Inertial control all-in-one machine and navigation control method thereof

Abstract

The invention provides an inertial control all-in-one machine and a navigation control method thereof.The inertial control all-in-one machine comprises an inertial navigation assembly, and the inertial navigation assembly collects inertial measurement signals of a carrier, obtains motion parameters of the carrier and sends the motion parameters to a missile-borne computer; the missile-borne computer receives the motion parameters of the inertial navigation assembly, communicates with a carrier and other subsystems on the carrier, and sends control signals to the electric comprehensive control combination and other subsystems; and the electric comprehensive control combination receives a control signal of the missile-borne computer and executes a corresponding control action. Through the integrated design of the inertial navigation system and the control system, the overall miniaturization of the missile system is realized, and the time period and the economic cost of the design of the missile system are reduced.

Description

Technical Field

[0001] This invention relates to the field of inertial navigation technology, and in particular to an integrated inertial control system and its navigation control method. Background Technology

[0002] Improving the precision strike capability of missile weapons is the most important offensive and defensive means in current and future defense applications. Inertial systems and control systems are core components of missile weapon systems and are key to achieving precision missile guidance.

[0003] The inertial system and control system in a missile weapon system are highly correlated, but due to the high degree of specialization and technical complexity of both, traditional missile systems typically design the two systems as two independent components. Modern missile weapon systems place higher demands on product size, weight, and cost. The traditional separate design of the inertial and control systems suffers from problems such as large size, heavy weight, complex structure, and numerous electrical interfaces, making it difficult to meet the current needs of missile weapon systems for miniaturization, low cost, and intelligence. Summary of the Invention

[0004] The purpose of this invention is to provide an integrated inertial control system that integrates inertial navigation components, onboard computer, and electrical integrated control modules to solve the problem of the separate design of inertial systems and control systems in the prior art, which makes integration difficult.

[0005] To achieve the above objectives, the technical solution adopted by the present invention is: an integrated inertial control system, the integrated inertial control system comprising:

[0006] An inertial navigation component is provided, which collects inertial measurement signals from the carrier, obtains the carrier's motion parameters, and sends the motion parameters to the onboard computer.

[0007] The onboard computer receives motion parameters from the inertial navigation component and communicates with the carrier aircraft and other subsystems on the carrier, sending control signals to the electrical integrated control assembly and other subsystems.

[0008] An electrical control unit receives control signals from the onboard computer and executes corresponding control actions.

[0009] Furthermore, the inertial control integrated machine also includes an external interface, and the inertial navigation component, the onboard computer, and the electrical integrated control assembly are respectively connected to the external interface for communication.

[0010] Furthermore, a surge suppression filter is connected in series between the external interface and the electrical control assembly.

[0011] Furthermore, the inertial navigation component includes a fiber optic gyroscope, at least one accelerometer, and a data processing circuit.

[0012] Furthermore, the inertial navigation component also includes a secondary power supply and a power-down sustaining circuit, with the secondary power supply electrically connected to the power-down sustaining circuit.

[0013] Furthermore, the data processing circuit includes a navigation circuit and an IF conversion circuit. The navigation circuit is communicatively connected to the fiber optic gyroscope, and the IF conversion circuit is communicatively connected to both the accelerometer and the navigation circuit.

[0014] Furthermore, the navigation circuit includes an ARM chip.

[0015] The present invention also provides a navigation control method for the above-mentioned inertial control integrated machine, the method comprising the following steps:

[0016] S1: The inertial navigation component collects the inertial measurement signals of the carrier, obtains the motion parameters of the carrier, and sends the motion parameters to the onboard computer;

[0017] S2: Obtain the relative motion information between the carrier and the target through an external device, and send the relative motion information to the onboard computer;

[0018] S3: The onboard computer calculates the rudder deflection angle of the carrier based on the motion parameters and the relative motion information, and controls the flight attitude of the carrier.

[0019] Before acquiring the inertial measurement signal, the inertial navigation component uses information provided by the main inertial navigation system on the carrier to complete the initial transfer alignment of the sub-inertial navigation system in the carrier.

