Guidance control device and method
By integrating the inertial measurement unit and the onboard computer to form a guidance and control device, the problems of large missile equipment size, complex power supply and high cost are solved, and the miniaturization and high-precision measurement and control of the missile are realized.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- Filing Date
- 2022-02-23
- Publication Date
- 2026-03-27
AI Technical Summary
Traditional inertial measurement units and missile-borne computers are independent units, resulting in large missile equipment size, complex power supply and high cost, which is not conducive to the miniaturization of missiles.
The inertial measurement unit and the onboard computer are integrated to form a guidance and control device, which includes an acceleration measurement device, an angular velocity measurement device, an information processing circuit, a power amplifier circuit, and an onboard computing device. This enables analog-to-digital conversion and compensation processing, and the integrated structure is compact.
This enabled the miniaturization of the missile, reduced equipment size and power supply complexity, lowered costs, and improved measurement accuracy and control efficiency.
Smart Images
Figure CN114518721B_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The present application relates to the technical field of guidance control, in particular to a guidance control device and method. BACKGROUND
[0002] The on-board equipment for attitude measurement of ground-to-air missile generally consists of an inertial measurement device and a missile-borne computer, the inertial measurement device can measure angular velocity and visual acceleration components along three axes of the missile body coordinate system, and the missile-borne computer can solve, stabilize control and adjust parameters of missile motion parameters and communicate data with other on-board equipment.
[0003] The traditional inertial measurement device and missile-borne computer are generally composed of two independent units, which are large in size, complex in power supply and high in cost. With the increasing range and flight time of the missile, the miniaturization and high precision of the on-board equipment are required, and the separate composition of the inertial measurement device and the missile-borne computer is not conducive to the development trend of missile miniaturization. SUMMARY
[0004] The purpose of the present application is to provide a guidance control device and method, which integrates the inertial measurement device and the missile-borne computer together, and is compact in structure, which is conducive to the development of missile miniaturization.
[0005] To achieve the above purpose, the present application provides the following scheme:
[0006] The present application provides a guidance control device, comprising:
[0007] an acceleration measurement device, an angular velocity measurement device, an information processing circuit, a power amplification circuit and a missile-borne computing device;
[0008] The information processing circuit is connected with the acceleration measurement device and the angular velocity measurement device respectively; the acceleration measurement device collects acceleration analog signals of the missile; the angular velocity measurement device collects angular velocity analog signals of the missile; the information processing circuit receives the acceleration analog signals and the angular velocity analog signals, performs analog-digital conversion operation on the acceleration analog signals and the angular velocity analog signals, obtains acceleration digital signals and angular velocity digital signals, and performs compensation processing on the acceleration digital signals and the angular velocity digital signals to obtain compensated data information;
[0009] The missile-borne computing device is connected with the information processing circuit and the power amplification circuit respectively, and the power amplification circuit is connected with a rudder; the missile-borne computing device receives the compensated data information transmitted by the information processing circuit, sends the compensated data information to the power amplification circuit; the power amplification circuit sends the compensated data information to the rudder after power amplification processing.
[0010] Optionally, the missile-borne computing device specifically comprises:
[0011] a digital signal processing circuit, a memory and a missile-borne computing controller;
[0012] the missile-borne computing controller is connected with the digital signal processing circuit and the information processing circuit respectively, and the memory is connected with the digital signal processing circuit; the digital signal processing circuit transmits the data stored in the memory to the missile-borne computing controller, and the missile-borne computing controller performs missile test operation by using the stored data.
[0013] Optionally, it further comprises:
[0014] a first power supply and a second power supply;
[0015] the first power supply is connected with the acceleration measuring device, the angular velocity measuring device and the information processing circuit respectively;
[0016] the second power supply is connected with the power amplification circuit and the missile-borne computing device respectively.
[0017] Optionally, the missile-borne computing controller is connected with a valve controller, the missile-borne computing controller transmits valve control instructions to the valve controller, and the missile-borne computing controller receives valve feedback signals sent by the valve controller.
