RU240095U1 - Integrated strapdown navigation device for unmanned aircraft systems - Google Patents

Abstract

The utility model relates to navigation, specifically to devices for unmanned aircraft systems (UAS). The solution comprises an inertial unit with switching and a power source, GNSS and IFRNS modules, a reference frequency generator, and a computing module with interfaces for receiving data from correctors: an airborne signal system, as well as auxiliary sensors such as a magnetometer, a digital image sensor (DISS), or a lidar. Furthermore, the claimed device is capable of integrating GNSS, DISS, and IFRNS data using an invariant extended Kalman filter, which takes into account the measurement errors of the auxiliary sensors. The claimed device is distinguished by the presence of an IFRNS module and the device's implementation of an algorithm for identifying and eliminating GNSS interference. The technical result consists of increased navigation accuracy in the presence of GNSS interference. 1 Fig.

Description

This utility model pertains to navigation technologies, specifically strapdown inertial navigation units, and can be used in unmanned aerial vehicles (UAVs). This device is designed to determine the three-dimensional coordinates, rotation angles, and speed of an aircraft in the absence of external navigation signals and under the influence of active electronic jamming.

Key requirements for navigation devices include autonomy, continuity, and high accuracy in determining the parameters of the unmanned vehicle's movement, with reliability that ensures safe control in conditions of possible interference.

Various integrated navigation devices used in aircraft are known. For example, patent RU2614192C1 describes a method for estimating inertial information errors and correcting them using Doppler velocity and drift measurement (DISS) data. This solution utilizes a classic discrete Kalman filtering procedure, and the correction is limited to data from a single correction device—the DISS. The system does not provide comprehensive integration of signals from multiple navigation aids and is not designed for stable operation in the absence of external navigation signals for extended periods.

Patent RU2646954C2 describes a method for correcting a strapdown inertial navigation system (SINS) for roll and pitch angles. This solution implements adaptive Kalman filtering with variable gain depending on flight parameters and turbulence conditions, using information from an airborne data system. A disadvantage of this method is its narrow functional focus—correcting only attitude parameters (roll and pitch), without comprehensive processing of signals from multiple different correctors and without the ability to estimate and compensate for sensor errors in a single model.

The closest technical solution to the claimed utility model is the navigation system patented by RU2640964C1, which comprises a strapdown inertial navigation system (SINS), an airborne signal system, satellite and radio navigation systems, a magnetometer, a laser rangefinder, and optoelectronic and astrocorrection systems. A distinctive feature of this system is the multiple redundancy of sensors, computers, and processing units, as well as the use of software and algorithmic modules to ensure high fault tolerance and survivability.

The disadvantage of this solution is that it focuses primarily on ensuring survivability and fault tolerance through architectural redundancy, while failing to disclose methods for the comprehensive, high-precision integration of signals from the SINS (strap-down inertial navigation system), GNSS (Global Navigation Satellite System), IPRNS (pulse-phase radio navigation systems) and auxiliary correctors in a single high-order Kalman filter. It does not provide for adaptive assessment and compensation of errors in sensors and correctors, nor does it provide for the extrapolation of errors in the absence of navigation signals.

The objective of this utility model is to create a device capable of compensating for the shortcomings of analogs and the prototype and to provide a technical result consisting of increasing the accuracy and noise immunity of the integrated UAS navigation system, including in conditions of active electronic interference in various operating modes: with correction by GNSS, by IFRNS and in conditions of completely autonomous operation.

The proposed utility model proposes an integrated strapdown navigation device for unmanned aerial systems (UAS), comprising an inertial unit (IU) with internal switching capabilities and a secondary power source, a GNSS receiver/computer module, an IFRNS receiver/computer module, a reference frequency generator, and a computing module for navigation algorithms with interfaces for receiving information from primary correctors: an airborne signal system (ASS), including a baroaltimeter. Auxiliary (optional) correctors may also be used: a three-axis magnetometer, a DISS or lidar, and a digital video camera.