[0020] The navigation control method of the present invention does not rely on guidance commands as a transition. Instead, it directly calculates the rudder deflection angle of the carrier based on the attitude information of the carrier during flight and the relative motion information between the carrier and the target, thereby controlling the missile flight. This helps to improve control accuracy and enhance the stability and reliability of the system.

[0021] Furthermore, the initial transfer alignment includes establishing an observation model based on velocity and attitude matching, wherein the observation model is:

[0022] ,

[0023] in, It is a velocity observation. It is an attitude observation measurement.

[0024] Furthermore, in the observation model: the include ,in, The velocity output of the main inertial navigation system. For the velocity output of the sub-inertial navigation system;

[0025] ,in, The attitude error of the sub-inertial navigation system; The attitude error of the main inertial navigation system. This is the attitude matrix output by the sub-inertial navigation system; The installation error angle between the main inertial navigation system and the sub-inertial navigation system.

[0026] This invention employs a speed + attitude matching method, balancing alignment accuracy and convergence speed, enabling rapid alignment transfer and meeting the requirements of a small inertial control system. Compared with existing technologies, the advantages of this invention are: the miniaturized inertial control system integrates inertial navigation components, onboard computer, and electrical control system in its structure, and achieves coordinated control and information interaction among the three, thus realizing the integrated design of the inertial system and control system in terms of structure and function; the navigation circuit adopts an ARM chip solution, reducing the interfaces between different products and the interface circuits required for level conversion, reducing the overall hardware quantity by nearly 30%; the onboard computer does not rely on guidance commands as a transition, but directly calculates the rudder deflection angle of the carrier based on the attitude information during the carrier's flight and the relative motion information between the carrier and the target, thereby controlling the missile's flight, which helps to improve control accuracy and enhance the stability and reliability of the system. Attached Figure Description

[0027] To more clearly illustrate the technical method of the present invention, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are only one embodiment of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0028] Figure 1 This is an external structural diagram of the inertial control integrated machine according to an embodiment of the present invention;

[0029] Figure 2 This is an exploded view of the inertial control integrated machine according to an embodiment of the present invention;

[0030] Figure 3 These are top and side views of the inertial control integrated machine according to an embodiment of the present invention. Figure 3 (a) is a top view of the inertial control integrated machine according to an embodiment of the present invention. Figure 3 (b) A side view of the inertial control integrated machine according to an embodiment of the present invention;

[0031] Figure 4 This is a block diagram of the electrical principle of a traditional inertial navigation system;

[0032] Figure 5This is a system software functional block diagram of a traditional inertial navigation system and control system;

[0033] Figure 6 This is an electrical principle block diagram of the inertial control integrated machine according to an embodiment of the present invention;

[0034] Figure 7 This is a software functional block diagram of the inertial navigation component of the inertial control integrated machine according to an embodiment of the present invention;

[0035] Figure 8 This is a circuit functional block diagram of the onboard computer of the inertial control integrated machine according to an embodiment of the present invention;

[0036] Figure 9 This is an electrical principle block diagram of the electrical integrated control assembly of the inertial control system according to an embodiment of the present invention.

[0037] In the attached image:

[0038] 1. External connector; 2. Electrical control board; 3. Internal vibration damping block; 4. Internal vibration damper; 5. Accelerometer; 6. Platform assembly; 7. Data processing board; 8. Leak detection port cover; 9. Positioning pin; 10. Base assembly; 11. Power board; 12. Gyroscope mainboard; 13. Light source board; 14. Light source; 15. Fiber optic ring assembly; 16. Launch computer board; 17. Filter; 18. Top cover assembly. Detailed Implementation

[0039] The present invention will now be described in detail with reference to the accompanying drawings and embodiments. It should be noted that, unless otherwise specified, the embodiments and features described herein can be combined with each other. For ease of description, the terms "upper," "lower," "left," and "right" used below only indicate that they correspond to the upper, lower, left, and right directions in the accompanying drawings and do not limit the structure.