[0018] Optionally, the missile-borne computing controller comprises a communication interface, an analog-digital conversion interface, a valve interface, a power supply interface, an input switch value interface and an output switch value interface;
[0019] the information processing circuit is connected with the missile-borne computing controller through the communication interface;
[0020] the power amplification circuit is connected with the missile-borne computing controller through the analog-digital conversion interface;
[0021] the valve controller is connected with the missile-borne computing controller through the valve interface;
[0022] the second power supply is connected with the missile-borne computing controller through the power supply interface;
[0023] the input switch value interface and the output switch value interface are used for connecting a missile cable network to the missile-borne computing controller.
[0024] Optionally, the angular velocity measuring device is an optical fiber gyroscope, and the acceleration measuring device is a quartz flexible accelerometer.
[0025] Optionally, the memory comprises a FLASH memory and an SRAM memory.
[0026] Optionally, the acceleration measuring device comprises an X-axis accelerometer, a Y-axis accelerometer and a Z-axis accelerometer; the X-axis accelerometer, the Y-axis accelerometer and the Z-axis accelerometer are connected with the information processing circuit.
[0027] The application further provides a guidance control method applied to the guidance control device, and the method comprises the following steps:
[0028] The information processing circuit acquires the acceleration analog signal of the missile collected by the acceleration measuring device and the angular velocity analog signal of the missile collected by the angular velocity measuring device;
[0029] The information processing circuit performs an analog-digital conversion operation on the acceleration analog signal and the angular velocity analog signal, obtains and compensates the acceleration digital signal and the angular velocity digital signal, and transmits the compensated data information to the missile-borne computing device;
[0030] The missile-borne computing device transmits the compensated data information to the power amplification circuit;
[0031] The power amplification circuit transmits the compensated data information to the rudder after power amplification processing.
[0032] Compared with the prior art, the application has the following beneficial effects:
[0033] The application provides a guidance control device and method, which comprises an acceleration measuring device, an angular velocity measuring device, an information processing circuit, a power amplification circuit and a missile-borne computing device; the information processing circuit is connected with the acceleration measuring device and the angular velocity measuring device; the information processing circuit receives the acceleration analog signal of the missile collected by the acceleration measuring device and the angular velocity analog signal of the missile collected by the angular velocity measuring device, performs an analog-digital conversion operation on the acceleration analog signal and the angular velocity analog signal, obtains and compensates the acceleration digital signal and the angular velocity digital signal, and obtains compensated data information; the information processing circuit of the application integrates the functions of digital-analog conversion and compensation processing; the missile-borne computing device is connected with the information processing circuit and the power amplification circuit; the power amplification circuit is connected with a rudder; the missile-borne computing device transmits the compensated data information to the power amplification circuit; and the power amplification circuit transmits the compensated data information to the rudder after power amplification processing. The acceleration measuring device, the angular velocity measuring device, the information processing circuit, the power amplification circuit and the missile-borne computing device are integrated together, so that the structure is compact and the missile can be miniaturized. BRIEF DESCRIPTION OF DRAWINGS
[0034] In order to make the technical solutions in the embodiments of the present application or the prior art clearer, the accompanying drawings needed in the embodiments will be briefly introduced below. Obviously, the accompanying drawings in the following description only need to be some embodiments of the present application, and other drawings can be obtained by those skilled in the art without any creative effort on the basis of these drawings.
[0035] Figure 1 The working principle diagram of the inertial measurement device in the embodiment of the present application;
[0036] Figure 2 The working principle diagram of the missile-borne computer in the embodiment of the present application;
[0037] Figure 3 The structural diagram of the guidance control device in the embodiment of the present application;
[0038] Figure 4 The flow chart of the guidance control method in the embodiment of the present application. DETAILED DESCRIPTION
[0039] The technical solutions in the embodiments of the present application will be described clearly and completely below with reference to the accompanying drawings in the embodiments of the present application. Obviously, the described embodiments are only some embodiments of the present application, but not all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without any creative effort fall within the protection scope of the present application.
[0040] The purpose of the present application is to provide a guidance control device and method, which integrates the inertial measurement device and the missile-borne computer together, so as to be compact in structure and be beneficial to the miniaturization development of the missile.
[0041] In order to make the above-mentioned purposes, features and advantages of the present application more obvious and easy to understand, the present application will be further described in detail below with reference to the accompanying drawings and specific embodiments.
