in it. From the powder gases formed in the chamber of the starting engine, a delayed-action beam igniter is triggered, which ignites the march charge igniter, from which the march charge is triggered at a safe distance for the anti-aircraft gunner. The reactive force created by the main engine accelerates the rocket to the main speed and maintains this speed in flight.
Wing block
The wing unit is designed for aerodynamic stabilization of the rocket in flight, creating lift in the presence of angles of attack and maintaining the required speed of rotation of the rocket on the trajectory.
The wing block consists of a body 3 (Fig. 41), four folding wings and a mechanism for their locking.
The folding wing consists of a plate 7, which is fastened with two screws 7 to the liners 2 and 8, put on the axis 4, placed in the hole in the body.
The locking mechanism consists of two stoppers 6 and a spring 5, with the help of which the stoppers are released and lock the wing when opened. After the spinning rocket takes off from the tube, under the action of centrifugal forces, the wings open. To maintain the required speed of rotation of the rocket in flight, the wings are deployed relative to the longitudinal axis of the wing unit at a certain angle.
The wing block is fixed with screws on the main engine nozzle block. There are four protrusions B on the body of the wing block for connecting it to the starting engine using an expandable connecting ring.
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Rice. 42. Pipe 9P39(9P39-1*)
1 - front cover; 2 and 11 - locks; 3 - block of sensors; 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 - lever of the heating mechanism; 18. 31 and 32 - springs; 19 38 - clamps; 20 - connector; 21 - rear rack; 22 - side connector mechanism; 23 - handle; 24 - front pillar; 25 - fairing; 26 - nozzles; 27 - board; 28 - pin contacts; 29 - guide pins; 30 - stopper; 33 - thrust; 34 - fork; 35 - body; 36 - button; 37 - eye; A and E - labels; B and M - holes; B - fly; G - rear sight; D - triangular mark; Zh - cutout; And - guides; K - bevel; L and U - surfaces; D - groove; Р and С – diameters; F - nests; W - board; Shch and E - gasket; Yu - overlay; I am a shock absorber;
*) Note:
1. Two variants of pipes can be in operation: 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
The homing head is an automatic device that is installed on a guided weapon in order to ensure high targeting accuracy.
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.
According to the principle of operation, the following homing heads are distinguished:
1) passive - receiving the energy radiated 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.
According to the type of energy received, the homing heads are divided into radar, optical, acoustic.
The acoustic homing head functions using audible sound and ultrasound. Its most effective use is in water, where sound waves decay 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 works using electromagnetic waves in the optical range. They are mounted on controlled means of destruction of ground, air and sea targets. Guidance is carried out by a source of infrared radiation or by the reflected energy of a laser beam. On guided means of destruction of ground targets, related to non-contrast, passive optical homing heads are used, which operate on the basis of an optical image of the terrain.
Radar homing heads work using electromagnetic waves in the radio range. Active, semi-active and passive radar heads are used on controlled means of destroying ground, air and sea targets-objects. On controlled means of destruction of non-contrasting ground targets, active homing heads are used, which operate on radio signals reflected from the terrain, or passive ones that operate on the radiothermal radiation of the terrain.
This text is an introductory piece. From the book Locksmith's Guide by Phillips Bill From the book Locksmith's Guide by Phillips Bill author Team of authorsDividing head A dividing head is a device used to set, clamp and periodically rotate or continuously rotate small workpieces processed on milling machines. In tool shops of machine-building enterprises
From the book Great Encyclopedia of Technology author Team of authorsTurret The turret is a special device in which various cutting tools are installed: drills, countersinks, reamers, taps, etc. The turret is an important component of turret lathes (automatic and
From the book Great Encyclopedia of Technology author Team of authorsHoming head A homing head is an automatic device that is installed on a guided weapon in order to ensure high targeting accuracy. The main parts of the homing head are: a coordinator with
From the book Great Soviet Encyclopedia (DE) of the author TSB From the book Great Soviet Encyclopedia (VI) of the author TSB From the book Great Soviet Encyclopedia (GO) of the author TSB From the book Great Soviet Encyclopedia (MA) of the author TSB From the book Great Soviet Encyclopedia (RA) of the author TSB From the book The Big Book of the Amateur Angler [with a colored insert] author Goryainov Alexey GeorgievichSinker head Today, this device is often referred to as a jig head. It resembles a large mormyshka with a fixing ring and a stopper for the bait. Spinning sinkers-heads serve mainly for horizontal wiring of soft baits and can vary in weight and