[0040] Figure 1 This is an external structural diagram of the inertial control integrated machine according to an embodiment of the present invention. Figure 2 This is an exploded view of the inertial control integrated machine according to an embodiment of the present invention. Figure 3 These are top and side views of the inertial control integrated machine according to an embodiment of the present invention. Figure 3 (a) is a top view of the inertial control integrated machine according to an embodiment of the present invention. Figure 3 (b) A side view of the inertial control integrated machine according to an embodiment of the present invention. Figure 2As shown, the integrated inertial control unit mainly consists of an upper cover component 18, a base component 10, and functional modules installed between the two. The functional modules include core components such as the platform assembly 6, the onboard computer board 16, the fiber optic ring assembly 15, the gyroscope motherboard 12, the accelerometer 5, the electrical integrated control board 2, the data processing board 7, and the power supply board 11. They also include auxiliary structures such as an internal vibration damper 4, an internal vibration damping block 3, a light source 14, a light source board 13, and a filter 17. An external connector 1 is located on one side of the upper cover component 18 for communication with external devices. The base component is equipped with positioning pins 9 and a leak detection port cover plate 8 for overall positioning and sealing maintenance. All the above components are positioned and fixed by positioning pins 9, fasteners, and other structures, and are assembled into a compact integrated inertial control device.

[0041] Figure 4 This is a block diagram of the electrical principle of a traditional inertial navigation system. The inertial navigation system is designed using a DSP chip, with FPGA assistance. Figure 5 This is a functional block diagram of the system software of a traditional inertial navigation system and control system. The inertial navigation system completes data acquisition, temperature compensation, error compensation, and navigation calculation, while the control system software, based on the projectile motion information calculated by the inertial navigation system, realizes projectile motion control according to guidance laws, control algorithms, etc.

[0042] The electrical principle of the inertial control system (ICS) of this invention will be described in detail below. The ICS of this invention employs integrated guidance and control technology, with a single product performing calculations for navigation, guidance, and attitude control equations, thereby controlling missile flight in real time.

[0043] Figure 6 This is an electrical schematic diagram of the inertial control integrated machine according to an embodiment of the present invention, as shown below. Figure 6 As shown, the integrated inertial control system includes an inertial navigation component, an onboard computer, and an electrical control assembly. The inertial navigation component collects inertial measurement signals from the carrier (e.g., a missile), obtains the carrier's motion parameters, and transmits these parameters to the onboard computer via an RS422 serial port. The onboard computer receives the motion parameters from the inertial navigation component and communicates with the carrier aircraft and other subsystems on the carrier, sending control signals to the electrical control assembly and other subsystems. The electrical control assembly receives the control signals from the onboard computer and executes corresponding control actions. The inertial navigation component, the onboard computer, and the electrical control assembly are all connected to external interfaces, and all three are connected to external devices through these interfaces.

[0044] The inertial navigation system comprises a platform assembly, an IF converter circuit (including temperature measurement), a navigation circuit, a secondary power supply, and a power-down sustaining circuit. The platform assembly integrates a three-axis integrated fiber optic gyroscope and three accelerometers for measuring the angular velocity and linear acceleration of the carrier. The secondary power supply provides 5V and ±15V power to the internal components of the integrated system, as well as 12V power to the servos. The power-down sustaining circuit ensures the integrated system continues to operate normally for 50ms after a power outage. The navigation circuit reads data transmitted by the three-axis integrated fiber optic gyroscope at a 2kHz frequency via an RS422 serial port, demodulates the received data, and obtains the three-axis angular velocity pulses. The navigation circuit, in conjunction with and controlling the IF converter circuit, acquires the output signals from the three accelerometers to obtain the three-axis acceleration pulses. It performs temperature compensation, error compensation, and navigation calculations to obtain the carrier's position, velocity, attitude, and other motion parameters in the three axes of the ground coordinate system, and transmits these parameters to the onboard computer via the RS422 serial port.

