RU2509041C1 - Method of defining spacecraft orbit - Google Patents

Abstract

FIELD: aircraft engineering.

SUBSTANCE: navigation parameters of spacecraft orbit are measured with the help of ground measuring stations. Said measured parameters are transmitted to processing center. Said measured parameters are converted, time moments of said measurements are recorded as measurements program while navigation parameter magnitudes measured by ground stations are used as measurements results. Converted navigation parameters are preprocessed to define orbit parameters by converted said parameters and vector of initial approximation of target parameters.

EFFECT: higher reliability of definition.

1 dwg, 2 tbl

Description

FIELD OF THE INVENTION

The invention relates to monitoring systems for the flight of spacecraft and can be used to determine the orbit of a spacecraft (SC) from the navigation measurements of its parameters.

The prior art.

There is a method of determining the parameters of the orbit of a spacecraft, which consists in measuring the trajectory parameters, transferring to the control complex the totality of the measured values of the trajectory parameters with their subsequent accumulation and processing [1]. According to the invention, four steps of processing path data are used. After the first of them, data on the orientation of the orbital plane are stored and filtered along with similar data accumulated over an interval of several tens of days. The exact parameters of the orbit plane are determined using them in the next step, where estimates of the four parameters of motion in the orbit plane are obtained, which are stored and filtered along with similar data accumulated over an interval of several days. They determine the exact values of the motion parameters in the plane of the orbit. The disadvantages of the method are:

- the need for multi-stage processing of the measured trajectory parameters, which increases the duration of the calculations to determine the parameters of the orbit;

- the impossibility of determining the parameters of the orbit in the event that the initial approximation vector of the desired parameters is outside the narrow boundaries of the allowable region for finding the initial approximation parameters.

A known method for determining the parameters of the orbit of a spacecraft [2], the purpose of which is to increase the accuracy and reliability of determining the parameters of the spacecraft motion. The technical result is achieved by increasing the number of specified motion parameters used to evaluate the orbital trajectory.

The disadvantage of this method is the significant dependence of the solution on the accuracy of the initial approximation of the desired parameters and the impossibility of determining the orbit if the initial approximation of the desired parameters is outside the boundaries of the allowable region.

A known prototype method for determining the parameters of the orbit of the spacecraft, in which they develop a program for measuring the navigation parameters of the orbit of the spacecraft for ground-based measuring stations, measure the navigation parameters of the orbit using ground-based measuring stations, transmit the measured navigation parameters to a computer center in which preliminary processing of the results of measurements of navigation parameters, determine the orbit parameters from the measured navigation parameters and the initial approximation vector, and parameters of the orbit, while the time points for measuring the navigation parameters are used as the measurement program, and the values of the navigation parameters measured by ground stations as the measurement results [3, p.170-172], [4, p.302-303] .

When using the described method for determining the orbit, the following types of navigation parameters can be measured: oblique range, radial speed, orientation angles of the line of sight [5, p.149].

In the described method, when processing the measurement results, the navigation parameters are recorded as functions of time. The main source data used to determine the orbit parameters of the spacecraft are the results of measurements of the navigation parameters and the initial approximation of the desired parameters [3, p. 171], [6, p. 39], which include:

- a program of measurements of navigation parameters of the orbit, presented in the form of a set of consecutive moments of time at which measurements of one or more types of navigation parameters of the orbit were performed;

- a set of measured navigation parameters, each of which corresponds to a value of its time;

is the vector of the initial approximation of the desired orbital parameters, which is the values of the orbital parameters describing the spatial motion of the spacecraft, and the corresponding time instant.

The vector of initial approximation of the sought-for orbital parameters does not describe the orbital motion accurately enough. Therefore, when determining the parameters of the orbit, it is necessary to refine it according to the results of navigation measurements.

The problem of determining the parameters of the orbit belongs to a wide class of inverse problems - the problems of estimating the parameters of a system, in those cases when the estimated parameters of the system are not available to direct measurement. The possibility of finding a solution to such problems depends on the error with which the initial approximation of the estimated parameters is given. Thus, the possibility of determining the orbit parameters from the measurement results depends on the error of the initial approximation vector of the desired orbit parameters. With large deviations of the initial approximation vector from the values of the required parameters (in other words, from the actual values of the orbit parameters), the orbit parameters cannot be determined even if there are the required number of measurements of the navigation parameters made with high accuracy [6, p. 42].

In practical problems of determining the orbit, large deviations of the parameters of the initial approximation vector can occur in the following adverse cases:

1) after the stage of launching the spacecraft into orbit, when the parameters of the obtained orbit significantly differ from the a priori specified orbital parameters;

2) for long intervals between navigation definitions of the parameters of the orbital motion, when the actual values of the parameters of the orbit significantly deviate from the predicted values due to the influence of factors of the space environment, not taken into account in the mathematical model of the motion of the spacecraft.

