Method of orientation and independent navigation for space vehicle intended for earth and near space monitoring

FIELD: aerospace engineering.

SUBSTANCE: astrohardware and Earth sounding hardware data is used to accurately determine spacecraft orbit and coordinates and orient it. Sounding data, apart from its designed purpose, is subjected to additional processing to be made in separate frames of the Earth image. This allows determining direction to the Earth center ("to plot local vertical"), correct current spacecraft coordinates and its orientation. Coordinates derived from combination of the Earth sounding area frames on one and several orbital passes are processed together to produce orbit parametres. Orbit parametre estimations are used for navigation-ballistic computations effected aboard the spacecraft till updating at another flight path. Note here that to do with aforesaid tasks, spacecraft needs no usual navigation data from ground appliances or satellite navigation systems, e.g. GLONASS. It does not required special means of orientation relative to the Earth centre (or horizon), say, infrared vertical. Model analyses have shown good chances of providing high accuracy in spacecraft current orientation and coordinate data on high-elliptical or geo-stationary orbit.

EFFECT: expanded operating performances in independent pilotless operation.

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The present invention relates to space and can be used in the design, construction and operation of a global information space monitoring of the Earth and near-earth space, used for high-precision guidance and navigation separate SPACECRAFT (SC) and groups of several mutually related SPACECRAFT operating in Autonomous automatic mode in the composition of the monitoring systems on highly-elliptical and geostationary orbits, preferred for continuous global observations of the Earth and near-earth space [1].

It is known that joint flight with the spacecraft (SC) "Union" on Board the spacecraft "Apollo" along with the use of navigation data received from the ground, was a task of Autonomous navigation, i.e. in the on-Board hardware QC was determined orbital parameters (and coordinates QC at a given point in time) without any interaction with the hardware other space or ground-based complexes or systems. The way of Autonomous navigation, and orientation of spacecraft "Apollo", which may be regarded as analogous to, is the following [2].

During the flight QC astronauts periodically determine the orientation of the axes of the platform in stellar inertial space. the La this astronaut using the handlebars expands QC so, to approve the optical system of the telescope sextant chosen two (or more) of the navigation star. By clicking astronaut issues a command to register in the on-Board computers (computer) data, determining the current orientation of the platform in the inertial coordinate system, the rotation angle of the optical system and the position of the axes of the platform with respect to the axes QC, as well as the time centering of the star in the sextant. Glioblastoma retains its orientation in inertial stellar space, the withdrawal does not exceed 0.1° per day. The desired orientation of the spacecraft in inertial space support, using the data on the misalignment of the axes CC and platform. For orbital or another connected with the Earth orientation attract navigation information received from the ground control Center, flight, or generated off-line on-Board QC. For solving the problem of Autonomous navigation astronauts in several points of the orbit using the onboard optical system to measure the height of the navigation stars above the visible horizon of the Earth or the angles between the directions on the stars and on targets with known coordinates on the Earth's surface (benchmarks). Measurement data is processed in the computer, and determine the current orbital parameters QC, which are used for navigation and ballistic calculations.

Insufficient the om method orientation and Autonomous navigation of spacecraft "Apollo" is that its implementation requires the participation of the operator in space. This eliminates the possibility of using this method on Board unmanned SPACECRAFT tasked with monitoring systems of the Earth.

The closest in technical essence (selected as a prototype) is, in our opinion, the method of orientation and navigation series of SPACECRAFT "meteor-M"designed for remote sensing of the earth and near-Earth space [3]. The essence solved by the prototype tasks is as follows.

With a given periodicity from the onboard astropart receive measured values of angles φibetween the construction axes AC and directions assigned to the navigation stars whose coordinates on the celestial sphere (right ascension

αiand declination δi) is known. At the subscriber side equipment of satellite navigation (ASN) of the GLONASS system periodically executes the determination of the coordinates and velocities of the motion of the SPACECRAFT in connected with the Earth coordinate system is the vector of initial conditions X(tn). The received vector X(tn) is used for navigation and ballistic calculations, including determination of SPACECRAFT coordinates α, δ at a given time t. Using the coordinates KA α, δ and navigational stars αi, δiand the angles φibetween building axes of the SPACECRAFT and the directions on the navigation of the stars, determine the current time t, the orientation angles of the construction of the axes KA associated with the Earth coordinate system.