[0045] The onboard computer receives measurement data from the inertial navigation system and calculates the carrier's attitude, position, and velocity information. It provides multiple communication interfaces, including one RS422 interface, a dual-redundant 1553B bus, seven RS422 interfaces, and an I / O detection channel, for information exchange with external devices. Specifically, the onboard computer communicates with the carrier aircraft via the A and B dual-redundant 1553B buses, receiving commands and data from the carrier aircraft and reporting various missile information and statuses as required. It assists in functions such as inertial navigation alignment, parameter setting, and launch control. Based on guidance laws and flight control algorithms, it performs calculations and communicates with other subsystems on the missile, completing timing organization, measurement, calculation, and output of various control signals. Simultaneously, it outputs ignition control signals and monitors the onboard power status.

[0046] The electrical control unit (ECU) activates the battery, switches power supplies, and drives the ignition based on control signals from the onboard computer. The ECU receives an external 28V power input and uses a surge suppression filter to filter and protect the input power, while also meeting the system's requirements for reverse connection protection, filtering, and spike suppression. The ECU provides multiple ignition control channels and external power supply channels to drive external devices and provide them with operating power. Simultaneously, the ECU communicates with the onboard computer through corresponding interfaces to monitor and control ignition control signals and the status of the external power supply.

[0047] The electrical connections and working principles of each module will be described in detail below.

[0048] Figure 7This is a software functional block diagram of the inertial navigation component of the inertial control integrated machine according to an embodiment of the present invention. The navigation circuit adopts an ARM chip solution. Specifically, the ARM chip is connected to an SDRAM module to cache a large amount of intermediate data generated during the navigation calculation process; a temperature-compensated crystal oscillator provides a high-precision clock signal to the ARM chip to ensure the timing synchronization of the entire circuit; and LED indicators are used to output the working status of the circuit, realizing visual monitoring of the circuit operation.

[0049] The raw analog signal from the accelerometer is first conditioned and converted by the IF circuit, then transmitted to the ARM chip via the IF circuit's acquisition signal line. Simultaneously, the ARM chip sends control commands to the IF circuit via its control signal line, controlling the timing of the signal acquisition process. The gyroscope's digital signal communicates bidirectionally with the ARM chip via an RS422 differential transceiver module. The ARM chip receives the angular velocity measurement data output by the gyroscope and sends trigger control signals to the gyroscope.

[0050] The external communication links employ an isolation design to enhance anti-interference capabilities. The navigation circuit communicates bidirectionally with both the onboard computer circuit and external test equipment via isolation circuits and RS422 differential transceiver modules: the communication link with the computer circuit is used to transmit navigation calculation results and control commands; the communication link with external test equipment is used for system debugging, performance verification, and data calibration.

[0051] The navigation circuit uses an ARM chip solution, which offers lower cost and power consumption compared to a DSP chip solution, and eliminates the need for an FPGA. Table 1 below compares the use of DSP chips and ARM chips in the navigation circuit. Furthermore, the ARM chip integrates 2MB of FLASH and 512KB of SRAM, reducing the need for interfaces between different products and level conversion circuits, thus reducing the overall hardware quantity by nearly 30%. ARM utilizes the SIMD instruction set to deeply optimize matrix operations, FFT, and filtering algorithms in the navigation algorithm through NEON intrinsics or assembly language, enabling data-level parallelism.

[0052] Table 1 Comparison of DSP Chips and ARM Chips

[0053]

[0054] Figure 8This is a circuit functional block diagram of the onboard computer of the inertial control system according to an embodiment of the present invention. The processor uses an HD-S3370, integrating a DSP and an FPGA, which achieve high-speed data interaction through an EMIF interface. The DSP is responsible for navigation calculation, guidance law calculation, and flight control algorithm implementation, while the FPGA assists in logic control, program loading, and timing management tasks, and communicates with the flash memory through an SPI interface. An externally connected SDRAM (32M×16Bit) is used to cache intermediate data, and a DC / DC power supply module supplies power to the internal circuits.

[0055] The onboard computer receives measurement data from the inertial navigation system via RS422 and outputs a 5ms synchronization signal to ensure timing alignment. It communicates with the carrier aircraft via a 1553B bus and interacts with external devices through multiple RS422 interfaces. These external devices include monitoring equipment, telemetry, data links, fuzes, servos, altimeters, seekers, and debugging equipment. The onboard computer outputs seven ignition control signals and one power control signal to the electrical control assembly, which receives and drives the corresponding ignition actuators and power switching devices. The electrical control assembly outputs thermal battery detection signals, DC power supply status detection signals, and deployment permission detection signals to the onboard computer. Simultaneously, the onboard computer receives status signals from external devices collected by the electrical control assembly through 16 I / O detection channels, enabling real-time monitoring of the missile system's status.