The disadvantage of the described prototype method is the inability to determine the parameters of the orbit of the spacecraft from the measured navigation parameters if the initial approximation vector is outside the boundaries of the allowable region for finding the initial approximation parameters. It is important to emphasize that the boundaries of the admissible region are determined by the used method of determining the orbit.

Disclosure of the invention.

The problem to which the claimed invention is directed, is to provide the ability to determine the parameters of the orbit with significant deviations of the parameters of the initial approximation vector from actual values, in other words, from the desired values.

The technical result achieved by the claimed invention is to increase the reliability of the technological process of determining the orbit with a significant deviation of the initial approximation vector from the desired orbit parameters due to a significant expansion of the boundaries of the initial approximation parameters.

The basis of the invention is a new procedure for processing measurements of navigation parameters of the orbit. The new set of actions proposed in the method provides for the conversion of the results of navigation measurements and provides a significant reduction in the accuracy requirements for setting the initial approximation of the desired orbit parameters.

The essence of the invention lies in the fact that to achieve the above technical results in the method of determining the orbit, a program is developed for measuring the navigation parameters of the orbit of the spacecraft for ground-based measuring stations, the navigation parameters of the orbit of the spacecraft are measured using ground-based measuring stations, the measured navigation parameters are transmitted to a computer center, which carry out preliminary processing of the measurement results of navigation parameters, determine the orbit parameters from the measured n in the navigation parameters and the initial approximation vector of the desired orbit parameters, according to the invention, during preliminary processing and determining the orbit, the timing of the measurement of the navigation parameters is used as the measurement results, and the values of the navigation parameters measured by ground stations are used as the measurement program.

The essential features characterizing the invention.

1) Changing the procedure for processing the measurement results of the navigation parameters of the orbit, namely, that the moments of time at which measurements were taken, when determining the parameters of the orbit, are used as the results of measurements, and the corresponding navigation moments of the orbit, actually measured by ground stations, are used as a program of measurements.

2) Perform the following set of sequential actions to determine the parameters of the orbit:

a) compilation of a navigation measurement program, during the development of which time intervals are planned for measurements by each measuring station;

b) measure the navigation parameters of the orbit using measuring stations in accordance with the program of navigation measurements and obtain the measured values of the navigation parameters as a function of time;

c) transmit the measured navigation parameters and the corresponding time points to the computer center;

d) in the computing center, the results of the measurements are converted and the navigation parameters of the orbit measured by the ground stations are recorded as a program of measurements, and the time points corresponding to the navigation parameters are recorded as the measured values of the navigation parameters of the orbit;

e) carry out preliminary processing of the converted navigation parameters of the orbit, while the measured values, in other words, the values of the function, use time instants, as an independent argument - the values of the linear or speed parameter of the trajectory, actually measured by the ground station;

f) determine the parameters of the orbit according to the results of the converted navigation parameters of the orbit that have undergone preliminary processing, and the vector of the initial approximation of the desired parameters of the orbit.

The implementation of paragraph "g" of the described sequence of actions significantly distinguishes the proposed set of actions from the sequence of the prototype method. Due to the fact that, in accordance with paragraph "d", the conversion of the measurement results is carried out, in the claimed method the algorithms for the preliminary processing and determination of the orbit differ from the corresponding algorithms of the prototype method.

The content of the conversion of the results of navigation measurements when performing the sequence of actions of the claimed method allows the following formalized interpretation. The measurement results of the navigation parameters of the orbit obtained by the measuring stations will be presented in the form of a set Z

$$\text{Z}={{\text{[z}}_{\text{one}}{\text{(t}}_{\text{one}}),...,z_j(t_j),...,z_m(t_m)]}^T,\left(\mathrm{one}\right)$$

where z _j is the measured value of the navigation parameter; t _j is the time of the jth measurement, m is the number of measurements.

The measurement program will be presented in the form

$$P={[t_{\mathrm{one}},...,t_j,...,t_m]}^T,\left(2\right)$$

When using the prototype method, the determination of the orbit is carried out by sequential iterations, and the calculation is reduced to minimizing the least squares functional of the form

$${\displaystyle \sum _{j=\mathrm{one}}^m{[z(t_j)-z_p(t_j,X_k)]}^2{\omega }_j\to \underset{X}{\min ,}}\left(3\right)$$

where z _p (t _j , X _k ) is the calculated value of the parameter z _j (t _j ) in some calculated orbit given by the vector of parameters X _k , ω _j is the weight of the measurements; X _k is the vector of the orbit parameters corresponding to some iteration k = 0, 1, 2 ...; X ₀ - the vector of the initial approximation of the desired parameters of the orbit.

