A space-based measurement device for inertial navigation-assisted satellite navigation

Abstract

The present invention discloses a space-based measurement device for inertial navigation-assisted satellite navigation, belonging to the field of satellite navigation technology. The device includes a navigation measurement subsystem, a power interface module and a communication interface module; the power interface module receives an external electrical signal and provides electric energy to the navigation measurement subsystem; the communication interface module performs two-way communication with an external platform, provides the navigation measurement subsystem with the positioning solution information and original measurement information of the external platform, and provides the navigation result of the navigation measurement subsystem to the external platform; the navigation measurement subsystem obtains the GNSS satellite signal of the satellite, MEMS inertial navigation data and the positioning solution information and original measurement information provided by the external platform, performs combined navigation filtering and data fusion, and obtains the positioning information of the aircraft. The present invention is suitable for navigation systems and space-based measurement and control systems for high-dynamic flights, can improve the accuracy of high-dynamic receivers in complex dynamic environments, and has highly reliable navigation capabilities.

Description

Technical Field

[0001] The present invention relates to the field of satellite navigation technology, and in particular to a space-based measurement device for inertial navigation-assisted satellite navigation. Background Art

[0002] The global satellite navigation system can provide real-time position, speed and time information around the clock, and has become an indispensable source of navigation information for aircraft. Aircraft have complex dynamic characteristics during flight, and their navigation systems need to meet the requirements of high dynamics, high precision and high reliability. At the same time, there is an urgent need for high-precision measurement of flight trajectories during flight tests. Existing positioning and speed control methods cannot achieve high-precision, continuous and stable navigation in harsh environments such as high dynamics and high overload. It is necessary to study high-precision measurement devices that can adapt to complex environments. Summary of the invention

[0003] The present invention provides a space-based measurement device for inertial navigation-assisted satellite navigation, which aims to solve the problems of high-precision, continuous and stable positioning and speed measurement in harsh environments such as high dynamics and large overloads, improve the measurement accuracy of the space-based measurement system, and provide a demonstration for space-based measurement terminals for weapons and launch vehicle platforms.

[0004] A first aspect of the present invention provides a space-based measurement device for inertial navigation-assisted satellite navigation, the device comprising:

[0005] Navigation measurement subsystem, power interface module and communication interface module;

[0006] The power interface module receives an external electrical signal and provides electrical energy to the navigation measurement subsystem;

[0007] The communication interface module performs two-way communication with the external platform, provides the positioning solution information and original measurement information of the external platform to the navigation measurement subsystem, and provides the navigation result of the navigation measurement subsystem to the external platform;

[0008] The navigation measurement subsystem obtains the GNSS satellite signals, MEMS inertial navigation data, and positioning solution information and original measurement information provided by the external platform, performs combined navigation filtering and data fusion, and obtains the positioning information of the aircraft.

[0009] According to the device of the first aspect of the present invention, the navigation measurement subsystem includes a GNSS dual-frequency antenna, a radio frequency receiving unit, an integrated navigation module and a MEMS inertial navigation module;

[0010] The RF receiving unit includes a RF front end, a baseband correlator and a capture module; the RF front end converts the dual-frequency GNSS satellite signal received by the GNSS dual-frequency antenna into a digital intermediate frequency signal, the capture module obtains the digital intermediate frequency signal, determines the code phase and the carrier phase based on the digital intermediate frequency signal, and sends the code phase and the carrier phase to the integrated navigation module; the baseband correlator obtains the digital intermediate frequency signal;

[0011] The integrated navigation module includes a code and carrier digitally controlled oscillator, a code tracking loop, a carrier tracking loop, a factor graph pseudorange fusion processing module, a factor graph pseudorange rate fusion processing module, a pseudorange and pseudorange rate predictor and an integrated navigation filter; wherein:

[0012] The GNSS dual-frequency antenna receives a GNSS satellite signal, wherein the GNSS satellite signal is a dual-frequency GNSS satellite signal; the GNSS satellite signal is converted into a digital intermediate frequency signal through a radio frequency front end;