[0042] The missile-borne equipment for the attitude measurement of the ground-to-air missile is generally composed of an inertial measurement device and a missile-borne computer. The inertial measurement device can correctly measure the angular velocity and the visual acceleration components along the three axes of the missile body coordinate system. The missile-borne computer can calculate, stabilize and control the missile motion parameters, and is responsible for the data communication with other equipment on the missile.
[0043] The working principle diagram of the inertial measurement device is as follows Figure 1As shown, the filter circuit of the inertial measurement device outputs the primary power after electromagnetic filtering, which is supplied to the AC power supply and the three-phase power supply, respectively, and each power supply is converted into the secondary power required in the inertial measurement device to supply power to the components in the inertial measurement device. The analog signals are input into the three accelerometers and the two gyroscopes, respectively, the two gyroscopes (X-axis dynamic tuning gyroscope and Y-axis Z-axis dynamic tuning gyroscope) are sensitive to the angular velocity information of the missile, and the three accelerometers (X-axis accelerometer, Y-axis accelerometer and Z-axis accelerometer) are sensitive to the acceleration information of each axis of the missile. After signal processing by the re-balance amplification circuit and the filter circuit, the signals are sent to the A / D conversion circuit for analog-digital conversion, and after error compensation by the error compensation software, the angular velocity, acceleration and temperature compensation information are output, wherein the digital quantity before compensation is 12 signals, the increment after compensation is 6 signals, the total quantity after compensation is 6 signals, and the temperature compensation is 1 signal. Finally, the signals are sent to the central information processor through the communication interface for solving and parameter adjustment by the flight control software.
[0044] The working principle diagram of the missile-borne computer is as shown in Figure 2 The missile-borne computer software is a monitoring software, and the missile-borne computer hardware includes a digital signal processor, a serial communication controller, a D / A converter, an A / D converter, a switching value input / output module, a switching conversion circuit and an external memory. The serial communication controller and the switching value input / output module are realized by an FPGA device, the D / A converter and the A / D converter are both 12 bits, the external memory includes a FLASH memory and an SRAM memory, and the SRAM of the digital signal processor is 34K*32bit. The digital signal processor is connected with the external memory, the D / A converter, the A / D converter, the serial communication controller and the switching value input / output module, the switching conversion circuit is connected with the switching value input / output module, the electrical system is connected with the switching conversion circuit, the inertial unit, the fuze, the uncontrolled instrument, the launch vehicle, the seeker and the telemetry / testing system are connected with the serial communication controller, the switching conversion circuit is connected with the fuze, the uncontrolled instrument is connected with the switching conversion circuit, and the rudder system is connected with the D / A converter, the A / D converter and the switching conversion circuit, respectively.
[0045] From the development direction of the current ground-to-air missile weapon system equipment, the missile range and flight time are getting longer and longer, the miniaturization and high precision of the missile equipment are getting higher and higher, and the demand for integrated inertial measurement and missile-borne computer is getting more and more urgent. The current inertial measurement device and missile-borne computer are independent units, which have large volume, complex power supply, large power consumption and high cost, and are not conducive to the development trend of missile miniaturization. Based on this, the application provides a guidance control device.
[0046] As Figure 3As shown, the guidance control device provided by the application comprises: an acceleration measuring device 1, an angular velocity measuring device 2, an information processing circuit 3, a power amplification circuit 4 and a missile-borne computing device 5. The acceleration measuring device 1 comprises an X-axis accelerometer, a Y-axis accelerometer and a Z-axis accelerometer, and the X-axis accelerometer, the Y-axis accelerometer and the Z-axis accelerometer are connected with the information processing circuit 3. The angular velocity measuring device 2 is an optical fiber gyroscope. The information processing circuit 3 contains an inertial measurement work software.
[0047] The information processing circuit 3 is connected with the acceleration measuring device 1 and the angular velocity measuring device 2 respectively; the acceleration measuring device 1 collects an acceleration analog signal of the missile; the angular velocity measuring device 2 collects an angular velocity analog signal of the missile; the information processing circuit 3 receives the acceleration analog signal and the angular velocity analog signal, performs an analog-digital conversion operation on the acceleration analog signal and the angular velocity analog signal, obtains and compensates the acceleration digital signal and the angular velocity digital signal, and obtains data information after compensation. The method for compensating the acceleration digital signal and the angular velocity digital signal is the same as the method for compensating by the error compensation software in the prior art. Figure 1
[0048] The missile-borne computing device 5 is connected with the information processing circuit 3 and the power amplification circuit 4 respectively, and the power amplification circuit 4 is connected with a rudder. The missile-borne computing device 5 receives the data information after compensation transmitted by the information processing circuit 3, generates a rudder command signal according to the data information after compensation, sends the rudder command signal to the power amplification circuit 4, and the power amplification circuit 4 sends the rudder command signal to the rudder 14 after power amplification processing. Optionally, the rudder 14 is a hydraulic rudder. The rudder 14 sends a rudder feedback signal to the power amplification circuit 4 after receiving the rudder command.
