CN117288186A - GNSS/SINS integrated navigation signal detection and anti-interference method under satellite disturbance environment - Google Patents

GNSS/SINS integrated navigation signal detection and anti-interference method under satellite disturbance environment Download PDF

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CN117288186A
CN117288186A CN202311442073.1A CN202311442073A CN117288186A CN 117288186 A CN117288186 A CN 117288186A CN 202311442073 A CN202311442073 A CN 202311442073A CN 117288186 A CN117288186 A CN 117288186A
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satellite
signal
pseudo
receiver
interference
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陈帅
庄舜
徐川
侯志宽
薛超
丁虎山
宋华
丁鹏飞
沈开淦
陈依玲
赵大想
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Nanjing University of Science and Technology
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Nanjing University of Science and Technology
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    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
    • G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
    • G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
    • G01C21/165—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation combined with non-inertial navigation instruments
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
    • G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
    • G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
    • G01S19/13—Receivers
    • G01S19/21—Interference related issues ; Issues related to cross-correlation, spoofing or other methods of denial of service
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
    • G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
    • G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
    • G01S19/13—Receivers
    • G01S19/35—Constructional details or hardware or software details of the signal processing chain
    • G01S19/37—Hardware or software details of the signal processing chain
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
    • G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
    • G01S19/38—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system
    • G01S19/39—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system the satellite radio beacon positioning system transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
    • G01S19/42—Determining position
    • G01S19/45—Determining position by combining measurements of signals from the satellite radio beacon positioning system with a supplementary measurement
    • G01S19/47—Determining position by combining measurements of signals from the satellite radio beacon positioning system with a supplementary measurement the supplementary measurement being an inertial measurement, e.g. tightly coupled inertial
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02D—CLIMATE CHANGE MITIGATION TECHNOLOGIES IN INFORMATION AND COMMUNICATION TECHNOLOGIES [ICT], I.E. INFORMATION AND COMMUNICATION TECHNOLOGIES AIMING AT THE REDUCTION OF THEIR OWN ENERGY USE
    • Y02D30/00—Reducing energy consumption in communication networks
    • Y02D30/70—Reducing energy consumption in communication networks in wireless communication networks

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  • Engineering & Computer Science (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Computer Networks & Wireless Communication (AREA)
  • Signal Processing (AREA)
  • Automation & Control Theory (AREA)
  • Position Fixing By Use Of Radio Waves (AREA)

Abstract

The invention discloses a GNSS/SINS integrated navigation signal detection and anti-interference method under a satellite disturbance environment, which comprises the following steps: firstly, selecting satellites with elevation angles larger than elevation angle cut-off angles, and distributing weight factors W to each satellite i Sum signal confidence C i The method comprises the steps of carrying out a first treatment on the surface of the Then, according to the satellite weight factors and the satellite signal confidence degrees, carrying out satellite availability calculation and satellite selection work; then adopting a multi-stage combined detection algorithm of the deception jamming signals to detect whether the system is in a deception jamming environment; and finally, adjusting the influence degree of inertial navigation on the power of the captured signal of the receiver and the tracking threshold through an inertial auxiliary loop technology to enable the interference-signal ratio to reach the system requirement. The invention has simple algorithm and knowledgeThe method has the advantages of high speed, high identification accuracy, capability of isolating deception signals and capability of improving the anti-interference capability of the receiver.

Description

GNSS/SINS integrated navigation signal detection and anti-interference method under satellite disturbance environment
Technical Field
The invention relates to the technical field of integrated navigation anti-interference, in particular to a GNSS/SINS integrated navigation signal detection and anti-interference method under a satellite disturbance environment.
Background
Satellite navigation systems are affected by a variety of factors, including urban environments, jamming, and fraud, which pose a significant challenge to the accuracy and reliability of satellite navigation systems, requiring specialized research and countermeasures.
First, urban environments are one of the important factors affecting the performance of satellite navigation systems. High-rise buildings, narrow streets and other obstructions can cause problems with signal attenuation, multipath effects, signal intermittence, etc., which can lead to continual changes in satellite availability, adversely affecting positioning and navigation accuracy.
Secondly, the use of electronic interference devices by adversaries or lawbreakers to interfere with the reception and processing of satellite signals can also seriously affect the performance of satellite navigation systems. In order to cope with such interference, it is necessary to enhance the anti-interference performance of the receiver and to establish a safe and reliable navigation authentication mechanism and anti-spoofing algorithm to ensure that the navigation system uses a true and reliable signal for navigation positioning.
Aiming at the problem that satellite navigation is interfered, researchers are developing novel antenna design and positioning algorithms to improve the signal receiving capability of a navigation system in a complex environment; meanwhile, the multi-sensor fusion technology is explored and used, and navigation systems such as satellites and inertia are combined, so that the accuracy and reliability of the whole navigation are improved.
In the research of satellite signal anti-deception technology, deception detection technology has six major categories, such as navigation data information, space processing, radio frequency front end processing, baseband digital signal processing, positioning navigation operation result and machine learning, but most detection modes have the problem that the theory and engineering implementation have great difference, and in order to realize higher detection probability of deception interference signals, the operation amount must be greatly increased or the complexity of hardware design must be greatly increased.
The deception jamming has two jamming modes of a forwarding type and a generating type, wherein the generating type deception jamming refers to that an attacker generates a signal with misleading property and is used for a jamming or misleading receiver, and the deception jamming is mainly divided into two jamming types of a mutation type and a buffer type. Abrupt spoofing is simple to implement, and is often accompanied by abrupt signal changes, which are intended to confuse the receiver's interpretation and processing of the signal. Slowly varying spoofing interference refers to an attacker achieving an interference effect by gradually changing the properties or parameters of the signal over a longer period of time. Unlike abrupt spoofing, the gradual spoofing changes the signal in a more gradual and concealed manner, which aims to reduce the likelihood of detection. Such spoofing interference usually takes longer to realize slow change of signals, so as to mislead interpretation and judgment of signals by a receiving end, but the implementation difficulty of the interference method is high.
Disclosure of Invention
The invention aims to provide the GNSS/SINS integrated navigation signal detection and anti-interference method which has the advantages of simple algorithm, high identification accuracy and high identification speed, can isolate deception signals and improve the anti-interference capability of a receiver.
The technical solution for realizing the purpose of the invention is as follows: a GNSS/SINS combined navigation signal detection and anti-interference method under a satellite disturbance environment comprises the following steps:
step 1, setting elevation angle cut-off angle E 0 Selecting a cut-off angle E with an elevation angle greater than the elevation angle 0 Weight calculation is carried out on satellites of the model, and the elevation angle is smaller than the elevation angle cut-off angle E 0 The satellite of (2) does not perform weight calculation;
step 2, assigning a weight factor W to each satellite i ；
Step 3, assigning signal confidence degree C to each satellite i ；
Step 4, according to satellite weight factor W i And satellite signal confidence C i Performing availability calculation and satellite selection of satellites;
step 5, detecting whether the system is in a deception jamming environment or not by adopting a deception jamming signal multi-stage joint detection algorithm;
and 6, adjusting the influence degree of inertial navigation on the power of the captured signal of the receiver and the tracking threshold through an inertial auxiliary loop technology to enable the interference-to-signal ratio to reach the system requirement.