[0013] The capture module sends the digital intermediate frequency signal to a code and carrier digitally controlled oscillator; the code and carrier digitally controlled oscillator determines a code phase and a carrier phase based on an input, and generates a local satellite signal based on the code phase and the carrier phase;

[0014] The baseband correlator processes the local satellite signal and the received digital intermediate frequency signal to obtain an orthogonal amplitude component and an in-phase amplitude component, the orthogonal amplitude component is recorded as an I value, and the in-phase amplitude component is recorded as a Q value;

[0015] Inputting the I value into the code tracking loop to obtain pseudorange data, the pseudorange data including pseudorange, code tracking error and code loop carrier-to-noise ratio; inputting the Q value into the carrier tracking loop to obtain pseudorange rate data, the pseudorange rate data including pseudorange rate, carrier tracking error and carrier loop carrier-to-noise ratio; inputting the pseudorange data and pseudorange rate data into the integrated navigation filter; the pseudorange and pseudorange rate estimator corrects the ionospheric error and tropospheric error of the pseudorange, and inputs the corrected pseudorange into the integrated navigation filter;

[0016] The combined navigation filter receives the inertial navigation solution result obtained by the MEMS inertial navigation module performing inertial navigation solution on the aircraft, and performs combined filtering based on the corrected pseudorange, the pseudorange rate data and the inertial navigation solution result to obtain the positioning information of the aircraft; the inertial navigation solution and combined filtering module in the MEMS inertial navigation module outputs the filtered pseudorange and pseudorange rate;

[0017] The pseudorange and pseudorange rate estimator obtains a pseudorange estimation value and a pseudorange rate estimation value based on the filtered pseudorange and pseudorange rate; the factor graph pseudorange fusion processing module performs factor graph fusion processing on the pseudorange estimation value and the pseudorange data output by the code tracking loop to obtain a first fusion result; the factor graph pseudorange rate fusion processing module performs factor graph fusion processing on the pseudorange rate estimation value and the pseudorange rate data output by the carrier tracking loop to obtain a second fusion result; the first fusion result and the second fusion result are input into the code and carrier digital control oscillator;

[0018] The MEMS inertial navigation module includes an inertial measurement device and an inertial navigation solution and combined filtering module;

[0019] The inertial measurement device is used to obtain acceleration information and angular velocity information of the aircraft;

[0020] The inertial navigation solution and combined filtering module outputs the filtered pseudorange and pseudorange rate, and inputs the filtered pseudorange and pseudorange rate into the combined navigation filter.

[0021] According to the device of the first aspect of the present invention, the pseudorange and pseudorange rate estimator corrects the ionospheric error and tropospheric error of the pseudorange, wherein the corrected ionospheric error is:

[0022]

[0023] in, is the measured value after ionospheric delay correction, that is, the corrected ionospheric error, is the frequency of carrier L1, is the frequency of carrier L2, is the pseudorange measurement value of carrier L1, is the pseudorange measurement value of carrier L2.

[0024] According to the device of the first aspect of the present invention, the corrected tropospheric error is:

[0025]

[0026] in, , , They represent the corrected tropospheric error, dry delay, and wet delay, respectively. and is the mapping function.

[0027] According to the device of the first aspect of the present invention, a Kalman filter is used to perform factor graph fusion processing, and the state equation for fusion is:

[0028]

[0029] in, is the predicted state vector at time k, is the forecast error covariance matrix at time k, is the state transition matrix, is the predicted state vector at time k-1, is the forecast error covariance matrix at time k-1, is the control input matrix, is the control input vector, is the process noise covariance matrix, is the transpose.

[0030] According to the device of the first aspect of the present invention, the state update formula for fusion is:

[0031]

[0032] in, is the Kalman gain, is the observation matrix, is the observation vector, is the observation noise covariance matrix, is the observation matrix, is the updated state variable, is the Kalman gain, is the updated error covariance matrix, A is the unit matrix.

[0033] According to the device of the first aspect of the present invention, the GNSS dual-frequency antenna is used to acquire GNSS signals of multiple satellites.

