1. Introduction
Today’s global navigation satellite systems (GNSS) have excellent performance and high positioning accuracy in most scenarios. With the support of GNSS constellation, they can satisfy the daily positioning requirements of military and civilian navigation. However, the inherent vulnerability of constellation GNSSs significantly degrade their positioning ability or even make them completely unable to locate under extreme conditions, such as during wartime. To obtain locations precisely in emergency environments, a rapidly deployed navigation and positioning system is imperative. Single-satellite systems show apparent advantages with fast deployment and low cost compared to GNSSs. Thus, it is an essential supplement to GNSSs. Existing single-satellite positioning techniques adopt observation values at different times from the only satellite to locate. The literature [
1] proposed a frequency-measurement-based single-satellite positioning method that utilized the frequency difference of arrival during satellite movement. However, this method has low positioning accuracy and is limited to stationary or small dynamic targets. The literature [
2] found the geometric relationship between location and rate of phase difference, Doppler frequency and satellite ephemeris position to obtain positioning results but it is quite sensitive to the parameters. The literature [
3] also proposed an angular-measurement-based single-satellite positioning method that utilized the angular-velocity relationship between satellite and target to achieve positioning. However, it requires high accuracy of satellite ephemeris. The literature [
4,
5] proposed a Doppler-based single-satellite positioning method that utilized frequency difference to construct the positioning hyperboloid and integrated it with the inertial navigation system (INS) to achieve positioning. However, these methods are more suitable for static positioning, with low positioning accuracy and fast divergence for dynamic targets. The literature [
6] proposed radial acceleration measurement based on a single-satellite positioning method. This method is simple since it only requires radial acceleration but, as the data source is limited, its reliability is poor. The literature [
7] proposed a beam-scanning single-satellite positioning method with high positioning accuracy for static targets. However, it works in active location mode, which needs interaction with the data center. The literature [
8] also proposed a BP neural-network-based single-satellite positioning method. The positioning results were stable but it takes a lot of time to perform the learning and updating process, so practical applications are limited. Researcher [
9] also proposed an INS-based Doppler integration single-satellite positioning method that has a good positioning effect for both dynamic and static targets. However, since the final filtering result depends on the least-squares solution result, the stability of this method is limited. In the above algorithms, the main method is to transplant the GPS positioning method directly to the single-satellite positioning system. The problem of keeping the target immovable in single-satellite positioning systems is neglected. To answer this question, an inertial-navigation-aided single-satellite highly dynamic positioning algorithm is proposed. First, we computed and compensated the displacement of the moving target by INS [
10]. Then, we utilized the integral Doppler pseudorange to construct an equal-frequency cone surface and use the height information of the moving target to establish the height surface. Next, we obtained the coarse positioning solution with these surfaces. Finally, the displacement value provided by INS was calculated and corrected in the UKF filter to improve the single-satellite positioning result.
2. Single-Satellite System Model
Under the conditions of GNSS denial, a single-satellite positioning system could be an effective complement [
11]. Most targets, such as vehicles, personnel, ships and aircraft are dynamic, that is to say, their track, direction and movement time are uncertain. Therefore, in this paper, inertial navigation is the motion of the target that must be compensated by the INS. The principle of the inertial-navigation-aided single-satellite positioning method is shown in
Figure 1.
Here,
denotes the satellite position at time
i,
denotes the pesudorange between satellite and target at time
i and
denotes the position of the target at time
i. Then, the pesudorange at the three successive moments, as shown in
Figure 1, can be described as
where
represents the error caused by the receiver-clock offset.
With the aid of inertial navigation [
12], we can obtain the position of the target at these three successive moments as
where
represent the displacement of the target by the INS from the
ith time point to the
th time point, respectively. we substituted (4) and (5) into Equations (2) and (4) and the pseudorange between the dynamic targets and the satellites is given in simplified form:
There are four unknowns in the three above equations, so we could not settle the position [
13]. In addition, we included the height measurement that was directly obtained from the altimeter at time
i to determine the target position, since it satisfies
The initial estimation of the target position at time point
ith can be obtained by a least-mean-square algorithm. Based on the Taylor expansion, the estimation problem can be described as
where
and
Equation (
10) contains four unknown parameters. The least-squares method was utilized to jointly solve the dynamic results at each iteration [
14],
, as follows:
By repeating iterations until increment met a certain accuracy threshold, we could obtain the initial estimation of the target’s three-dimensional coordinates.