The disadvantages of this method, limiting its application to SPACECRAFT monitoring system are:

- increase in the complexity of the onboard hardware KA monitoring system due to the inclusion in its membership of the subscriber's equipment GLONASS;

the dependence of the success of the monitoring objectives from the health of the GLONASS system and the quality (accuracy) of the information received from it, i.e. the absence of the Autonomous functioning of the monitoring system, important in terms of, for example, its military applications and commercial relations;

the GLONASS system is designed for customer service on the Earth's surface and in near-earth space, including satellite orbit, the low and medium height, and is not intended for SPACECRAFT operating on highly-elliptical and geostationary orbits.

The aim of the invention is to solve problems of orientation and navigation-ballistic ensure KA monitoring system of the Earth and near-earth space on geo and HEO orbits, operating in automatic mode in terms of Autonomous unmanned operation.

This goal is achieved due to the fact that h is on the way orientation and Autonomous navigation of SPACECRAFT monitoring system of the Earth to determine the orientation of the SPACECRAFT measured using on-Board Astroparticle angles between the construction axes AC and directions on assigned navigational stars with known coordinates of right ascension αiand declination δiand use the current coordinates of the SPACECRAFT on the celestial sphere αpand δpcalculated using the initial conditions X(tn)intended for navigation and ballistic calculations. To Refine the estimated data of the current coordinates αp, δpand orientation of the SPACECRAFT to determine the exact initial conditions X(tnperform special additional processing information generated by onboard equipment sensing in the process of solving problems for its intended purpose.

When processing information of each individual frame sensing in the j-th row of its sweep allocate the elements of the resolution (with coordinates (j, ij(n)), which correspond to the boundaries of the "space - to-Earth" (j, ij(1)) and "Earth - to-space" (j, ij(2)). To highlight these elements of the permission form in the sliding mode along the j-th row of scan pairs of adjacent samples of limited length {gji}1and {gji}2containing data on the radiation parameters of the gjispace and Earth adopted in the separate elements of the resolution. Estimate the mathematical expectation M1[g], M2[g] and the variance of D1[g], D2[g] the radiation received in the first and second is Minich samples. From the set of pairs of adjacent samples generated in the j-th row, distinguish those in which the difference of the estimates of M1[g] compared to M2[g] and D1[g] D2[g] take extreme (Max, min) values. Coordinate values separating adjacent sample selected pairs, used to denote the coordinates of the boundaries of the "space - to-Earth" (j, ij(1)) and "Earth - to-space" (j, ij(2)). Coordinates j, ij(p)elements of the resolution selected in the set of lines in the frame, remember, accumulate, process, least-squares and determine the coordinates of the j0, i0the center of the Earth, observed in the field. Calculate the difference Δj, Δi measured coordinates of the j0, i0and the coordinates of the center of the Earth j(0)i(0)corresponding to the calculated current SPACECRAFT coordinates αp, δp. Using the difference Δj, Δi correct calculated values of the orientation angles KA (used to calculate the coordinates of the object sensing), and determine corrective amendments to the calculated coordinates KA αp, δpthan determine the actual (measured) coordinates KA αto, δtoat time ttobinding information of the k-th frame sensing of the earth's surface. Measured coordinates KA αto, δtoobtained in the frame-like sensing on the river is a working-class area of the n-th round of flight KA, remember, accumulate and form in the array of measured coordinates (α, δ)n. When forming the array of coordinates (α, δ)nyou can compress the information contained in this array. For this set of coordinates αto, δtoobtained in the frame-like sensing in the working portion of the n-th revolution of the flight SPACECRAFT is divided into subsets corresponding to the zones of smoothing. The resulting optimal smoothing coordinates αm, δmform array (α, δ)ninstead of the coordinates αto, δto. The coordinates of the array (α, δ)nprocess together, using the method of least squares, and define the vector of the current parameters of the orbit R(tnat the time tnthe n-th revolution of the flight SPACECRAFT. This provides the minimum sum of squared weighted residuals (differences) of the processed coordinates α, δ array (α, δ)nand their estimated values αp, δpcalculated using the parameters of the vector R(tnand working a stochastic model of the SPACECRAFT motion, accepted for processing. When determining the weight matrix of residuals sum of the covariance matrix of errors processed SPACECRAFT coordinates α, δ and the corresponding methodological errors working model of the motion of the SPACECRAFT. Applying the parameters of the vector R(tn) create new initial conditions X(tthe used in subsequent navigation and ballistic calculations.