[0034] The beneficial technical effects brought by the present invention include:

[0035] (1) The present invention adopts a GNSS dual-frequency receiver to eliminate ionospheric errors. The dual-frequency receiver can directly measure the ionospheric delay using dual-frequency measurement values ​​without the aid of any mathematical model, and can control the ionospheric error within a smaller range, thereby better achieving high-precision positioning.

[0036] (2) The scalar and vector tracking loops of the inertial navigation-aided satellite used in the present invention can resolve the contradiction between dynamic performance and noise suppression in a high dynamic environment. The pseudorange and pseudorange rate of the satellite are estimated by inertial navigation, combined with the pseudorange and pseudorange rate errors output by the satellite tracking loop filter, and the Doppler frequency is estimated based on a multi-source fusion algorithm to control the carrier and code numerically controlled oscillator frequency of the local signal in the tracking loop, thereby achieving the closure of the tracking loop. The present invention can improve the tracking capability of the receiver in a high dynamic environment.

[0037] (3) The present invention adopts a tight coupling method to enable the combined system to have better performance. The tight coupling method can more accurately predict the pseudorange and Doppler frequency shift of the satellite navigation signal, and the predicted value of the corrected inertial navigation can effectively detect the credibility of the actual measurement value of the satellite navigation. In addition, the positioning output of the inertial navigation subsystem is usually not related to the measurement value of the satellite navigation system. Therefore, the use of a tight coupling combination method can effectively realize the complementarity of the functional characteristics of the two navigation systems, thereby improving the positioning performance of the satellite navigation subsystem and the entire combined system.

[0038] (4) The present invention is applicable to navigation systems and space-based measurement and control systems for high-dynamic flights, can improve the accuracy of high-dynamic receivers in complex dynamic environments, and has highly reliable navigation capabilities.

[0039] (5) The present invention can be implemented based on a traditional inertial measurement unit and a traditional satellite navigation receiver hardware architecture, which can improve navigation performance and facilitate the transformation of traditional navigation systems. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] Figure 1 A schematic diagram of the structure of a space-based measurement device for inertial navigation-aided satellite navigation according to the present invention;

[0041] Figure 2 It is a detailed structural schematic diagram of the space-based measurement device for inertial navigation-assisted satellite navigation according to the present invention. DETAILED DESCRIPTION

[0042] The present invention is described in detail below in conjunction with the accompanying drawings and embodiments.

[0043] like Figure 1-Figure 2 As shown, the present invention proposes a space-based measurement device for inertial navigation-assisted satellite navigation, comprising:

[0044] Navigation measurement subsystem, power interface module and communication interface module;

[0045] The power interface module receives an external electrical signal and provides electrical energy to the navigation measurement subsystem;

[0046] The communication interface module performs two-way communication with the external platform, provides the positioning solution information and original measurement information of the external platform to the navigation measurement subsystem, and provides the navigation result of the navigation measurement subsystem to the external platform;

[0047] The navigation measurement subsystem obtains the GNSS satellite signals, MEMS inertial navigation data, and positioning solution information and original measurement information provided by the external platform, performs combined navigation filtering and data fusion, and obtains the positioning information of the aircraft.

[0048] The navigation and measurement subsystem includes a GNSS dual-frequency antenna, a radio frequency receiving unit, an integrated navigation module and a MEMS inertial navigation module.

[0049] The RF receiving unit includes a RF front end, a baseband correlator and a capture module; the RF front end converts the dual-frequency GNSS satellite signal received by the GNSS dual-frequency antenna into a digital intermediate frequency signal, the capture module obtains the digital intermediate frequency signal, determines the code phase and the carrier phase based on the digital intermediate frequency signal, and sends the code phase and the carrier phase to the integrated navigation module; the baseband correlator obtains the digital intermediate frequency signal;

[0050] The integrated navigation module includes a code and carrier digitally controlled oscillator, a code tracking loop, a carrier tracking loop, a factor graph pseudorange fusion processing module, a factor graph pseudorange rate fusion processing module, a pseudorange and pseudorange rate predictor and an integrated navigation filter; wherein:

[0051] The GNSS dual-frequency antenna receives a GNSS satellite signal, wherein the GNSS satellite signal is a dual-frequency GNSS satellite signal; the GNSS satellite signal is converted into a digital intermediate frequency signal through a radio frequency front end;

[0052] The capture module sends the digital intermediate frequency signal to a code and carrier digitally controlled oscillator; the code and carrier digitally controlled oscillator determines a code phase and a carrier phase based on an input, and generates a local satellite signal based on the code phase and the carrier phase;

[0053] The baseband correlator processes the local satellite signal and the received digital intermediate frequency signal to obtain an orthogonal amplitude component and an in-phase amplitude component, the orthogonal amplitude component is recorded as an I value, and the in-phase amplitude component is recorded as a Q value;

[0054] Inputting the I value into the code tracking loop to obtain pseudorange data, the pseudorange data including pseudorange, code tracking error and code loop carrier-to-noise ratio; inputting the Q value into the carrier tracking loop to obtain pseudorange rate data, the pseudorange rate data including pseudorange rate, carrier tracking error and carrier loop carrier-to-noise ratio; inputting the pseudorange data and pseudorange rate data into the integrated navigation filter; the pseudorange and pseudorange rate estimator corrects the ionospheric error and tropospheric error of the pseudorange, and inputs the corrected pseudorange into the integrated navigation filter;

[0055] The combined navigation filter receives the inertial navigation solution result obtained by the MEMS inertial navigation module performing inertial navigation solution on the aircraft, and performs combined filtering based on the corrected pseudorange, the pseudorange rate data and the inertial navigation solution result to obtain the positioning information of the aircraft; the inertial navigation solution and combined filtering module in the MEMS inertial navigation module outputs the filtered pseudorange and pseudorange rate;

[0056] The pseudorange and pseudorange rate estimator obtains a pseudorange estimation value and a pseudorange rate estimation value based on the filtered pseudorange and pseudorange rate; the factor graph pseudorange fusion processing module performs factor graph fusion processing on the pseudorange estimation value and the pseudorange data output by the code tracking loop to obtain a first fusion result; the factor graph pseudorange rate fusion processing module performs factor graph fusion processing on the pseudorange rate estimation value and the pseudorange rate data output by the carrier tracking loop to obtain a second fusion result; the first fusion result and the second fusion result are input into the code and carrier digitally controlled oscillator to realize closed-loop control of the code and carrier digitally controlled oscillator.

[0057] The GNSS dual-frequency antenna is used to acquire GNSS signals from multiple satellites.

[0058] When the baseband correlator is used for the first time, the baseband correlator processes the local satellite signal and the initial digital intermediate frequency signal to obtain an orthogonal amplitude component and an in-phase amplitude component, and inputs the orthogonal amplitude component and the in-phase amplitude component into the code tracking loop and the carrier tracking loop respectively to generate a pseudorange and a pseudorange rate respectively. The local satellite signal is determined based on an initial code phase and an initial carrier phase, and the initial code phase and the initial carrier phase are determined by a code and carrier numerically controlled oscillator based on the initial digital intermediate frequency signal. When used subsequently, the baseband correlator processes the local satellite signal and the subsequent digital intermediate frequency signal to obtain an orthogonal amplitude component and an in-phase amplitude component, and inputs the orthogonal amplitude component and the in-phase amplitude component into the code tracking loop and the carrier tracking loop respectively to generate a pseudorange data and a pseudorange rate data respectively. The local satellite signal is determined based on a subsequent code phase and a subsequent carrier phase, and the subsequent code phase is determined based on a subsequent digital intermediate frequency signal and the first fusion result, and the subsequent carrier phase is determined based on a subsequent digital intermediate frequency signal and the second fusion result.

[0059] The MEMS inertial navigation module includes an inertial measurement device and an inertial navigation solution and combined filtering module;

[0060] The inertial measurement device is used to obtain acceleration information and angular velocity information of the aircraft;

[0061] The inertial navigation solution and combined filtering module outputs the filtered pseudorange and pseudorange rate, and inputs the filtered pseudorange and pseudorange rate into the combined navigation filter.