Further, the elevation angle cut-off angle E is set in the step 1 0 Selecting a cut-off angle E with an elevation angle greater than the elevation angle 0 Weight calculation is carried out on satellites of the model, and the elevation angle is smaller than the elevation angle cut-off angle E 0 The satellite of (2) does not perform weight calculation, and is specifically as follows:
in an actual scene, if the satellite elevation angle is too low, multipath effect is caused, and the measured pseudo-range is different, so that the positioning accuracy is affected, and therefore, the elevation angle cut-off angle E is set according to the application scene 0 E is used for positioning the ground due to more shielding interference factors such as buildings 0 Set to 10 °, low altitude aircraft position E 0 Set to 5 °. When the elevation angle of the satellite i is lower than the elevation angle cut-off angle E 0 When it is, its weight factor W i Directly set to 0, and do not participate in the weight calculation process.
Further, step 2 is to assign a weight factor W to each satellite i The method is characterized by comprising the following steps:
setting a weight factor W for each acquired satellite i Indicating the influence of the satellite on the positioning accuracy,the greater the weight factor, the greater the effect that the satellite plays in the current positioning; setting the observation matrix as H after n satellites are captured by the current receiver n The observed component of satellite i is h i The observed quantity matrix of the rest n-1 satellites is H n-1 The following formula can be obtained:
according to Sherman-Morrison formula, when A is E R n×n Is a reversible matrix, u, v E R n For column vectors, the inverse of the matrix is:
segment(s)According to formula (2) and formula (3), then:
order theAnd (3) finishing (4) to obtain:
the weights of the satellite i in the current n satellites affecting the satellite distribution are:
in Deltah ij，j Is delta A i Wherein j = 1,2,3,4; w (W) i Weights for satellites i at the current timeFactors.
Further, assigning a signal confidence level C to each satellite as described in step 3 i The method is characterized by comprising the following steps:
setting signal confidence degree C of each satellite i For representing the signal authenticity of the satellite, a lower value indicates that the satellite signal deviates from the estimated real signal and is more difficult to correct, whereas a higher value indicates that the satellite signal is more authentic; the formula is as follows:
wherein C is adaptire_i For the signal self-adaptive confidence of the satellite, according to the signal-to-noise ratio SNR and the pseudo-range error rho err And (3) adaptively adjusting the parameter size and the change condition, and reducing the confidence coefficient of the satellite when the signal-to-noise ratio or the pseudo-range information value has larger jump.
Further, the method for adjusting the signal self-adaptive confidence of the satellite specifically comprises the following steps:
determining signal confidence C of satellite according to signal-to-noise ratio SNR jump i The method is as follows:
in dSNR ave_i Average jump value of signal-to-noise ratio of ith satellite, SNR cur_i For the current signal-to-noise value of the ith satellite,setting the last signal-to-noise value of the ith satellite, n being the count in a period of time and 5 seconds;
threshold dSNR for jump judgment i The method comprises the following steps:
dSNR i ＝dSNR ave_i *1.2 (9)
the next signal-to-noise ratio jump value is compared with a threshold dSNR i Comparing, when the signal to noise ratio jump variable exceeds dSNR i The confidence of the satellite i is reduced; decrease dC of confidence each time i The method comprises the following steps:
the lower the original confidence of the satellite is, the larger the reduction of the satellite confidence is; and when the signal-to-noise ratio jump variable does not exceed the signal-to-noise ratio average jump value, the confidence of the satellite is fixedly improved by 0.05.
Further, the step 4 is based on the satellite weight factor W i And satellite signal confidence C i The availability calculation and satellite selection work of satellites are carried out, and the method specifically comprises the following steps:
step 4.1, according to the satellite weight factor W i And satellite signal confidence C i The priority of the satellite is calculated, and the formula is as follows:
in the formula of Priority i Indicating the priority of the satellite i for positioning, and n indicates the number of currently available satellites;
step 4.2, when the satellite is in an interference-free environment, the influence degree of the satellite weight factor and the confidence level on the satellite selection priority is equal; when the environment is interfered, the influence degree of the confidence coefficient on the satellite selection priority is unchanged, the influence proportion of the weight factor on the satellite selection priority is adjusted along with the number of currently available satellites, the number of satellites is increased based on the minimum requirement of 4 satellites for positioning, and the influence degree of the weight factor on the satellite selection priority is reduced.
Further, in step 5, a multi-level joint detection algorithm for the spoofing interference signal is adopted to detect whether the system is in a spoofing interference environment, which is specifically as follows:
step 5.1, detecting the signal-to-noise ratio, which is as follows:
under the condition of no interference, the signal-to-noise ratio of satellite signals is high; under the condition that satellite signals are interfered by deception signals, the change of signal to noise ratio is as follows:
where SNR' represents the signal-to-noise value, P, of the receiver after being subjected to a spoofing signal s For the power of the real satellite signal, P i To deceive the power of the signal N 0 Is the system noise power spectral density.
Converting the real satellite signal power change in the formula (15) into the distance information of the receiver and the satellite which is easier to acquire, wherein the change of the signal-to-noise ratio of the receiver after being subjected to deception signals is as follows:
where α is a parameter related to the distance between the receiver and the satellite;
under the interference of the deception signal, the signal-to-noise ratio of the satellite signal is reduced due to the influence of the deception signal power, and the higher the deception signal power is, the lower the signal-to-noise ratio is, namely alpha P i Is the added value of the satellite signal noise base received by the receiver, and the deception jamming signal can increase the noise base of the received signal;
setting a corresponding threshold value according to the actual condition of the receiver, and judging whether the receiver is attacked by deception signals or not; the signal-to-noise ratio threshold is set as follows:
In the SNR 0 For the signal-to-noise threshold, SNR i_visible The signal to noise ratio of the ith visible satellite is given, and N is the number of visible satellites;
when the threshold value is exceeded, giving an alarm that the deception jamming is possible, and reducing the confidence of the satellite i according to a formula (10);
step 5.2, detecting the clock difference of the receiver, which is specifically as follows:
the pseudo-range equation of a single satellite after satellite clock correction is as follows:
wherein ρ is i Representing the pseudorange of the ith satellite to the target antenna, (x) s ，y s ，z s ) Is the three-dimensional position of the satellite, (x) u ，y u ，z u ) For three-dimensional position of target object, Δt u For the local clock-difference to be the local clock-difference,is the sum of ionosphere, troposphere and other error terms;
after N pseudo-range equations of the formula are obtained, simultaneous equation sets and linearization processing are carried out to obtain:
G·ΔX u ＝b (19)
wherein DeltaX is u ＝[Δx u Δy u Δz u Δt u ] T Is a state variable;
the least squares solution of equation (19) is:
Δx 0 ＝(G T G) -1 G T b (22)
due to Deltax 0 ＝[Δx u Δy u Δz u Δt u ] T The clock difference delta t can be obtained through the method (22) u ；
The common error of the pseudo-range observed quantity can greatly influence the clock difference, the submersion of the deception signal can cause the fluctuation of the clock difference, and the deception signal can be detected according to the change; because the variable is directly used as the judgment basis of the deception signal, a plurality of judgment criteria are needed to prevent the outlier in the data from influencing the detection process:
Δt u_pre0 -Δt u_pre1 ≤2×(Δt u_pre1 -Δt u ) (23)