[0049] The optical fiber gyroscope 2 adopts a passive output mode and is triggered by a 0.5ms periodic timing interrupt signal generated by the information processing circuit 3 and is output through an RS-422 interface. The acceleration measuring device 1 receives a model flight signal, adopts an accelerometer as an acceleration sensor, and three accelerometers (an X-axis accelerometer, a Y-axis accelerometer and a Z-axis accelerometer) are orthogonally installed, the sensitive axes coincide with the missile body coordinate system, and respectively sense the projection components of the visual acceleration of the missile movement in the missile body coordinate system. The visual acceleration components sensed by the accelerometer are provided to the information processing circuit 3 in the form of an analog quantity.
[0050] The hardware circuit part of the information processing circuit 3 mainly comprises a DSP (Digital Signal Process) device, a ∑-△ type AD converter, peripheral auxiliary circuit and interface circuit. After the DSP is powered on and runs with the inertial measurement software, the information processing circuit 3 triggers the fiber-optic gyroscope output periodically with a 0.5ms timing signal. The DSP device simultaneously samples three channels of accelerometer analog signals and one channel of temperature signals. After two cycles of three channels of fiber-optic gyroscope, three channels of accelerometer and one channel of temperature digital signals are obtained, the inertial measurement software accumulates the data of two cycles of each channel of the fiber-optic gyroscope and the accelerometer to form 2.5ms increment compensation pre-digital quantity, and takes the data of the second cycle as total compensation pre-digital quantity. After the increment and total compensation pre-digital quantity are compensated by circuit zero, system error and temperature, the interrupt signal is triggered and transmitted to the missile-borne computing device 5 through the RS-422 serial interface according to the communication protocol.
[0051] The guidance control device further comprises a first power supply 6 and a second power supply 7; the first power supply 6 is connected with the acceleration measurement device 1, the angular velocity measurement device 2 and the information processing circuit 3 respectively; the second power supply 7 is connected with the power amplification circuit 4 and the missile-borne computing device 5 respectively. The power supply voltage of the first power supply 6 and the second power supply 7 is 27V, the first power supply 6 is a power supply component of LXD28L515, and the second power supply 7 is a power supply component of SWM-810. The power supply component of LXD28L515 serves as a secondary power supply to supply power to the accelerometer, the fiber-optic gyroscope and the information processing circuit 3 respectively. The power supply component of SWM-810 supplies power to the power amplification circuit 4 and the missile-borne computing device 5.
[0052] Optionally, the acceleration measurement device 1, the angular velocity measurement device 2, the information processing circuit 3, the power amplification circuit 4, the first power supply 6 and the second power supply 7 are combined as an inertial measurement power amplification combination 8.
[0053] The missile-borne computing device 5 specifically comprises a digital signal processing circuit 9, a storage and a missile-borne computing controller 10. The storage comprises a FLASH storage 11 and a SRAM storage 12. The missile-borne computing controller 10 is connected with the digital signal processing circuit 9 and the information processing circuit 3 respectively, and the storage is connected with the digital signal processing circuit 9. The digital signal processing circuit 9 transmits the data stored in the storage to the missile-borne computing controller 10, and the missile-borne computing controller 10 performs missile test operation by using the stored data. The digital signal processing circuit 9 is a DSP (V33) and contains measurement and control monitoring software.
[0054] The missile-borne computing controller 10 is connected with a valve controller 13. The missile-borne computing controller 10 transmits valve control instructions to the valve controller 13, and the missile-borne computing controller 10 receives the valve feedback signals sent by the valve controller 13.