[0056] Figure 9 This is an electrical principle block diagram of the electrical integrated control unit of the inertial control system according to an embodiment of the present invention. The electrical integrated control unit receives two 28V DC main power supplies, two thermal battery power supplies, and a launch permission signal through an external interface. The 28V DC power supply 1, after passing through a surge suppression filter and a power conversion relay, is integrated with other input power supplies through battery grid-connection logic to output three system power supplies for external devices. A power failure maintenance circuit maintains power supply to critical circuits when the main power supply is interrupted. At the same time, it has built-in thermal battery detection, DC power supply detection, and launch permission signal detection modules to monitor the power status and launch conditions in real time.

[0057] The electrical integrated control unit receives 7 ignition control signals (including 6 ignition control signals and 1 ignition thermal battery activation control signal) and pyrotechnic short-circuit switching control signals from the onboard computer. It then uses 4 dual-channel ignition modules to achieve 6 ignition drive and 1 ignition thermal battery activation output, driving the pyrotechnic short-circuit switching device to power the missile system and complete the ignition execution function.

[0058] This invention also provides a navigation control method for an inertial control system. The method includes the following steps:

[0059] S1: The inertial navigation component collects the inertial measurement signals of the carrier, obtains the carrier's motion parameters, and sends the motion parameters to the onboard computer;

[0060] S2: Acquire relative motion information between the carrier and the target through external devices, and send the relative motion information to the onboard computer;

[0061] S3: The onboard computer calculates the rudder deflection angle of the carrier based on motion parameters and relative motion information, and controls the flight attitude of the carrier.

[0062] Before acquiring inertial measurement signals, the inertial navigation component uses information provided by the main inertial navigation system on the carrier to complete the initial transfer alignment of the sub-inertial navigation system in the carrier.

[0063] The onboard computer does not use guidance commands as a transition. Instead, it directly calculates the missile's control deflection angle from the missile's attitude information during flight and the relative motion information of the missile body and the target, thus controlling the missile's flight. This avoids the problems of post-flight alignment taking up flight time and affecting the initial trajectory accuracy, resulting in high control accuracy and stability.

[0064] Specifically, before entering navigation mode, the integrated system needs to use information provided by the main inertial navigation system (INS) on the carrier aircraft to complete the initial transfer alignment of the carrier's sub-INS. Transfer alignment consists of two processes: coarse alignment and fine alignment. The coarse alignment process uses the attitude, velocity, and position information given by the main INS as initial values ​​to begin attitude and navigation calculations. Fine alignment, on the other hand, uses a Kalman filter algorithm to estimate the real-time misalignment angle. Once the estimated value reaches the required accuracy, the real-time calculated attitude is corrected. The Kalman filter first updates the state variables over time using the state equation, then calculates theoretical observations using an observation model, and compares these with the actual measured velocity and attitude differences to obtain residuals. Finally, the predicted values ​​are corrected based on these residuals. Once the estimated value reaches the required accuracy, the estimated misalignment angle is used to correct the real-time calculated attitude, thus completing the fine alignment. Different matching forms of transfer alignment can be obtained depending on the measurement information in the Kalman filter algorithm. Compared to other matching forms, the velocity + attitude matching form is suitable for rapid transfer alignment.