[0062] The space-based measurement device for inertial navigation-assisted satellite navigation of the present invention is small in size and light in weight, thus meeting the demand for miniaturization; and adopts a wide power input module and a universal communication interface module, thus meeting the demand for universalization.

[0063] Furthermore, the pseudorange and pseudorange rate estimator corrects the ionospheric error and tropospheric error of the pseudorange, wherein the corrected ionospheric error is:

[0064]

[0065] in, is the measured value after ionospheric delay correction, that is, the corrected ionospheric error, is the frequency of carrier L1, is the frequency of carrier L2, is the pseudorange measurement value of carrier L1, is the pseudorange measurement value of carrier L2;

[0066] The corrected tropospheric error is:

[0067]

[0068] in, , , They represent the corrected tropospheric error, dry delay, and wet delay, respectively. and is the mapping function.

[0069] The Kalman filter is used for factor graph fusion processing, and the state equation for fusion is:

[0070]

[0071] in, is the predicted state vector at time k, is the forecast error covariance matrix at time k, is the state transition matrix, is the predicted state vector at time k-1, is the forecast error covariance matrix at time k-1, is the control input matrix, is the control input vector, is the process noise covariance matrix, is the transpose.

[0072] The state update formula for fusion is:

[0073]

[0074] in, is the Kalman gain, is the observation matrix, is the observation vector, is the observation noise covariance matrix, is the observation matrix, is the updated state variable, is the Kalman gain, is the updated error covariance matrix,A is the unit matrix.

[0075] The factor graph pseudorange fusion processing module and the factor graph pseudorange rate fusion processing module optimize the code and carrier numerically controlled oscillators based on the factor graph optimization principle. Each factor corresponds to an observation quantity, and its probability model is determined by the sensor probability model.

[0076] Furthermore, the attitude update module of the MEMS inertial navigation module performs error correction on the error-compensated result and outputs updated attitude data; the inertial navigation solution and combined filtering module solves the corrected data to obtain an inertial navigation-assisted second pseudorange rate estimate and a second pseudorange estimate.

[0077] The above specific embodiments only describe the design principle of the present invention. The shapes and names of the components in the description may be different and are not limited. Therefore, those skilled in the art in the field of the present invention may modify or replace the technical solutions recorded in the above embodiments; and these modifications and replacements do not deviate from the creative purpose and technical solutions of the present invention and should all fall within the protection scope of the present invention.