[0055] The missile-borne computing controller 10 comprises a communication interface, an analog-digital conversion interface, a valve interface (FM interface), a power supply interface, an input switch quantity interface and an output switch quantity interface. The communication interface is a serial communication interface, and the communication interface comprises a GZ interface for communication with the information processing circuit 3, a YC interface for communication with a telemetry system, a YX interface for communication with a fuse, a YM interface for communication with a non-controlled instrument, a DY interface for communication with a guidance head, and a DD interface for communication with a launching vehicle. The analog-digital conversion interface comprises a DA interface (D / A conversion interface, digital-analog conversion interface) for sending rudder commands to the power amplifier circuit 4 and an AD interface (A / D conversion interface, analog-digital conversion interface) for receiving rudder feedback sent by the power amplifier circuit 4. The missile-borne cable network is connected to the input switch quantity interface and the output switch quantity interface respectively. The controllers of the input switch quantity interface, the output switch quantity interface, the A / D conversion interface, the D / A conversion interface and the serial communication interface circuit are realized by the FPGA, and the missile-borne computing controller 10 is realized by the FPGA.
[0056] The missile-borne computing device 5 performs serial communication with the ground equipment, collects input switch signals and driving output switch signals, has a software reset blocking function, loads programs from a FLASH program memory to an SRAM memory, uploads and downloads programs or data from the FLASH program memory by the ground equipment, uploads the instrument test software or commands by the ground test equipment, can complete the test and inspection of the hardware functional components, and can test the technical parameters of the rudder, the gas valve and the inertial measurement unit.
[0057] The inertial measurement device, the missile-borne computer and the rudder power amplifier are integrated in the application, and the measurement of the angular velocity of three axes and the acceleration of three axes, the output channel number of the four-way rudder power amplifier, the serial communication channel number of seven ways, the input switch quantity of four ways, the output switch quantity of nine ways, the D / A conversion of four ways and the A / D conversion of four ways can be completed. The guidance control device has small volume, light weight and mass not greater than 3.0 kg.
[0058] The high-precision fiber-optic gyroscope is used as the angular velocity sensitive element and the quartz flexible accelerometer is used as the acceleration sensitive element in the application, the measurement precision is high, the angular velocity measurement range is -550° / s~+550° / s, the angular velocity zero offset stability is not greater than 0.002° / s, the acceleration measurement range is -55g~+55g, and the acceleration offset stability is not greater than 2*10 -3 g.
[0059] The guidance control device can realize measurement of angular velocity and visual acceleration of a missile, output of a control instruction of a rudder, output of forwarding related information to a missile-borne device, and completion of a missile launching task by a missile launching vehicle.
[0060] As shown in Figure 4 The application further provides a guidance control method. Figure 3 The method is applied to the guidance control device shown in
[0061] In step 401, the information processing circuit obtains an acceleration analog signal of the missile collected by the acceleration measurement device and an angular velocity analog signal of the missile collected by the angular velocity measurement device.
[0062] In step 402, the information processing circuit performs an analog-digital conversion operation on the acceleration analog signal and the angular velocity analog signal, obtains an acceleration digital signal and an angular velocity digital signal, performs compensation processing on the acceleration digital signal and the angular velocity digital signal, and transmits the compensated data information to the missile-borne computing device.
[0063] In step 403, the missile-borne computing device transmits the compensated data information to the power amplification circuit.
[0064] In step 404, the power amplification circuit performs power amplification processing on the compensated data information and transmits the processed data information to the rudder.
[0065] It is to be understood that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting; it is not intended to exclude myriad other embodiments of the present application that other inventors can develop based on the description herein. It must be noted that as used herein, the singular forms "a", "an" and "the" include plural referents unless the context clearly dictates otherwise. The terms "includes" and "comprising," as well as derivatives thereof, mean that various embodiments include, but are not limited to, the listed material or step or steps. The terms "substantially," "essentially," "approximately," "about" or "perpendicular" and the like, are defined as being close to as understood by one of ordinary skill in the art, and in one non-limiting embodiment the close to is within 10%, in another embodiment within 5%, and in another embodiment within 1% and in yet another embodiment within 0.1%. Other definitions will be apparent to one of ordinary skill in the art.
[0066] Each of the embodiments in the present specification is described in a related manner, and the same or similar parts among the embodiments can be referred to each other. Each of the embodiments focuses on the difference from other embodiments. In particular, for the system embodiments, since they are basically similar to the method embodiments, the description is relatively simple, and the relevant parts can be referred to the part of the description of the method embodiments.