[0065] The observation model employs a velocity-attitude matching approach, utilizing velocity and attitude error information between the master and sub-inertial navigation systems (INS) for calibration. Since the accuracy of the master INS is significantly higher than that of the sub-INS, the navigation information output by the master INS is used as the true value. Because the master INS is installed at the aircraft's center of gravity, where vibration response is relatively low, pre-filtering is typically performed on each set of velocity differences during calibration to eliminate errors caused by high-frequency vibrations. After lever arm effect compensation and ignoring the zero bias of the master INS accelerometer, the velocity error differential equation is obtained as follows:

[0066] ,

[0067] in, These represent the computational navigation coordinate systems of the main inertial navigation system and the sub-inertial navigation system, respectively. Let η be the zero bias of the accelerometer in the sub-inertial navigation system, and η be the uncorrelated noise. Calculate the specific force vector in the navigation coordinate system for the sub-inertial navigation system. This is the misalignment angle vector between the sub-inertial navigation system and the main inertial navigation system. This is the projection of the Earth's rotational angular velocity onto the sub-inertial navigation coordinate system. Calculate the projection of the angular velocity of the navigation coordinate system relative to the Earth into its own coordinate system for the sub-inertial navigation system. This is the velocity error vector between the sub-inertial navigation system and the main inertial navigation system in the sub-inertial navigation calculation navigation coordinate system.

[0068] The differential equation for the misalignment angle is:

[0069]

[0070] in, For the drift of the gyroscope with inertial guidance, For noise terms, Calculate the average angular velocity of the navigation coordinate system relative to the inertial space for the sub-inertial navigation system.

[0071] The state equation for velocity + attitude matching transfer alignment is established as follows:

[0072]

[0073] in, For the accelerometer zero-bias model, τ is the zero-biased decay term over time. a For the relevant time constant, White noise to drive zero bias variation. This is a gyroscope drift model. τ is the decay term of drift over time. g For the relevant time constant, White noise to drive drift changes. This indicates interference noise.

[0074] The velocity-attitude matching observation model uses the velocity error and attitude between master inertial navigation systems as observations: ,

[0075] in, It is a velocity observation. It is an attitude observation measurement.

[0076] Velocity Observation The composition is ,in, The velocity output of the main inertial navigation system. The velocity output of the sub-inertial navigation system.

[0077] Attitude observation It can be obtained by matching the attitude matrices of the sub-inertial navigation system and the main inertial navigation system, as shown below:

[0078] ,

[0079] in, The attitude error of the sub-inertial navigation system; The attitude error of the main inertial navigation system. This is the attitude matrix output by the sub-inertial navigation system; The installation error angle between the main inertial navigation system and the sub-inertial navigation system.

[0080] The state equation describes the time-varying characteristics of state variables such as the sub-inertial navigation misalignment angle, gyroscope drift, and accelerometer bias; the observation model establishes the mapping relationship between the velocity difference and attitude difference between the main inertial navigation system and the sub-inertial navigation system. Specifically, the misalignment angle between the main inertial navigation system and the sub-inertial navigation system... This can lead to errors in the sub-INS attitude matrix, which in turn affects the projection accuracy of the specific force in the navigation coordinate system, resulting in a difference in the velocity output between the main INS and the sub-INS. At the same time, by performing attitude maneuvers on the carrier aircraft, richer misalignment angle information can be included in the velocity difference and attitude difference, thereby improving the observability of the system and completing accurate estimation of the sub-INS misalignment angle, gyroscope drift, and accelerometer zero bias in a shorter time.

[0081] Experimental verification of the inertial control integrated machine of the present invention showed that its overall weight and volume were significantly reduced, which can meet the requirements of miniaturized missile systems. At the same time, it meets the requirements of missile systems for multiple interfaces and accuracy indicators, product acceptance and testing. Its main technical indicators are shown in Table 2.

[0082] Table 2 Main Technical Specifications of the Inertial Control Unit

[0083]

[0084] This invention presents a miniaturized, integrated inertial and control system that combines an inertial navigation system and a control system. By employing miniaturized and lightweight design methods, it integrates functional modules such as the inertial navigation system, onboard computer, and electrical system. The integrated inertial and control design significantly reduces the size and weight of the inertial navigation system, which aligns with the trend of miniaturization in missile systems and is suitable for small missile systems.

[0085] Specifically, the inertial control unit includes an inertial navigation component, an onboard computer, and an electrical integrated control assembly. The inertial navigation component includes a three-axis integrated fiber optic gyroscope, three quartz flexural accelerometers, and supporting data processing circuitry for measuring the attitude, position, and velocity of the carrier. The onboard computer performs calculations based on guidance laws and flight control algorithms and issues various control signals. The electrical integrated control assembly performs functions such as power filtering, power conversion, and ignition drive.