Wherein Δt is u_pre0 、Δt u_pre1 Respectively representing the clock difference values of the previous time and the previous time, and if the clock difference jump amplitude of the current time is larger, early warning and saving delta t u_pre1 Is a new Δt u _ pre0 For continuing to compare the jump amplitude at the next moment, and alarming that the jump amplitude is possibly attacked by deceptive jamming signals when the jump amplitude exceeds twice; under normal conditions, the clock error jump has upper and lower limits, a wild value exists at part of the time, when an interference signal appears, the clock error jump amplitude exceeds the range of the upper and lower limits, at the moment, the occurrence of the interference signal can be judged, and pseudo-range rate detection are carried out to confirm the existence of a deception signal;
and 5.3, according to the pseudo-range and pseudo-range rate information, realizing inertial auxiliary pseudo-range and pseudo-range rate consistency detection, wherein the method comprises the following steps of:
by analyzing the change of the pseudo range and the pseudo range rate, judging whether the satellite signal is interfered and deceptively, wherein the formula is as follows:
wherein ρ is u 、Pseudo-range and pseudo-range rate, ρ, of antenna relative to satellite output by GNSS receiver m 、/>Pseudo range, pseudo range rate, deltaρ and/or Deltaρ of carrier relative to satellite calculated by combining inertial navigation information with satellite ephemeris>Respectively isError between the two calculation modes;
for the ith satellite, its corresponding ρ m 、The calculation method is as follows:
wherein ρ is m_i 、Pseudo-range and pseudo-range rate of SINS relative to satellite i, (x) m ，y m ，z m )、/>Position, speed, < >/of inertial system in ECEF coordinate system respectively>The position and the speed of the satellite i under the ECEF coordinate system are respectively;
normally, the inertial system has very small error divergence in a short time, Δρ,The values are very small; however, when the abrupt deception signal exists, the position, speed or time information of the receiver is biased, and the pseudo range rate calculated by the receiver are biased; according to Δρ, ++>The setting method refers to the formulas (8) and (9) and judges that the spoofing is received when the values of the two exceeds the set thresholdInterference of the spoofing signal;
when the slow-change type deception signal exists, the periodic enhancement detection is designed, the characteristic of high precision in a short time of inertial navigation is utilized, the calculated pseudo range rate of the inertial navigation speed is periodically constructed, and the statistical test quantity is constructed with the calculated pseudo range rate of the satellite navigation result, so that the slow-change type deception detection is realized; since inertial navigation position information is obtained by integrating velocity information, errors in position information accumulate faster, and pseudo-range information ρ calculated from position in equation (24) m The error accumulation of (a) is faster than the pseudo-range rate information Therefore, adopting the consistency detection of the pseudo range rate to carry out the statistics of the detection quantity on the pseudo range rate sequence;
deriving a satellite pseudo-range equation of formula (18), and obtaining a pseudo-range rate equation, wherein the formula is as follows:
in the middle ofPseudo-range rate for the ith satellite, +.>Variable value representing the distance between satellite and receiver, < >>Representing the receiver clock drift +.>Representing the ionosphere, troposphere delay rate of change, which is less sensitive to time-change and is therefore omitted;
will beExpressed as the satellite to receiver motion velocity vector difference, the formula (27) is:
in the formula, v i The motion velocity vector of the ith satellite is v is the motion velocity vector of the receiver; when the original satellite signal is processed by the spoofing interference source to generate a spoofing signal or directly forward, the speed vector inconsistency can cause jump of the pseudo range rate of the actually acquired satellite signal due to change of the signal direction, so that the statistical test quantity of the residual error of the pseudo range rate of the receiver is constructed as follows:
in the method, in the process of the invention,theoretical residual error representing pseudo-range rate between receiver and satellite i, and using inertial navigation speed information short-time high-precision characteristic to calculate pseudo-range rate +.>Substitution of theoretical pseudo-range rate->
In the inspection period, when the residual exceeds the residual threshold caused by the set inertial navigation drift, determining the signal of the satellite as a deceptive signal, calculating the mean value and standard deviation of the residual, and identifying abnormal values by using the statistical data, wherein the formula is as follows:
In the method, in the process of the invention,representing the mean of the residual, s ∈ Represents the standard deviation of the residual error, n represents the number of samples, ε i A residual representing the ith observation; based on the nature of the normal distribution, about 99.7% of observations will be at 3s ∈ Thus when the observed value exceeds 3s ∈ Is determined to be an outlier.
Further, in the step 6, the influence degree of inertial navigation on the power of the captured signal and the tracking threshold of the receiver is adjusted by using the inertial auxiliary loop technology, so that the interference-to-signal ratio reaches the system requirement, which is as follows:
in the inertial auxiliary capturing process, the Doppler frequency shift, the frequency error and the code shift caused by the Doppler frequency shift and the pseudo-range information are calculated by utilizing the position, the speed and the acceleration information output by the inertial navigation system and combining the position and the speed information of the satellite, and the calculated Doppler frequency shift is calculatedPN code frequency offset τ and code shift Δτ dyn And respectively transmitting the signals to a numerical control oscillator NCO, a voltage-controlled oscillator VCC and a code generator PN, and setting initial values of code phase and Doppler frequency shift by using time positions provided by GNSS/SINS to finish the capturing process.
Further, the structure of the inertial auxiliary loop is specifically as follows:
the code loop adopts a carrier-assisted second-order delay phase-locked loop DLL; the carrier wave loop adopts an inertial auxiliary third-order phase-locked loop (PLL); a first order low pass filter is built in the inertial auxiliary part The loop filter of the third-order phase-locked loop PLL is selected as follows:
obtaining external frequency auxiliary deviation through speed deviation and clock frequency deviation estimated by a navigation filter, adding a reference input signal, and obtaining external frequency auxiliary information through an inertial filter after differentiation;
the phase error output by the phase discriminator is the difference value between the reference input signal and the phase of the output signal of the external phase noise and voltage-controlled oscillator;
the phase error is filtered by a loop filter of a third-order phase-locked loop (PLL), added with external frequency auxiliary information and integrated to obtain the output signal phase of the voltage-controlled oscillator.
Compared with the prior art, the invention has the remarkable advantages that: (1) Meanwhile, two main factors of satellite space geometric distribution and observation errors affecting positioning accuracy are considered, and the availability judgment of the current satellite is completed in real time, so that the satellite positioning method is used for completing positioning satellite selection work, and is simple in algorithm and high in recognition speed; (2) The advantage of GNSS/SINS integrated navigation is fully utilized, an algorithm is designed to enable the system to have good detection effect on mutation type and buffer type deception signals, the system also has certain suppression capability on interference signals, the recognition accuracy is high, deception signals can be isolated, and the anti-interference capability of a receiver is improved.
Drawings
FIG. 1 is a flow chart of a method for detecting GNSS/SINS integrated navigation signals and resisting interference in a satellite disturbance environment according to the present invention.
Fig. 2 is a graph of a satellite selection effect for a complex road segment in an embodiment of the present invention.
FIG. 3 is a graph showing the effect of satellite selection on horizontal positioning accuracy in a complex environment in accordance with an embodiment of the present invention.
FIG. 4 is a graph of the clock-difference transition result in an embodiment of the invention.
Detailed Description
The invention will now be described in further detail with reference to the drawings and to specific examples.