The claimed device is an integrated strapdown navigation device for unmanned aircraft systems, comprising a speed measurement module, the first through third outputs of which are connected to the twentieth through twenty-second inputs of the information integration unit, an air signal system module, the first and second outputs of which are connected to the twenty-sixth and twenty-seventh inputs of the information integration unit, a reference frequency generator, the first output of which is connected to the sixth input of the satellite navigation information unit, an absolute angular velocity projection meter, the first through third outputs of which are connected to the same inputs of the failure detection, localization and compensation unit and the first through third inputs of the initial data computer, an apparent acceleration vector projection meter, the first through third outputs of which are connected to the fourth through sixth inputs of the navigation parameter computer, the fourth through sixth inputs of the initial data computer and the fourth through sixth inputs of the information integration unit, a satellite navigation information unit, the input of which is connected through an amplifier to an antenna, the first through eighth outputs of which are connected to the inputs of thirteenth to nineteenth and twenty-fifth, respectively, information integration units, the second to fourth outputs of which are also connected to the seventh to ninth inputs of the initial data computer, the ninth output is connected to the first input of the pulse-phase radio navigation system module, a failure detection, localization and compensation unit, the first to third outputs of which are connected to the first to third inputs of the navigation parameter computer and the first to third inputs of the information integration unit, an initial data computer, the first to sixth outputs of which are connected to the seventh to twelfth inputs of the information integration unit and the seventh to twelfth inputs of the navigation parameter computer, a navigation parameter computer, the first to ninth outputs of which are directly connected to the navigation parameter consumer equipment, and characterized in that it includes an information integration unit, the first to third outputs of which are connected to the first to third inputs of the satellite navigation information unit, the fourth to twelfth outputs are connected to the navigation parameter consumer equipment, configured to detect interference, as well as a module pulse-phase radio navigation systems, the first and second outputs of which are connected to the twenty-third and twenty-fourth inputs of the information integration unit, the third output of which is connected to the fifth input of the satellite navigation information unit, also the output of the second module of the pulse-phase radio navigation systems and the fourth input of the satellite navigation information unit are connected with the possibility of two-way information exchange.

The three-dimensional coordinates, attitude angles, and velocity of the aircraft are determined by processing signals from the SINS and correctors using an invariant extended Kalman filter (IEKF). The estimated vector in the Kalman filter contains error estimates for the accelerometers and gyroscopes, including offsets and scale factors, estimates for individual components of the correction systems, and error estimates for the SINS. The differences between the corresponding parameters of the SINS and the correction systems are processed to extract error estimates for the SINS and its inertial sensors. These estimates are then used to compensate for these errors.

Improving the accuracy and noise immunity in conditions of electronic interference (EI) is achieved through the identification of interference and its elimination, as well as the simultaneous assessment of the errors of the SINS, IFRNS, GNSS and optional correctors (DISS, lidar, three-axis magnetometer), determined by the error variances calculated during the implementation of the IRFK algorithms.

The algorithmic support and its implementing functional blocks have been modified. The Kalman invariant filter is capable of processing an extended set of measurement data (DISS, IFRNS), improving the accuracy of coordinate determination both in the presence of radio navigation signals and in autonomous mode. The device is equipped with a GNSS signal interference identification and rejection unit, capable of identifying characteristic interference features and automatically excluding unreliable GNSS data from the integration. A distinctive feature is the integration of an extended set of data from devices (SINS, GNSS, IFRNS, and additional information sources) with the ability to work together through a specially implemented invariant filter, ensuring comprehensive error assessment and compensation, as well as stable system operation in the absence of external signals. The implemented constructive and algorithmic interaction of the SINS with IFRNS, GNSS receivers, and auxiliary correction units ensures increased accuracy and stability of the system by expanding the set of measurements taken into account and automatically rejecting distorted signals.

The device can be explained by the following drawing:

Fig. 1 shows a block diagram of the proposed utility model. In accordance with Fig. 1, the device comprises a satellite navigation information unit (SNI) 1 connected to an antenna 2, inputs of an information integration unit (IFU) 4, an initial data calculator (IDC) 3 and an input of a pulse-phase radio navigation system module 9, and a group of inputs of the satellite navigation information unit 1 is connected to the outputs of the following units: a reference frequency generator 12 (RFG), a pulse-phase radio navigation system module 9 (PPRSM), and an information integration unit 4. In addition, the initial data computer 3 has inputs connected to the outputs of the apparent acceleration vector projection meter 6 (AAVPM) and the absolute angular velocity projection meter 5 (AAVPM), and the information integration unit 4 has inputs connected to the outputs of the velocity measurement module (VMM), the pulse-phase radio navigation system module 9, the failure detection, localization and compensation unit 8 (FLU), the apparent acceleration vector projection meter 6, and the initial data computer 3; the output group of the information integration unit 4 is directly connected to the system. The navigation parameter computer 7 (NPC) has a group of inputs connected to the initial data computer 3, the failure detection, localization and compensation unit 8, the apparent acceleration vector projection meter 6, and a group of outputs directly connected to the system. In addition to the above, the block 8 for detecting, localizing and compensating for failure has a group of inputs connected to the absolute angular velocity projection meter 6.