3. Navigation-Data Fusion
Considering the complex and changeable motion of the target and the inherent error of the INS, the initial estimation of the target position from the previous section needs to further be improved [
15]. The traditional Kalman filter is based on the linear model, shows high positioning fluctuations and suffers divergence in sudden changes. Therefore, in this section, we present an unscented Kalman filter (UKF) -based inertial-navigation-aided single-satellite highly dynamic positioning algorithm.
Suppose is a l dimensional random variable that follows Gaussian distribution, that is, , where is an l dimensional vector and is a symmetric semipositive definite square matrix. is decomposed by Cholesky decomposition as , where is the lower triangular matrix. A new posterior random variable is obtained through nonlinear transformation.
In this case, unscented transformation (UT) is utilized to estimate the mean value and the covariance matrix of
. Choosing a set of
sampling points that meet the selection scheme as
where
and
represent the main scale parameter and
represents a minor scale parameter;
a indicates the extent to which the sampling point deviates from the expected value, which is usually a very small positive number. It can be seen that all
and
are symmetrical about
, so weight factors can be defined as
The predicted state mean and covariance matrix can be obtained through the weighted sum.
Propagating each sampling point through nonlinear function,
Then, the predicted measurement mean and covariance matrix can be obtained through the weighted sum.
The UKF replaced the traditional EKF to solve nonlinear problems through unscented transformation. This transformation approximates the probability of distribution rather than the nonlinear function and avoids calculation of the Jacobi matrix. It provides estimation accuracy approximately to the third-order EKF, equally computationally intensive with the first-order EKF. In a single-satellite system, the decremental of computation load has great significance for performance improvement in highly dynamic positioning.
In the inertial-navigation-aided single-satellite system, the augmented state is denoted as
where
,
and
represent the position, velocity and acceleration of the positioning target at time
t, respectively. The state-space equation model to describe the dynamic process of targets:
where,
denotes a dynamic matrix, dynamic matrix
represents speed variation and
represents acceleration variation;
denotes single-satellite-system noise at time t. In practice, the obtained samples are discrete values. The state-transition matrix obtained by discretizing
can be regarded as nontime-varying when the sampling interval is small.
The prediction and update steps of UKF are looped after initialization. Assuming that the state of a single-satellite system is a random variable , with expected value and covariance matrix , by supplying that UKF a prior estimate and covariance , the next step is to build a state-transition matrix . Let represents the sampling interval and represents the following state-transition matrix.
State quantity was utilized to construct the sampling points and calculate the unscented weight coefficients; represents the predicted estimate at sampling interval t and represents the posterior estimate at sampling interval t. In the time-update part of single-satellite positioning, the state transformation of the generated test point set is first carried out. Since UKF guarantees second-order accuracy, nonlinear transformation function can be rewritten as
The predicted estimate of
is:
Construction of measurement equation vector
by using a pseudorange and height values at
t time:
Similar to the least-squares part, the state variable substituting value is the target-state value at time; are the relative displacements of the three-dimensional coordinates between time i and , recorded by the inertial navigation system; represent the coordinates of the single satellite at time i; represents the actual pseudorange value between single satellite and target at time i; represents the estimated pseudorange value at time i; represents the target-height value at time i; represents the estimated target-height value at time i. Here, was utilized to calculate the pseudorange value of at the corresponding time.
All symbols in Equations (36–39) are scalars. For example,
represents the
kth column of the solution of vector
at time
i;
,
and
represent the
value of the
kth column of solution of vector
at time
. Then, predictive value
of
is as follows:
UT transformation was utilized to obtain a priori covariance matrix
In order to make the dynamic positioning prediction accurate, initialization was only performed once and the measurement-update part was as follows:
In order to make the dynamic positioning prediction accurate, initialization was only performed once and the measurement-update part was as follows:
The Kalman filter gives the estimated value at the new time point according to the a priori estimate. The of the state variable indicates the prediction estimate, right label-indicates the a priori value, indicates the Kalman filter coefficient matrix at time t and and are the prediction covariance error matrix of the single-satellite navigation system. The computational complexity of the UKF filter was smaller than that of the EKF with the same accuracy and tracking performance was also better than EKF. It is very suitable for dynamic-target location and has good robustness. Combining INS input compensation and the initial least-squares calculation value, the UKF filter can provide stable rough positioning for a considerable period of time and can meet the requirements of single-satellite navigation.