[0086] The miniaturized inertial control unit integrates inertial navigation, onboard computer, and electrical control system components. Through miniaturized design of each board and selection of components, it achieves structural fusion and seamless integration of the inertial and control systems. This integrated inertial control design avoids iterative design of multiple subsystems, thus reducing the design time and economic cost of the missile system. It also controls design costs, reduces reliance on hardware, and improves system reliability.

[0087] The inertial system and control system integration proposed in this invention is a novel system structure proposed in response to the traditional design methods of missiles. By using software and hardware integration, the product is designed to be low-cost and lightweight, ultimately achieving overall miniaturization and cost reduction of the missile system. It can be widely used in weapon systems such as rockets and guided bombs, and has broad application prospects.

[0088] The above embodiments should be understood as being used only to illustrate the present invention more clearly, and not to limit the scope of the present invention. After reading the present invention, any modifications of the present invention in various equivalent forms by those skilled in the art fall within the scope defined by the appended claims.

Claims

1. An integrated inertial control system, characterized in that, The inertial control integrated machine includes: An inertial navigation component is provided, which collects inertial measurement signals from the carrier, obtains the carrier's motion parameters, and sends the motion parameters to the onboard computer. The onboard computer receives motion parameters from the inertial navigation component and communicates with the carrier aircraft and other subsystems on the carrier, sending control signals to the electrical integrated control assembly and other subsystems. An electrical control unit receives control signals from the onboard computer and executes corresponding control actions.

2. The inertial control integrated machine according to claim 1, characterized in that, The inertial control integrated machine also includes an external interface, which is connected to external devices. The inertial navigation component, the onboard computer, and the electrical integrated control assembly are respectively communicatively connected to the external interface.

3. The inertial control integrated machine according to claim 2, characterized in that, A surge suppression filter is connected in series between the external interface and the electrical control assembly.

4. The inertial control integrated machine according to claim 1, characterized in that, The inertial navigation component includes a fiber optic gyroscope, at least one accelerometer, and data processing circuitry.

5. The inertial control integrated machine according to claim 4, characterized in that, The inertial navigation component also includes a secondary power supply and a power-down sustaining circuit, with the secondary power supply electrically connected to the power-down sustaining circuit.

6. The inertial control integrated machine according to claim 5, characterized in that, The data processing circuit includes a navigation circuit and an IF conversion circuit. The navigation circuit is communicatively connected to the fiber optic gyroscope, and the IF conversion circuit is communicatively connected to both the accelerometer and the navigation circuit.

7. The inertial control integrated machine according to claim 6, characterized in that, The navigation circuit includes an ARM chip.

8. A navigation control method for an inertial control integrated machine, characterized in that, The method includes the following steps: S1: The inertial navigation component collects the inertial measurement signals of the carrier, obtains the motion parameters of the carrier, and sends the motion parameters to the onboard computer; S2: Obtain the relative motion information between the carrier and the target through an external device, and send the relative motion information to the onboard computer; S3: The onboard computer calculates the rudder deflection angle of the carrier based on the motion parameters and the relative motion information, and controls the flight attitude of the carrier. Before acquiring the inertial measurement signal, the inertial navigation component uses information provided by the main inertial navigation system on the carrier to complete the initial transfer alignment of the sub-inertial navigation system in the carrier.

9. The navigation control method for an inertial control integrated machine according to claim 8, characterized in that, The initial transfer alignment includes establishing an observation model based on velocity and attitude matching, wherein the observation model is: ， in, It is a velocity observation. It is an attitude observation measurement.

10. The navigation control method for an inertial control integrated machine according to claim 9, characterized in that, In the observation model: include ,in, The velocity output of the main inertial navigation system. For the velocity output of the sub-inertial navigation system; ,in, The attitude error of the sub-inertial navigation system; The attitude error of the main inertial navigation system. This is the attitude matrix output by the sub-inertial navigation system; The installation error angle between the main inertial navigation system and the sub-inertial navigation system.

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