Referring to fig. 1, the method for detecting and resisting interference of a GNSS/SINS integrated navigation signal in a satellite disturbance environment according to the present invention includes the following steps:
a GNSS/SINS combined navigation signal detection and anti-interference method under a satellite disturbance environment comprises the following steps:
step 1, setting elevation angle cut-off angle E 0 Selecting a cut-off angle E with an elevation angle greater than the elevation angle 0 Weight calculation is carried out on satellites of the model, and the elevation angle is smaller than the elevation angle cut-off angle E 0 The satellite of (2) does not perform weight calculation;
step 2, assigning a weight factor W to each satellite i ；
Step 3, assigning signal confidence degree C to each satellite i ；
Step 4, according to satellite weight factor W i And satellite signal confidence C i Performing availability calculation and satellite selection of satellites;
Step 5, detecting whether the system is in a deception jamming environment or not by adopting a deception jamming signal multi-stage joint detection algorithm;
and 6, adjusting the influence degree of inertial navigation on the power of the captured signal of the receiver and the tracking threshold through an inertial auxiliary loop technology to enable the interference-to-signal ratio to reach the system requirement.
As a specific example, in step 1, an elevation angle cut-off angle E is set 0 Selecting a cut-off angle E with an elevation angle greater than the elevation angle 0 Weight calculation is carried out on satellites of the model, and the elevation angle is smaller than the elevation angle cut-off angle E 0 The satellite of (2) does not perform weight calculation, and is specifically as follows:
in an actual scene, if the satellite elevation angle is too low, multipath effect is caused, and the measured pseudo-range is different, so that the positioning accuracy is affected, and therefore, the elevation angle cut-off angle E is determined according to the application scene 0 Is set to 5-10 DEG, when the elevation angle of the satellite i is lower than the elevation angle cut-off angle E 0 When it is, its weight factor W i Directly set to 0, and do not participate in the weight calculation process.
As a specific example, in step 2, a weight factor W is assigned to each satellite i The method is used for rapidly giving the degree of distribution of each satellite in the whole satellite in the real-time captured satellites, and specifically comprises the following steps:
in the early stages of receiver design and performance analysis, the concept of a precision factor was defined, where the geometric precision factor is expressed as follows:
H in i，i (i=1, 2,3, 4) is (H T H) -1 Is H is the observation matrix.
The precision factor can be regarded as a linear mapping from measurement errors in observables to state estimation errors, since each row vector of the observation matrix H consists of the three directional cosine vectors of a satellite and a user and 1, the specific dimension and form is determined by the X-array setup, so (H) T H) -1 It is necessary to relate to the geometrical distribution of the satellites.
The satellites captured by the GNSS receiver at different times are constantly changing due to the relative movement of the satellites and the receiver, as are the relative positions of each captured satellite in the sky. Setting a weight factor W for each acquired satellite i The magnitude of the influence of the satellite on the positioning accuracy is represented, and the larger the weight factor is, the larger the satellite plays a role in the current positioning. Setting the observation matrix as H after n satellites are captured by the current receiver n The observation component of a satellite i is h i The observed quantity matrix of the rest n-1 satellites is H n-1 The following formula can be obtained:
according to Sherman-Morrison formula, when A is E R n×n Is a reversible matrix, u, v E R n For column vectors, the inverse of the matrix is:
is provided with According to formula (2) and formula (3), then:
order theAnd (3) finishing (4) to obtain:
the weights of the satellite i in the current n satellites affecting the satellite distribution are:
in Deltah ij，j (is DeltaA i Wherein j = 1,2,3,4; w (W) i The weight factor of the satellite i at the current moment is obtained.
As a specific example, in step 3, a signal confidence level C is assigned to each satellite i The satellite signal quality feedback device is used for quickly feeding back satellite signal quality change in a complex environment, and auxiliary weight factors are used for completing satellite selection work in the complex environment, particularly in a deception interference environment, and specifically comprises the following steps:
signal confidence C for each satellite i The lower the value is, the more the satellite signal deviates from the estimated real signal and is more difficult to correct, otherwise, the higher the satellite signal is; the formula is as follows:
wherein C is adaptire_i According to SNR, pseudo-range error ρ err And (3) adaptively adjusting the size and the change condition of the equal parameters, and reducing the confidence coefficient of the satellite when the signal-to-noise ratio or the pseudo-range information value has larger jump.
Determining signal confidence C of satellite according to signal-to-noise ratio SNR jump i The method is as follows:
in dSNR ave_i Average jump value of signal-to-noise ratio of ith satellite, SNR cur_i For the current signal-to-noise value of the ith satellite,for the last signal-to-noise value of the ith satellite, n is the count in a period of time, and can be set to be 5 seconds;
threshold dSNR for jump judgment i The method comprises the following steps:
dSNR i ＝dSNR ave_i *1.2 (9)
the next signal-to-noise ratio jump value is compared with a threshold dSNR i Comparing, the next signal-to-noise ratio jump variable exceeds dSNR i The confidence of the satellite i is reduced; decrease dC of confidence each time i The method comprises the following steps:
the lower the original confidence of the satellite, the greater the amount of decrease in the satellite confidence. And when the signal-to-noise ratio jump variable does not exceed the signal-to-noise ratio average jump value, the confidence of the satellite is fixedly improved by 0.05.
As a specific example, in step 4, the satellite weight factor W is used i And satellite signal confidence C i The availability calculation and satellite selection work of satellites are carried out, and the method specifically comprises the following steps:
step 4.1, according to the satellite weight factor W i And satellite signal confidence C i Calculating the priority of the satellite, and the formulaThe method comprises the following steps:
in the formula of Priority i Indicating the priority of satellite i for positioning and n indicating the number of currently available satellites.
And 4.2, when the environment is in a non-interference environment, the influence degree of the satellite weight factors and the confidence degrees on the satellite selection priority is equal, when the environment is interfered, the influence degree of the confidence degrees on the satellite selection priority is unchanged, the influence proportion of the weight factors on the satellite selection priority is adjusted along with the number of currently available satellites, the number of satellites is increased based on the minimum requirement of positioning on the number of satellites, and the influence degree of the weight factors on the satellite selection priority is smaller.
The satellite selecting effect diagram of the complex road section is shown in fig. 2. When satellite signals are interfered, positioning results without satellite selection generate errors with different sizes, after satellite selection is judged through the weight factors and the confidence coefficient, the satellite positioning results are basically stable before and after the satellite selection is interfered, the influence of the satellite selection on horizontal positioning accuracy in a complex environment is shown in a figure 3, and the positioning accuracy is reduced by 0-2 m due to the removal of part of satellites, but the positioning accuracy can still be maintained within 3 m under the conditions of large interference and serious reduction of satellite signal quality. The receiver used in the test is a Beidou single-frequency point receiver, the number of channels is 12, and theoretically, the multimode multi-frequency point receiver can obtain a better positioning result under the processing of the method.
As a specific example, in step 5, a multi-level joint detection algorithm for the spoofing interference signal is used to detect whether the system is in a spoofing interference environment, which is specifically as follows:
the manner in which the jammer signal is spoofed is varied, one common way being to fool the receiver by transmitting a dummy signal similar to the real signal, which can be achieved by transmitting the receiver with a dummy signal of a different modulation scheme, such as a pseudo random noise code, a pseudo frequency hopping code, etc. In addition, the deception jamming signal can influence the positioning result by changing the arrival time, phase, amplitude and other parameters of the signal.