[0067] The above merely describes the embodiments of the present application and is not intended to limit the present application. The present application can have various modifications and changes for those skilled in the art. Any modification, equivalent replacement, improvement, etc. within the spirit and principle of the present application shall be included in the scope of claims of the present application.
Claims
1. A guidance control device, characterized by comprising: The application relates to a missile-borne computing device and a missile test system. The missile-borne computing device comprises an acceleration measuring device, an angular velocity measuring device, an information processing circuit, a power amplification circuit and a missile-borne computing device. The information processing circuit is connected with the acceleration measuring device and the angular velocity measuring device respectively; the acceleration measuring device collects acceleration analog signals of a missile; the angular velocity measuring device collects angular velocity analog signals of the missile; the information processing circuit receives the acceleration analog signals and the angular velocity analog signals, carries out analog-digital conversion operation on the acceleration analog signals and the angular velocity analog signals, obtains acceleration digital signals and angular velocity digital signals, and carries out compensation processing on the acceleration digital signals and the angular velocity digital signals to obtain compensated data information. The missile-borne computing device is connected with the information processing circuit and the power amplification circuit respectively, the power amplification circuit is connected with a rudder, the missile-borne computing device receives the compensated data information transmitted by the information processing circuit, sends the compensated data information to the power amplification circuit, and sends the compensated data information to the rudder after power amplification processing.
2. The guidance control device of claim 1, wherein The missile-borne computing device specifically comprises a digital signal processing circuit, a memory and a missile-borne computing controller. The missile-borne computing controller is connected with the digital signal processing circuit and the information processing circuit respectively, and the memory is connected with the digital signal processing circuit; the digital signal processing circuit transmits data stored in the memory to the missile-borne computing controller, and the missile-borne computing controller carries out missile test operation by using the stored data. The application further comprises a first power supply and a second power supply.
3. The guidance control device of claim 2, wherein The first power supply is connected with the acceleration measuring device, the angular velocity measuring device and the information processing circuit respectively. The second power supply is connected with the power amplification circuit and the missile-borne computing device respectively. The missile-borne computing controller is connected with a valve controller, transmits valve control instructions to the valve controller, and receives valve feedback signals sent by the valve controller. The missile-borne computing controller comprises a communication interface, an analog-digital conversion interface, a valve interface, a power supply interface, an input switch value interface and an output switch value interface.
4. The guidance control device of claim 3, wherein The information processing circuit is connected with the missile-borne computing controller through the communication interface.
5. The guidance control device of claim 4, wherein The power amplification circuit is connected with the missile-borne computing controller through the analog-digital conversion interface. The valve controller is connected with the missile-borne computing controller through the valve interface. The second power supply is connected with the missile-borne computing controller through the power supply interface. The input switch value interface and the output switch value interface are used for connecting a missile cable network to the missile-borne computing controller. The angular velocity measuring device is an optical fiber gyroscope, and the acceleration measuring device is a quartz flexible accelerometer. The memory comprises a FLASH memory and an SRAM memory.
6. The guidance control device of claim 1, wherein The acceleration measuring device comprises an X-axis accelerometer, a Y-axis accelerometer and a Z-axis accelerometer; the X-axis accelerometer, the Y-axis accelerometer and the Z-axis accelerometer are connected with the information processing circuit.
7. The guidance control device of claim 2, wherein 8. The guidance control device of claim 1, wherein 9. A guidance control method characterized by, The method is applied to the guidance control device as claimed in any one of claims 1-8, and the method comprises: The information processing circuit acquires an acceleration analog signal of the missile collected by the acceleration measuring device and an angular velocity analog signal of the missile collected by the angular velocity measuring device; The information processing circuit performs an analog-digital conversion operation on the acceleration analog signal and the angular velocity analog signal, obtains an acceleration digital signal and an angular velocity digital signal, and performs compensation processing on the acceleration digital signal and the angular velocity digital signal, obtains compensated data information, and transmits the compensated data information to the missile-borne computing device; The missile-borne computing device transmits the compensated data information to the power amplification circuit; The power amplification circuit transmits the compensated data information to the rudder after power amplification processing.
Citation Information
Patent Citations
Guidance control device
CN216979620U