When using the inventive method, the results of navigation measurements are converted, and the measurement results are recorded as a set of Z *

$$Z*={[t_{\mathrm{one}}(z_{\mathrm{one}}),...,t_j(z_j),...,t_m(z_m)]}^T,\left(\mathrm{four}\right)$$

The converted measurements correspond to the measurement program, which we write in the form

$$P*={[z_{\mathrm{one}},...,z_j,...,z_m]}^T,\left(5\right)$$

In the claimed method, determining the orbit reduces to minimizing a functional of the form

$${\displaystyle \sum _{j=\mathrm{one}}^m{[t(z_j)-t_p(z_j,X_k)]}^2{\omega }_{{}_j}^{*}\to \underset{X}{\min ,}}\left(6\right)$$

- веса измерений.where t (z _j ) is the value of the measurement time of the parameter z _j ; X ₀ - the vector of the initial approximation of the desired parameters of the orbit; t _p (z _j , X _k ) is the calculated time value corresponding to the parameter z _j in the orbit X _k ; $${\omega }_j^{*}$$

- weight measurements.

When using the prototype method and functional (3), determining the parameters of the orbit reduces to minimizing the mismatch between the two values of the navigation parameter - measured and calculated, i.e. values corresponding to the same moment in time, one of which refers to the real, and the other to the calculated orbit.

When using the claimed method and functional (6), the determination of the parameters of the orbit is reduced to minimizing the mismatch between the two time instants of reaching the same value of the navigation parameter, while one moment corresponds to the actual orbital movement, the second to the movement in the calculated orbit.

Confirmation of the possibility of obtaining the claimed technical result when using the proposed method was obtained by conducting numerous experimental calculations. The calculation results confirmed a significant expansion of the boundaries of the allowable region for finding the initial approximation parameters when using the claimed method. The calculations are based on the results of field measurements of the navigation parameters of the orbit.

An example confirming a significant expansion of the boundaries when using the proposed method compared to the prototype method is shown in tables 1 and 2. The tables show the boundaries of the allowable region for finding the initial approximation of the prototype method and the claimed method.

The bounds are estimated based on the measurement of the radial velocity of the spacecraft, located in a circumcircular orbit 1000 km high with a draconic period of revolution T _d = 104 min. The measurements were performed relative to ground-based measuring stations over a flight interval of 25 minutes. The required spacecraft orbit parameters are set by Keplerian elements:

$$\begin{array}{cccc}{t}_0=0& a=7362\mathrm{to}m;& e=0.00275;& i=82.96gR\mathrm{but}d;\left(7\right)\\ & \Omega =156.45gR\mathrm{but}d;& v=128.72gR\mathrm{but}d;& \omega =231.27gR\mathrm{but}d.\end{array}$$

Table 1 presents the minimum and maximum values of the boundaries of the initial approximation parameters corresponding to the prototype method, and in table 2 to the claimed method. In columns 2 and 3, the values of the boundaries are shown, and in parentheses are the deviations of the boundary values from the values of the parameters of the vector (7). The last line shows the values of the permissible deviation of time t ₀ at which the initial conditions are specified.

Table 1 Assessment of the boundaries of the prototype method Orbit elements Boundary Estimates Min values Max. value one 2 3 a km 6999 (-363) 7742 (+380) e 0 (-0.00275) 0.06 (+0.055727) i, hail 71.96 (-11) 100.97 (+18) Ω, deg 146.45 (-10) 168.45 (+12) v, hail 119.73 (-9) 139.72 (+11) ω, deg 222.27 (-9) 242.27 (+11) t ₀ min -3 (-3) +2 (+2)

table 2 Assessment of the boundaries of the claimed method Orbit elements Boundary Estimates Min values Max. value one 2 four a km 6502 (-860) 10252 (+2890) e 0 (-0.00275) 0.12275 (+0.12) i, hail 30.96 (-52) 103.96 (+21) Ω, deg 141.45 (-15) 189.45 (+33) v, hail 69.72 (-59) 383.72 (+255) ω, deg 173.27 (-58) 481.27 (+250) t ₀ min -74 (-74) +17 (+17)

From tables 1, 2 it is seen that the boundaries of the region of acceptable values of the claimed method significantly exceeds the boundaries of the prototype method. So, for example, from the last rows of the tables it follows that the claimed method allows you to determine the orbit when deviating the time of binding of the initial conditions in the range of values [-74, +17] min, and the prototype method with significantly smaller boundaries of the allowable range of values - [-3, +2] min

The block diagram of a system designed to determine the orbit in accordance with the claimed method, is presented in figure 1.