In the specific implementation of the rogue interfering signal, the characteristics of the satellite signal such as modulation, coding, transmission, etc. need to be studied and understood in depth. For example, the pseudorandom noise code in a GPS signal is a long sequence code with a period of 1 millisecond and multiple repetitions occur during signal transmission. Thus, if a rogue can acquire the pseudo-random noise code of the GPS signal, a pseudo-signal similar to the real signal can be generated, and the effect of rogue interference signals can be achieved by changing the arrival time, the phase, the amplitude and other parameters.
The basic characteristics of the GPS signal can be expressed using the following formula:
s t (t)＝A(t)·cos[2πf c t+φ(t)] (12)
wherein s is t (t) represents the carrier wave of the GPS signal, A (t) represents the amplitude of the signal, f c Representing the carrier frequency of the signal and phi (t) represents the phase of the signal.
For the generated deception jamming, after the satellite signal parameters are obtained, the satellite receiver of the jamming source uses various technical means, such as frequency offset, code phase change, signal intensity change and the like to disguise the false navigation signal, so that the false navigation signal cannot be distinguished from the true navigation signal. The generated spoofing disturbance may be expressed by the following formula:
s f (t)＝A(t)·cos[2πf c t+φ(t)]+n(t) (13)
wherein s is f (t) represents a spoofing jamming signal, A (t) represents the amplitude of the signal, f c Represents the carrier frequency of the signal, phi (t) represents the phase of the signal, and n (t) represents the additive white gaussian noise.
For forward spoofing interference, the source receiver forwards the received satellite signal to the target receiver using a high power transmitting device. The spoofing source changes parameters such as time of arrival, phase, amplitude, etc. of the signal by intercepting and decoding the real signal, and then transmits the spurious signal to the receiver by various means. The forwarded fraud can be expressed by the following formula:
s r (t)＝s t (t-τ)+n(t) (14)
wherein s is r (t) represents the signal observed by the receiver, s t (t) represents a real signal, τ represents a propagation delay of the signal, and n (t) represents additive white gaussian noise. The above formula can be used to understand what effect the spoofing signal has on the satellite signal. According to the method shown in fig. 1, a spoofing signal multistage joint detection algorithm is adopted to detect whether a spoofing interference signal exists, a calculation mode of selecting priorities of satellites in a formula (11) is affected, and the confidence of the corresponding satellite is adaptively adjusted to a certain degree according to a detection result of a detection method described below. The deception jamming detection steps are as follows:
step 5.1, detecting the signal-to-noise ratio, which is as follows:
the signal-to-noise ratio of satellite signals is generally high without interference; under the condition that satellite signals are interfered by deception signals, the change of signal to noise ratio is as follows:
Where SNR' represents the signal-to-noise value, P, of the receiver after being subjected to a spoofing signal s For the power of the real satellite signal, P i To deceive the power of the signal N 0 Is the system noise power spectral density.
Converting the real satellite signal power change in the formula (15) into the distance information of the receiver and the satellite which is easier to acquire, wherein the change of the signal-to-noise ratio of the receiver after being subjected to deception signals is as follows:
where α is a parameter related to the distance between the receiver and the satellite.
It can be seen that under the influence of the spoofing signal, the signal-to-noise ratio of the satellite signal is reduced by the influence of the spoofing signal power, and the higher the spoofing signal power is, the lower the signal-to-noise ratio is, and the alpha P is i Is the added value of the satellite signal noise base received by the receiverSpoofing the interfering signal may increase the noise floor of the received signal;
the receiver can be preliminarily judged whether to suffer from the attack of the deception signal or not by setting the corresponding threshold value according to the actual condition of the receiver, and the signal-to-noise ratio threshold value is set as follows:
in the SNR 0 For the signal-to-noise threshold, SNR i_visible The signal to noise ratio of the ith visible satellite is given, and N is the number of visible satellites;
when the threshold is exceeded, an alarm is given that there is a potential for fraud, and the confidence of satellite i is reduced as per equation (10).
Step 5.2, detecting the clock difference of the receiver, which is specifically as follows:
the pseudo-range equation of a single satellite after satellite clock correction is as follows:
wherein ρ is i Representing the pseudorange of the ith satellite to the target antenna, (x) s ，y s ，z s ) Is the three-dimensional position of the satellite, (x) u ，y u ，z u ) For three-dimensional position of target object, Δt u For local clock difference, n ρi Is the sum of ionosphere, troposphere and other error terms;
after N pseudo-range equations of the formula are obtained, simultaneous equation sets and linearization processing are carried out to obtain:
G·ΔX u ＝b (19)
wherein DeltaX is u ＝[Δx u Δy u Δz u Δt u ] T Is a state variable.
The least squares solution of equation (19) is:
Δx 0 ＝(G T G) -1 G T b (22)
due to Deltax 0 ＝[Δx u Δy u Δz u Δt u ] T The clock difference delta t can be obtained through the method (22) u 。
The common error of the pseudo-range observed quantity does not affect the position calculation, but can greatly affect the clock difference, the submersion of the deception signal can cause the fluctuation of the clock difference, and the deception signal can be detected according to the change; because the variable is directly used as the judgment basis of the deception signal, a plurality of judgment criteria are needed to prevent the outlier in the data from influencing the detection process:
Δt u_pre0 -Δt u_pre1 ≤2×(Δt u_pre1 -Δt u ) (23)
wherein Δt is u_pre0 、Δt u _ pre1 Respectively representing the clock difference values of the previous time and the previous time, and if the clock difference jump amplitude of the current time is larger, early warning and saving delta t u_pre1 Is a new Δt u_pre0 For the next time the jump amplitude continues to be compared, and when more than two times it is alerted that it is likely to suffer from a rogue jamming signal attack. As shown in the clock-difference jump result diagram shown in fig. 4, under normal conditions, the clock-difference jump has upper and lower limits, and a wild value exists at some moments, when an interference signal occurs, the clock-difference jump amplitude exceeds the upper and lower limits, at this time, the occurrence of the interference signal can be judged, and pseudo-range rate detection are performed to confirm the existence of a spoofed signal.
And 5.3, realizing inertial auxiliary pseudo-range/pseudo-range rate consistency detection according to the pseudo-range and pseudo-range rate information, wherein the method comprises the following steps of:
spoofing typically involves modification or falsification of the satellite signal so that the receiver misbelieves that it is farther or closer from the satellite, which may result in the receiver calculating erroneous position information. In the GNSS/SINS integrated navigation, the inertial system may measure the motion state of the navigation system using the accelerometer and gyroscope in the inertial measurement unit IMU. Through the integration process, position, speed and attitude information of the navigation system can be obtained. The inertial system can estimate the approximate position and velocity information of each satellite in combination with the current motion state of the receiver and the satellite ephemeris, and then achieve receiver positioning by calculating the pseudo-range and the pseudo-range rate between the receiver and the satellites. Such information may assist in selecting the appropriate satellite. When satellite signals are deceptively interfered, the pseudo range and the pseudo range rate of the signals can be changed abnormally, so that whether the satellite signals are interfered and deceptively interfered can be judged by analyzing the change of the pseudo range and the pseudo range rate, and the calculation process is as follows:
Wherein ρ is u 、Pseudo-range and pseudo-range rate, ρ, of antenna relative to satellite output by GNSS receiver m 、/>Pseudo range, pseudo range rate, deltaρ and/or Deltaρ of carrier relative to satellite calculated by combining inertial navigation information with satellite ephemeris>Errors between the two calculation modes are respectively;
for the ith satellite, its corresponding ρ m 、The calculation method is as follows: />
Wherein ρ is m_i 、Pseudo-range and pseudo-range rate of SINS relative to satellite i, (x) m ，y m ，z m )、/>Position, speed, < >/of inertial system in ECEF coordinate system respectively>The position and the speed of the satellite i under the ECEF coordinate system are respectively;
normally, the inertial system has very small error divergence in a short time, Δρ,The values are very small; however, when the abrupt deception signal exists, the position, speed or time information of the receiver is biased, and the pseudo range rate calculated by the receiver are biased. According to Δρ, ++>The setting method refers to the formulas (8) and (9), and when the two are detected to exceed the set threshold, the probability of being interfered by the deception signal can be judged.