Claims

1. A space-based measurement device for inertial navigation-assisted satellite navigation, characterized in that: The device comprises: Navigation measurement subsystem, power interface module and communication interface module; The power interface module receives an external electrical signal and provides electrical energy to the navigation measurement subsystem; The communication interface module performs two-way communication with the external platform, provides the positioning solution information and original measurement information of the external platform to the navigation measurement subsystem, and provides the navigation result of the navigation measurement subsystem to the external platform; The navigation measurement subsystem acquires the GNSS satellite signals, MEMS inertial navigation data, and positioning solution information and original measurement information provided by the external platform, performs combined navigation filtering and data fusion, and obtains the positioning information of the aircraft; The navigation and measurement subsystem includes a GNSS dual-frequency antenna, a radio frequency receiving unit, an integrated navigation module and a MEMS inertial navigation module; The RF receiving unit includes a RF front end, a baseband correlator and a capture module; the RF front end converts the dual-frequency GNSS satellite signal received by the GNSS dual-frequency antenna into a digital intermediate frequency signal, the capture module obtains the digital intermediate frequency signal, determines the code phase and the carrier phase based on the digital intermediate frequency signal, and sends the code phase and the carrier phase to the integrated navigation module; the baseband correlator obtains the digital intermediate frequency signal; The integrated navigation module includes a code and carrier digitally controlled oscillator, a code tracking loop, a carrier tracking loop, a factor graph pseudorange fusion processing module, a factor graph pseudorange rate fusion processing module, a pseudorange and pseudorange rate predictor and an integrated navigation filter; wherein: The GNSS dual-frequency antenna receives a GNSS satellite signal, wherein the GNSS satellite signal is a dual-frequency GNSS satellite signal; the GNSS satellite signal is converted into a digital intermediate frequency signal through a radio frequency front end; The capture module sends the digital intermediate frequency signal to a code and carrier digitally controlled oscillator; the code and carrier digitally controlled oscillator determines a code phase and a carrier phase based on an input, and generates a local satellite signal based on the code phase and the carrier phase; The baseband correlator processes the local satellite signal and the received digital intermediate frequency signal to obtain an orthogonal amplitude component and an in-phase amplitude component, the orthogonal amplitude component is recorded as an I value, and the in-phase amplitude component is recorded as a Q value; Inputting the I value into the code tracking loop to obtain pseudorange data, the pseudorange data including pseudorange, code tracking error and code loop carrier-to-noise ratio; inputting the Q value into the carrier tracking loop to obtain pseudorange rate data, the pseudorange rate data including pseudorange rate, carrier tracking error and carrier loop carrier-to-noise ratio; inputting the pseudorange data and pseudorange rate data into the integrated navigation filter; the pseudorange and pseudorange rate estimator corrects the ionospheric error and tropospheric error of the pseudorange, and inputs the corrected pseudorange into the integrated navigation filter; The combined navigation filter receives the inertial navigation solution result obtained by the MEMS inertial navigation module performing inertial navigation solution on the aircraft, and performs combined filtering based on the corrected pseudorange, the pseudorange rate data and the inertial navigation solution result to obtain the positioning information of the aircraft; the inertial navigation solution and combined filtering module in the MEMS inertial navigation module outputs the filtered pseudorange and pseudorange rate; The pseudorange and pseudorange rate estimator obtains a pseudorange estimation value and a pseudorange rate estimation value based on the filtered pseudorange and pseudorange rate; the factor graph pseudorange fusion processing module performs factor graph fusion processing on the pseudorange estimation value and the pseudorange data output by the code tracking loop to obtain a first fusion result; the factor graph pseudorange rate fusion processing module performs factor graph fusion processing on the pseudorange rate estimation value and the pseudorange rate data output by the carrier tracking loop to obtain a second fusion result; the first fusion result and the second fusion result are input into the code and carrier numerically controlled oscillator; The MEMS inertial navigation module includes an inertial measurement device and an inertial navigation solution and combined filtering module; The inertial measurement device is used to obtain acceleration information and angular velocity information of the aircraft; The inertial navigation solution and combined filtering module outputs the filtered pseudorange and pseudorange rate, and inputs the filtered pseudorange and pseudorange rate into the combined navigation filter.

2. The device according to claim 1, characterized in that The pseudorange and pseudorange rate estimator corrects the ionospheric error and tropospheric error of the pseudorange, wherein the corrected ionospheric error is: in, is the measured value after ionospheric delay correction, that is, the corrected ionospheric error, is the frequency of carrier L1, is the frequency of carrier L2, is the pseudorange measurement value of carrier L1, is the pseudorange measurement value of carrier L2.

3. The device according to claim 2, characterized in that The corrected tropospheric error is: in, , , They represent the corrected tropospheric error, dry delay, and wet delay, respectively. and is the mapping function.

4. The device according to claim 3, characterized in that The Kalman filter is used for factor graph fusion processing, and the state equation for fusion is: in, is the predicted state vector at time k, is the forecast error covariance matrix at time k, is the state transition matrix, is the predicted state vector at time k-1, is the forecast error covariance matrix at time k-1, is the control input matrix, is the control input vector, is the process noise covariance matrix, is the transpose.

5. The device according to claim 4, characterized in that The state update formula for fusion is: in, is the Kalman gain, is the observation matrix, is the observation vector, is the observation noise covariance matrix, is the observation matrix, is the updated state variable, is the Kalman gain, is the updated error covariance matrix, and A is the unit matrix.

6. The device according to any one of claims 1 to 5, characterized in that The GNSS dual-frequency antenna is used to acquire GNSS signals from multiple satellites.

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