In order to improve the accuracy of the current orbital parameters arrays of coordinates (α, δ)nreceived on the sliding interval q of coils flight KA, remember accumulate and form together arrays of coordinates {(α, δ)n, (α, δ)n-1,..., (α, δ)n-q+1}. Coordinates α, δ from the set of arrays {(α, δ)n, (α, δ)n-1,..., (α, δ)n-q+1} handle together, using the method of least squares, and define the vector of the adjusted parameters of the orbit R(t0). At the same time provide the minimum sum of squared weighted residuals processed coordinates α, δ from the set of arrays {(α, δ)n, (α, δ)n-1,..., (α, δ)n-q+1} and their estimated values αp, δpcalculated using the parameters of the vector R(t0and working in a stochastic model the motion of the SPACECRAFT. When determining the weight matrix of residuals sum of the covariance matrix of errors processed SPACECRAFT coordinates α, δ and the corresponding methodological errors working model of the motion of the SPACECRAFT. Applying the parameters of the vector R(t0) create new initial conditions X(tnused in subsequent navigation and ballistic calculations.

In order to reduce the cost of computing resources required for the precise determination of the parameters of the bits, the vectors R(tn) and their covariance matrix, obtained at the sliding interval q of coils flight KA, remember, accumulate and form the set of vectors {R(tn), R(tn-1),..., R(tn-q+1)}. The parameters of the orbits of R(tn), R(tn-1),..., R(tn-q+1) set of vectors {R{tn), R(tn-1),...,R(tn-q+1)} filter using the least-squares method, and define the vector of exact values of the parameters of the orbit R(tf)that provide the minimum sum of the squares of the weighted residuals of the processed parameters of the orbits of R(tn), R(tn-1),...,R(tn-q+1from the set of vectors {R(tn), R(tn-1),...,R(tn-q+1)} and the estimated values of Rp(tn),

Rp(tn-1),...,Rp(tn-q+1), calculated using the parameters of the vector R(tfand working in a stochastic model the motion of the SPACECRAFT. When determining the weight matrix of residuals sum of the covariance matrix of the errors of the processed parameters of the orbit R(tn), R(tn-1),...,R(tn-q+1) and the corresponding methodological errors working model of the motion of the SPACECRAFT. Applying the parameters of the vector R(tf) create new initial conditions X(tnused in subsequent navigation and ballistic calculations.

Figure 1 shows the sequence of operations for processing data in the proposed method orientation and avtonomniy navigation KA monitoring system of the Earth.

In the diagram (Fig 1) indicated

1 - KA;

2 - onboard equipment monitoring system;

3 is the calculated coordinates of the center of the Earth j(0)i(0)in the field of sensing equipment monitoring system;

4 - known initial conditions X(tn), which can be used for navigation and ballistic calculations including the current coordinates of the SPACECRAFT;

5 - calculation of SPACECRAFT coordinates at the time of sensing;

6 is the calculated current position KA αp, δpat time tksensing;

7 is a side astroparticule;

8 - coordinate αi, δinavigation stars;

9 - calculation of the required values of the angles between the construction axes AC and directions on the navigation of the stars;

10 - calculated values of the angles φipbetween the construction axes AC and directions on the navigation of stars corresponding to the desired orientation of the SPACECRAFT;

11 - measured values of angles φibetween the construction axes AC and directions on the navigation of the stars;

12 - motion control of SPACECRAFT relative to its center of mass;

13 - radiation parameters gjispace and Earth, taken in the i-th element of the resolution of the j-th row scan frame sensing equipment monitoring system;

14 - processing of radiation parameters gjithe separating elements of the resolution corresponding to the boundary of m is waiting for the cosmos and the Earth;