The implementation difficulty of the delayed type deception signal is high, but the concealment and camouflage are high, and the damage is high, so that the delayed type deception signal needs to be detected. In the actual situation, the deception signal cannot be seamlessly cut into a receiver loop, and the deception signal can cause the change of observables, so that the periodic enhancement detection is designed, the characteristic of high accuracy in inertial navigation in a short time is utilized, the calculated pseudo range rate of the inertial navigation speed is periodically constructed, the statistical test quantity is constructed with the calculated pseudo range rate of the satellite navigation result, and the slow-change deception detection is realized.
Since inertial navigation position information is obtained by integrating velocity information, errors in position information accumulate faster, and pseudo-range information ρ calculated from position in equation (24) m The error accumulation of (a) is faster than the pseudo-range rate informationTherefore, only the pseudo range rate consistency detection is used for carrying out the verification quantity statistics on the pseudo range rate sequence.
Deriving a satellite pseudo-range equation of formula (18), obtaining a pseudo-range rate equation, and simplifying the expression as follows:
in the middle ofPseudo-range rate for the ith satellite, +.>Variable value representing the distance between satellite and receiver, < >>Representing the receiver clock drift +.>Indicating ionosphere, troposphere delay rate of change,/->Items are less sensitive to time variations and are therefore discarded;
will beRepresented as satellite and receiverMotion velocity vector difference, and the formula (27) is:
in the formula, v i V is the motion velocity vector of the receiver, which is the motion velocity vector of the ith satellite. When the original satellite signal is processed by the spoofing interference source to generate a spoofing signal or directly forward, the speed vector inconsistency can cause jump of the pseudo range rate of the actually acquired satellite signal due to change of the signal direction, so that the statistical test quantity of the residual error of the pseudo range rate of the receiver is constructed as follows:
in the method, in the process of the invention,theoretical residual error representing pseudo-range rate between receiver and satellite i, and using inertial navigation speed information short-time high-precision characteristic to calculate pseudo-range rate +. >Substitution of theoretical pseudo-range rate->
In the inspection period, when the residual exceeds the residual threshold caused by the set inertial navigation drift, determining the signal of the satellite as a deceptive signal, calculating the mean value and standard deviation of the residual, and identifying abnormal values by using the statistical data, wherein the formula is as follows:
/>
in the method, in the process of the invention,representing the mean of the residual, s ∈ Represents the standard deviation of the residual error, n represents the number of samples, ε i A residual representing the ith observation; based on the nature of the normal distribution, about 99.7% of observations will be at 3s ∈ Thus when a certain observed value exceeds 3s ∈ Is determined to be an outlier.
As a specific example, in step 6, the influence degree of inertial navigation on the captured signal power and the tracking threshold of the receiver is adjusted by using the inertial auxiliary loop technology, so that the interference-to-signal ratio reaches the system requirement, and the anti-interference capability of the receiver is improved, which is specifically as follows:
in the inertial auxiliary capturing process, the position, speed and acceleration information output by an inertial navigation system are utilized, the Doppler frequency shift, the frequency error and the code shift caused by the Doppler frequency shift and the pseudo-range information are calculated by combining the position and speed information of a satellite, the receiver loop is assisted by the derived information, and the calculated Doppler frequency shift is calculated PN code frequency offset τ and code shift Δτ dyn The initial values of code phase and Doppler frequency shift are set by utilizing the time positions provided by GNSS/SINS to finish the capturing process, and the bottom noise of the channel is reduced under the condition of inertial assistance.
The inertial auxiliary tracking loop structure is as follows: the code loop adopts a carrier-assisted second-order delay phase-locked loop DLL; the carrier wave loop adopts an inertial auxiliary third-order phase-locked loop (PLL); a first order low pass filter is built in the inertial auxiliary partThe loop filter of the third-order phase-locked loop PLL is selected as follows:
obtaining external frequency auxiliary deviation through speed deviation and clock frequency deviation estimated by a navigation filter, adding a reference input signal, and obtaining external frequency auxiliary information through an inertial filter after differentiation; the phase error output by the phase discriminator is the difference value between the reference input signal and the phase of the output signal of the external phase noise and voltage-controlled oscillator; the phase error is filtered by a loop filter of a third-order phase-locked loop (PLL), added with external frequency auxiliary information and integrated to obtain the output signal phase of the voltage-controlled oscillator.
In a GNSS receiver, a tracking threshold T thr For determining whether the satellite signal received by the receiver can be tracked, typically a constant value; when inertial assistance is performed, the tracking capability of the receiver to the satellite is enhanced, and the tracking threshold limit is reduced, and the formula is:
C/N in 0 S is the received signal power, N is the noise power, B is the tracking loop signal bandwidth, and T is the integration time.
From this, it can be seen that the inertial assist can significantly reduce the tracking loop signal bandwidth B and greatly reduce the channel bottom noise N, and the effect of the inertial assist is as follows:
S＝S′-ΔS (34)
T＝T thr -ΔT (35)
wherein S' is the signal power captured by the satellite receiver, and delta S and delta T are the signal power after inertial assistance and the reducible value of the tracking threshold respectively, and reflect the reduction degree of the navigation system on the dependence of satellite information. When the magnitude of Δs and Δt is reduced, the calculation amount and the power consumption increase, but according to equations (33) to (35), it can be obtained that increasing the magnitude of Δs and Δt can improve the interference-to-signal ratio, the tracking performance can be enhanced, and thus the anti-spoofing capability of the receiver can be improved.
In summary, the GNSS/SINS integrated navigation signal detection and anti-interference method under the satellite disturbance environment can realize more excellent satellite availability rapid judgment capability under the complex environment and ensure positioning accuracy; the detection of the deception signal is rapidly and accurately realized by adopting a deception signal multistage joint detection algorithm, and satellite star signals which receive the deception signal are marked and isolated; the detection accuracy and the comprehensiveness of the deception signals are further ensured by adopting the inertial auxiliary deception signal detection and the combined navigation technology.