The proposed model works as follows. Signals from gyroscopes , , proportional to the projections of the absolute angular velocity vector, and signals from accelerometers , , , proportional to the projections of the absolute acceleration vector, are fed to the initial data computer. In addition, the gyroscope signals , , are sent to the failure detection, localization and compensation unit (FDU), where, as a result of processing, the following are generated: , , , which have the meaning of reliable values of the projections of the absolute angular velocity vector. The initial data calculator (IDC) generates values , , , proportional to the values of latitude, longitude and altitude of the aircraft's location, as well as , , , proportional to the values of the aircraft's heading, roll, and pitch angles. The information integration unit receives the values , , , , , from the VND, , , from BOLKO, , , – accelerometer signals, (latitude), (longitude), (height), (northern component of speed), (eastern component of speed), from the satellite navigation information block, And from the air signal system, – latitude and longitude from the pulse-phase radio navigation system module and , , – projections of velocities from the velocity measurement module.

The algorithm for detecting interference created by electronic warfare systems consists of tracking discrepancies and their increments, which are used to detect the moments of interference occurrence and removal.

Variable values (latitude), (longitude), (height), (northern component of speed), (eastern component of speed), (vertical component of speed), come from the satellite navigation information unit, and (latitude), (longitude), (height), (northern component of speed), (eastern component of speed), (vertical component of velocity) – from the SINS.

Next, at step k, checks for each variable from the discrepancy is calculated:

And the increment of the residual relative to the previous step:

The main condition under which the program stops and resumes integration using SNA information:

If And , Where – the limiting value of the increment of residuals, and – variance of the variable , then this section is the beginning of the impact of the electronic warfare systems (or its end), and in this case the device stops (resumes) integration using GNSS information.

There are also two scenarios for starting to use GNSS after a long period of non-use:

1. If GNSS data are not used in the integration and at the same time for all variables fair And , and also for at least one variable fair , Where – critical value of the variance of the variable , then the program begins integration using GNSS information.

2. If GNSS data are not used in the integration and at the same time for all variables fair And , and also for at least one pair from fair , , Where And – critical values of the dispersion of variables, then the program begins integration using GNSS information.

During periods when correction using GNSS data is not possible, the IFRNS data is used as an independent source of aircraft position data. Calculated residuals And are passed to the Kalman filter and participate in the aggregation algorithms, thereby maintaining the accuracy of the autonomous mode.

Example of specific implementation:

A device containing an inertial unit (IU) with internal switching means and a secondary power source:

1. Gyroscopes Fisoptica VG103F1 RAEL.402139.103.20 TP.

2. Quartz pendulum accelerometers Er-S1-e1.

3. Inertial assembly: EMFAS.731143.3.002.

4. ADC board based on the LTC2440 chip

GNSS receiving and computing module, 2K-363E-62 TSUI.468157.137-01, IFRNS module TSUI.468157.097-01, reference frequency generator GK85-TS-1-10M-2E-8/ET-E-12V-2, computing module for navigation algorithms with interfaces for receiving information based on the Vortex86DX processor.

Claims (1)

An integrated strapdown navigation device for unmanned aircraft systems, comprising: a velocity measurement module, the first through third outputs of which are connected to the twentieth through twenty-second inputs of the information integration unit; an air signal system module, the first and second outputs of which are connected to the twenty-sixth and twenty-seventh inputs of the information integration unit; a reference frequency generator, the first output of which is connected to the sixth input of the satellite navigation information unit, an absolute angular velocity projection meter, the first through third outputs of which are connected to the like inputs of the failure detection, localization and compensation unit and the first through third inputs of the initial data computer; an apparent acceleration vector projection meter, the first through third outputs of which are connected to the fourth through sixth inputs of the navigation parameter computer, the fourth through sixth inputs of the initial data computer and the fourth through sixth inputs of the information integration unit; a satellite navigation information unit, the input of which is connected to an antenna via an amplifier, the first through eighth outputs of which are connected to the thirteenth through nineteenth and twenty-fifth inputs, respectively, of the information integration unit, the second through fourth outputs are also connected to the seventh through ninth inputs of the initial data computer, the ninth output is connected to the first input of the pulse-phase radio navigation systems module; a failure detection, localization and compensation unit, the first through third outputs of which are connected to the first through third inputs of the navigation parameter computer and the first through third inputs of the information integration unit; an initial data computer, the first through sixth outputs of which are connected to the seventh through twelfth inputs of the information integration unit and the seventh through twelfth inputs of the navigation parameter computer; A navigation parameter calculator, the first through ninth outputs of which are directly connected to the navigation parameter consumer equipment, and characterized in that it includes an information integration unit, the first through third outputs of which are connected to the first through third inputs of the satellite navigation information unit, the fourth through twelfth outputs are connected to the navigation parameter consumer equipment, configured to detect interference, as well as a pulse-phase radio navigation system module, the first and second outputs of which are connected to the twenty-third and twenty-fourth inputs of the information integration unit, the third output of which is connected to the fifth input of the satellite navigation information unit, and the second output of the pulse-phase radio navigation system module and the fourth input of the satellite navigation information unit are also connected with the possibility of two-way information exchange.