The system includes a unit for developing a navigation measurement program 1, ground measuring stations 2, means for transmitting the measurement results of navigation parameters 3, a unit for storing the results of navigation measurements 4, a unit for converting the results of navigation measurements 5, a preprocessing unit 6, a unit for determining orbit parameters 7, a unit for storing parameters orbits 8. The composition of the computing center, which performs the processing of navigation measurements, includes blocks 1, 4, 5, 6, 7, 8, which are combined in the drawing paragraph irnoy line. In this case, the first input of the orbit parameter determination unit 7 is connected to the output of the pre-processing unit 6, the second input of the orbit parameter determination unit 7 is connected to the second output of the orbit parameter storage unit 8, the first output of the orbit parameter storage unit 8 is connected to the input of the development block of the navigation measurement program 1 .

The system operates as follows.

The development block of the navigation measurement program 1 receives the orbit parameters coming from block 8, compiles a program for measuring navigation parameters and transmits the program to ground-based measuring stations 2, which measure the navigation parameters of the orbit in accordance with the measurement program. Means of transmitting the results of navigation measurements 3 transmit measurements to a computer center for processing navigation information, which are received in the storage unit for the results of navigation measurements 4.

The conversion unit for the results of navigation measurements 5 converts the measured navigation parameters in accordance with the procedure of the claimed method and records the times of measurement of the navigation parameters as measurement results, and the values of the navigation parameters measured by ground stations as a measurement program. Block 6 pre-processes the converted results of navigation measurements, which are received after processing in block 7.

In the block for determining the parameters of the orbit 7 receives the values of the vector of initial approximation of the desired parameters of the orbit from block 8 and the converted results of navigation measurements from block 6. Then block 7 determines the parameters of the orbit from the values of the vector of initial approximation and the converted results of navigation measurements. The calculated orbit parameters are received for storage in block 8.

Industrial applicability

The main technical result achieved by the claimed invention is to increase the reliability of the successful implementation of the technological cycle of determining the orbit according to the results of navigation measurements with a significant deviation of the initial approximation vector from the desired parameters of the orbit. The result is achieved by expanding the boundaries of the admissible region for finding the initial approximation parameters.

An advantage of the claimed invention is that its implementation in spacecraft flight control practice does not require measuring, firstly, the usual organizational measures that are currently being performed when measuring navigation parameters of the orbit, and secondly, the operating procedure of measuring stations and the actions of maintenance personnel. It is necessary to apply a new procedure for performing actions when processing the results of navigation measurements in a computer center.

The proposed method provides the ability to determine the parameters of the orbit of the spacecraft from the measured navigation parameters if the initial approximation vector is outside the narrow boundaries of the allowable region for finding the initial approximation parameters.

From the sequence of actions necessary for the implementation of the method it follows that the claimed method can be used to determine the parameters of the orbit of spacecraft and repeatedly reproduced.

Literature

1. Patent No. 2150414 RU, IPC7 B64G 3/00, G01S 3/42. The method of determining the parameters of the orbit of the spacecraft / Denisov K.I., Wompe A.A., claimed 01.02.1999, Publ. 06/10/2000.

2. Patent No. 2391265 of the Russian Federation, IPC B64G 3/00, G09B 23/00. The system for assessing the accuracy of determining the parameters of the motion of a satellite / Mikhailova V.P., Nemtsov V.I., Sadovnikov O.G., announced June 15, 2009, Publ. 05/10/2010.

3. Navigation support for the flight of the Salyut-6 orbital complex - Soyuz - Progress / I.K. Bazhinov, V. P. Gavrilov, V. D. Yastrebov et al. - Moscow: Nauka, 1985.

4. GLONASS. The principles of construction and operation. / Ed. A.I. Petrova, V.N. Kharisova. - M .: Radio engineering, 2005.

5. Ivanov N.M., Lysenko L.N. Ballistics and spacecraft navigation. - M.: Bustard, 2004.

6. Brandin V.N., Razorenov G.N. Determination of the trajectories of spacecraft. - M.: Mechanical Engineering, 1978.

Claims (1)

A method for determining the orbit of a spacecraft, in which a program for measuring the navigation parameters of the orbit of the spacecraft for ground-based measuring stations is developed, the navigation parameters of the orbit of the spacecraft are measured using ground-based measuring stations, the measured navigation parameters are transmitted to a computer center in which preliminary processing of the measurement results of the navigation parameters is carried out, determine the orbit parameters from the measured navigation parameters and the initial approximation vector of the desired pa orbit parameters, characterized in that during pre-processing and determining the orbit, time moments of measuring navigation parameters are used as measurement results, and values of navigation parameters measured by ground stations are used as a measurement program.

Images