Claims (9)

1. The GNSS/SINS integrated navigation signal detection and anti-interference method under the satellite disturbance environment is characterized by comprising the following steps:
step 1, setting elevation angle cut-off angle E 0 Selecting a cut-off angle E with an elevation angle greater than the elevation angle 0 Weight calculation is carried out on satellites of the model, and the elevation angle is smaller than the elevation angle cut-off angle E 0 The satellite of (2) does not perform weight calculation;
step 2, assigning a weight factor W to each satellite i ；
Step 3, assigning signal confidence degree C to each satellite i ；
Step 4, according to satellite weight factor W i And satellite signal confidence C i Performing availability calculation and satellite selection of satellites;
step 5, detecting whether the system is in a deception jamming environment or not by adopting a deception jamming signal multi-stage joint detection algorithm;
and 6, adjusting the influence degree of inertial navigation on the power of the captured signal of the receiver and the tracking threshold through an inertial auxiliary loop technology to enable the interference-to-signal ratio to reach the system requirement.
2. The method for detecting and resisting interference of integrated GNSS/SINS signals in a satellite disturbance environment according to claim 1, wherein the setting of the elevation cut-off angle E in step 1 0 Selecting an elevation angle greater than the elevation angleCut-off angle E 0 Weight calculation is carried out on satellites of the model, and the elevation angle is smaller than the elevation angle cut-off angle E 0 The satellite of (2) does not perform weight calculation, and is specifically as follows:
Elevation angle cut-off angle E according to application scene 0 Is set to 5-10 DEG, when the elevation angle of the satellite i is lower than the elevation angle cut-off angle E 0 When the weight factor W is used i Directly set to 0, and do not participate in the weight calculation process.
3. The method for detecting and resisting interference of integrated GNSS/SINS signals in a satellite disturbance environment according to claim 2, wherein step 2 is characterized by assigning a weight factor W to each satellite i The method is characterized by comprising the following steps:
setting a weight factor W for each acquired satellite i The influence of the satellite on the positioning precision is represented, and the larger the weight factor is, the larger the satellite plays a role in the current positioning; setting the observation matrix as H after n satellites are captured by the current receiver n The observed component of satellite i is h i The observed quantity matrix of the rest n-1 satellites is H n-1 The following formula is obtained:
according to Sherman-Morrison formula, when A is E R n×n Is a reversible matrix, u, v E R n For column vectors, the inverse of the matrix is:
is provided withAccording to formula (2) and formula (3), then:
order theAnd (3) finishing (4) to obtain:
the weights of the satellite i in the current n satellites affecting the satellite distribution are:
in Deltah ij，j Is delta A i Wherein j = 1,2,3,4; w (W) i The weight factor of the satellite i at the current moment is obtained.
4. The method for detecting and resisting interference of integrated GNSS/SINS signals in a satellite disturbance environment according to claim 3, wherein in step 3, a signal confidence level C is assigned to each satellite i The method is characterized by comprising the following steps:
setting signal confidence degree C of each satellite i The lower the value used to represent the signal authenticity of the satellite, the more difficult it is to correct the satellite signal from the estimated true signal, and conversely, the higher the satellite signal authenticity, the following formula is:
wherein C is adaptive_i For the signal self-adaptive confidence of the satellite, according to the signal-to-noise ratio SNR and the pseudo-range error rho err And (3) adaptively adjusting the parameter size and the change condition, and reducing the confidence coefficient of the satellite when the signal-to-noise ratio or the pseudo-range information value has larger jump.
5. The method for detecting and resisting interference of combined GNSS/SINS navigation signals in a satellite disturbance environment according to claim 4, wherein the method for adjusting the signal adaptive confidence of the satellite is specifically as follows:
determining signal confidence C of satellite according to signal-to-noise ratio SNR jump i The method is as follows:
in dSNR ave_i Average jump value of signal-to-noise ratio of ith satellite, SNR cur_i For the current signal-to-noise value of the ith satellite,for the last signal-to-noise value of the ith satellite, n is the count in 5 seconds;
Threshold dSNR for jump judgment i The method comprises the following steps:
dSNR i ＝dSNR ave_i *1.2 (9)
the next signal-to-noise ratio jump value is compared with a threshold dSNR i Comparing, the next signal-to-noise ratio jump variable exceeds dSNR i The confidence of the satellite i is reduced; decrease dC of confidence each time i The method comprises the following steps:
the lower the original confidence of the satellite is, the larger the reduction of the satellite confidence is; and when the signal-to-noise ratio jump variable does not exceed the signal-to-noise ratio average jump value, the confidence of the satellite is fixedly improved by 0.05.
6. The method for GNSS/SINS integrated navigation signal detection and interference suppression in a satellite disturbance environment according to claim 5, wherein said step 4 is based on a satellite weight factor W i And satellite signal confidence C i Enter intoThe satellite availability calculation and satellite selection work is as follows:
step 4.1, according to the satellite weight factor W i And satellite signal confidence C i The priority of the satellite is calculated, and the formula is as follows:
in the formula of Priority i Indicating the priority of the satellite i for positioning, and n indicates the number of currently available satellites;
step 4.2, when the satellite is in an interference-free environment, the influence degree of the satellite weight factor and the confidence level on the satellite selection priority is equal; when the environment is interfered, the influence degree of the confidence coefficient on the satellite selection priority is unchanged, the influence proportion of the weight factor on the satellite selection priority is adjusted along with the number of currently available satellites, the number of satellites is increased based on the minimum requirement of 4 satellites for positioning, and the influence degree of the weight factor on the satellite selection priority is reduced.
7. The method for detecting and resisting interference of GNSS/SINS integrated navigation signals in a satellite disturbance environment according to claim 6, wherein in step 5, a multi-level joint detection algorithm for rogue disturbance signals is adopted to detect whether the system is in a rogue disturbance environment, specifically as follows:
step 5.1, detecting the signal-to-noise ratio, which is as follows:
under the condition of no interference, the signal-to-noise ratio of the satellite signal is higher than a set value; under the condition that satellite signals are interfered by deception signals, the change of signal to noise ratio is as follows:
where SNR' represents the signal-to-noise value, P, of the receiver after being subjected to a spoofing signal s For the power of the real satellite signal, P i To deceive the power of the signal N 0 Power spectral density for system noise;
the other receiver receives the deceptive signal and changes the signal-to-noise ratio as follows:
where α is a parameter related to the distance between the receiver and the satellite;
under the interference of the deception signal, the signal-to-noise ratio of the satellite signal is reduced due to the influence of the deception signal power, and the higher the deception signal power is, the lower the signal-to-noise ratio is, namely alpha P i Is the added value of the satellite signal noise base received by the receiver, and the deception jamming signal can increase the noise base of the received signal;
Setting a corresponding threshold value according to the actual condition of the receiver, and judging whether the receiver is attacked by deception signals or not; the signal-to-noise ratio threshold is set as follows:
in the SNR 0 For the signal-to-noise threshold, SNR i_visible The signal to noise ratio of the ith visible satellite is given, and N is the number of visible satellites;
when the threshold value is exceeded, giving an alarm that the deception jamming is possible, and reducing the confidence of the satellite i according to a formula (10);
step 5.2, detecting the clock difference of the receiver, which is specifically as follows:
the pseudo-range equation of a single satellite after satellite clock correction is as follows:
wherein ρ is i Representing the pseudorange of the ith satellite to the target antenna, (x) s ，y s ，z s ) Is the three-dimensional position of the satellite, (x) u ，y u ，z u ) Is the three-dimensional position of the target object, c is the speed of light,Δt u for the local clock-difference to be the local clock-difference,is the sum of ionosphere, troposphere and other error terms;
after N pseudo-range equations of the formula are obtained, simultaneous equation sets and linearization processing are carried out to obtain:
G·ΔX u ＝b (19)
wherein DeltaX is u ＝[Δx u Δy u Δz u Δt u ] T Is a state variable;
the least squares solution of equation (19) is:
Δx 0 ＝(G T G) -1 G T b (22)
due to Deltax 0 ＝[Δx u Δy u Δz u Δt u ] T The clock difference delta t can be obtained through the method (22) u ；
The common error of the pseudo-range observed quantity can influence the clock difference, the submersion of the deception signal can cause the fluctuation of the clock difference, and the deception signal is detected according to the fluctuation; because the variable is directly used as the judgment basis of the deception signal, a plurality of judgment criteria are needed to prevent the outlier in the data from influencing the detection process:
Δt u_pre0 -Δt u_pre1 ≤2×(Δt u_pre1 -Δt u ) (23)
Wherein Δt is u_pre0 、Δt u_pre1 Respectively representing the clock difference value of the previous moment and the previous moment, and if the clock difference jump amplitude of the current moment is larger than a set value, pre-warning and storingΔt u_pre1 Is a new Δt u_pre0 For continuing to compare the jump amplitude at the next moment, and when the jump amplitude exceeds twice, the alarm possibly suffers from attack of deceptive jamming signals; under normal conditions, the clock error jump has upper and lower limits, a wild value exists at part of the time, when an interference signal appears, the clock error jump amplitude exceeds the range of the upper and lower limits, at the moment, the occurrence of the interference signal is judged, and pseudo-range rate detection are carried out to confirm the existence of a deception signal;
and 5.3, according to the pseudo-range and pseudo-range rate information, realizing inertial auxiliary pseudo-range and pseudo-range rate consistency detection, wherein the method comprises the following steps of:
by analyzing the change of the pseudo range and the pseudo range rate, judging whether the satellite signal is interfered and deceptively, wherein the formula is as follows:
wherein ρ is u 、Pseudo-range and pseudo-range rate, ρ, of antenna relative to satellite output by GNSS receiver m 、/>Pseudo range, pseudo range rate, deltaρ and/or Deltaρ of carrier relative to satellite calculated by combining inertial navigation information with satellite ephemeris>Errors between the two calculation modes are respectively;
for the ith satellite, its corresponding ρ m 、The calculation method is as follows:
wherein ρ is m_i 、Pseudo-range and pseudo-range rate of SINS relative to satellite i, (x) m ，y m ，z m )、/>Position, speed, < >/of inertial system in ECEF coordinate system respectively>The position and the speed of the satellite i under the ECEF coordinate system are respectively;
normally, the error divergence of the inertial system is smaller than a set value, deltaρ,The values are all smaller than the set value; however, when the abrupt deception signal exists, the position, speed or time information of the receiver is biased, and the pseudo range rate calculated by the receiver are biased; according to Δρ, ++>The setting method refers to the formula (8) and the formula (9), and when the two are detected to exceed the set threshold value, the interference of the deception signal is judged;
when the delayed type deception signal exists, the periodic enhancement detection is designed, the characteristic of high precision in a short time of inertial navigation is utilized to periodically construct the calculated pseudo range rate of the inertial navigation speed, and the calculated pseudo range rate of the satellite navigation result is utilized to construct the statistical test quantityRealizing slow change type deception detection; since inertial navigation position information is obtained by integrating velocity information, errors in position information accumulate faster, and pseudo-range information ρ calculated from position in equation (24) m The error accumulation of (a) is faster than the pseudo-range rate informationTherefore, adopting the consistency detection of the pseudo range rate to carry out the statistics of the detection quantity on the pseudo range rate sequence;
deriving a satellite pseudo-range equation of formula (18), and obtaining a pseudo-range rate equation, wherein the formula is as follows:
in the middle ofPseudo-range rate for the ith satellite, +.>Variable value representing the distance between satellite and receiver, < >>Representing the receiver clock drift +.>Representing the ionosphere, troposphere delay rate of change, which is less sensitive to time-change and is therefore omitted;
will beExpressed as the satellite to receiver motion velocity vector difference, the formula (27) is:
in the formula, v i The motion velocity vector of the ith satellite is v is the motion velocity vector of the receiver; when the original satellite signal is processed by the spoofing interference source to generate a spoofing signal or directly forward, the speed vector inconsistency can cause jump of the pseudo range rate of the actually acquired satellite signal due to change of the signal direction, so that the statistical test quantity of the residual error of the pseudo range rate of the receiver is constructed as follows:
in the method, in the process of the invention,theoretical residual error representing pseudo-range rate between receiver and satellite i, and using inertial navigation speed information short-time high-precision characteristic to calculate pseudo-range rate +.>Substitution of theoretical pseudo-range rate- >
In the inspection period, when the residual exceeds the residual threshold caused by the set inertial navigation drift, determining the signal of the satellite as a deceptive signal, calculating the mean value and standard deviation of the residual, and identifying abnormal values by using the statistical data, wherein the formula is as follows:
in the method, in the process of the invention,representing the mean of the residual, s ∈ Represents the standard deviation of the residual error, n represents the number of samples, ε i A residual representing the ith observation; based on the nature of the normal distribution, about 99.7% of observations will be at 3s ∈ Thus when the observed value exceeds 3s ∈ Is determined to be an outlier.
8. The method for detecting and resisting interference of GNSS/SINS integrated navigation signals in a satellite disturbance environment according to claim 7, wherein in step 6, the influence degree of inertial navigation on the receiver acquisition signal power and the tracking threshold is adjusted by using an inertial auxiliary loop technology, so that the interference-to-signal ratio reaches the system requirement, specifically as follows:
in the inertial auxiliary capturing process, the Doppler frequency shift, the frequency shift and the code shift caused by the Doppler frequency shift and the pseudo-range information are calculated by utilizing the position, the speed and the acceleration information output by the inertial navigation system and combining the position and the speed information of the satellite, and the calculated Doppler frequency shift is calculated PN code frequency offset τ and code shift Δτ dyn And respectively transmitting the signals to a numerical control oscillator NCO, a voltage-controlled oscillator VCC and a code generator PN, and setting initial values of code phase and Doppler frequency shift by using time positions provided by GNSS/SINS to finish the capturing process.
9. The method for detecting and resisting disturbance of integrated navigation signal of GNSS/SINS in satellite disturbance according to claim 8, wherein the inertial auxiliary loop has the following structure:
the code loop adopts a carrier-assisted second-order delay phase-locked loop DLL; the carrier wave loop adopts an inertial auxiliary third-order phase-locked loop (PLL); a first order low pass filter is built in the inertial auxiliary partThe loop filter of the third-order phase-locked loop PLL is selected as follows:
obtaining external frequency auxiliary deviation through speed deviation and clock frequency deviation estimated by a navigation filter, adding a reference input signal, and obtaining external frequency auxiliary information through an inertial filter after differentiation;
the phase error output by the phase discriminator is the difference value between the reference input signal and the phase of the output signal of the external phase noise and voltage-controlled oscillator;
the phase error is filtered by a loop filter of a third-order phase-locked loop (PLL), added with external frequency auxiliary information and integrated to obtain the output signal phase of the voltage-controlled oscillator.
CN202311442073.1A 2023-10-31 2023-10-31 GNSS/SINS integrated navigation signal detection and anti-interference method under satellite disturbance environment Pending CN117288186A (en)

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