15 - coordinates (j, ij(n)selected elements of the resolution corresponding to the boundary between space and Earth;

16 - processing of the coordinates (j, ij(n)selected items permissions, determining the size and position of the observed shape of the Earth in the box frame sensing;

17 - coordinates of the center of figure of the Earth j0, i0in the field of sensing;

18 - the coordinates of the SPACECRAFT;

19 - measured values of the current coordinates of the KA - right ascension αkand declination δk;

20 - smoothing of the measured SPACECRAFT coordinates αk, δk;

21 - smoothed coordinates KA αm, δm;

22 - array formation SPACECRAFT coordinates, obtained from measurements on one turn of the flight SPACECRAFT;

23 - array (α, δ)nthe SPACECRAFT coordinates on one turn of the flight SPACECRAFT;

24 - processing array of SPACECRAFT coordinates at one stage of its flight, to determine the current settings of the orbit;

25 - vector R(tn) current orbital parameters obtained from measurements on one turn of the flight SPACECRAFT;

26 - replacing the initial conditions X(tn) on their new values, corresponding to the orbital parameters obtained at the current stage of the flight SPACECRAFT in the information processing system onboard equipment monitoring;

27 - the formation of a set of arrays SPACECRAFT coordinates on the sliding interval q turns his flight;

28 is a set of arrays SPACECRAFT coordinates {(α, δ)n, (α, δ)n-1,..., (α, δ)n-q+1}received on the sliding interval q turns his flight;

29 - determination of the adjusted parameters of the orbit of the SPACECRAFT coordinates, distributed on q turns his flight;

30 - vector R(t0revised current orbital parameters obtained by processing the SPACECRAFT coordinates, distributed on q turns his flight;

31 to form a set of vectors of the current orbital parameters obtained at the sliding interval q of coils flight KA;

32 - the set {R(tn), R(tn-1),...,R(tn-q+1)} vectors of the current orbital parameters obtained in the interval q of coils flight KA;

33 - the definition of the vector of exact orbital parameters by filtering a set of vectors of the current orbital parameters, distributed on the interval q of coils flight KA;

34 - vector R(tf) precise orbit parameters obtained by filtering the set of values of its parameters, distributed on the interval q of coils flight KA;

L - logical element in the sequence diagram of the type "if" and "either-or";

P - marking of break lines between elements of the sequence of steps performed to reduce the intersections of the lines on the drawing.

The proposed method is implemented as follows.

Required orientat the Yu building axes KA (1) determine the interests of the guidance field frame sensing equipment monitoring (2) on a given plot of the observed surface of the Earth. In particular, if you set the calculated coordinate values of the j(0)i(0)(3) the Earth's center coinciding with the center of the field frame sensing. Thus the optical axis of the instrument sensing parallel construction axis KA, which should be focused in the center of the Earth along the radius vector of the SPACECRAFT with the coordinates of right ascension α and declination δ on the celestial sphere [4]. To perform the required orientation of the SPACECRAFT using known (existing a priori in the on-Board complex KA) values of the initial conditions X(tn) (4) and count (5) coordinates KA αp, δp(6) the current sensing tk. From the almanac of onboard astropart (7) select and appoint the navigation stars with known coordinates αi, δi(8). Define (9) the calculated values of the angles φip(10) between the directions on the stars and building axes KA corresponding to the desired orientation KA considering its estimated coordinates on the celestial sphere αp, δp(see figure 2). Measured Astroparticle angles φi, (11) due to the operation of the control system (12), which required the reversal of the SPACECRAFT relative to its center of mass, leads to the corresponding calculated values φIPthan provide a desired orientation of the SPACECRAFT.

To Refine the estimated data of the current coordinates and orientation To the perform special processing of information (13), generated by onboard equipment monitoring system (2). Handle (14) the radiation parameters gji(13) of the cosmos and of the Earth, taken in the i-x elements of resolution j-x scan lines of the frame sensing equipment monitoring systems, as follows. In the sliding mode along the j-line scanner frame formed of a pair of adjacent samples {gji}1and {gji}2containing data on the radiation parameters of the gjiin the elements of the resolution of the row numbers (coordinates) from i-h to i-1 for the first sample and the second sample is from i to i+h-1 (see figure 3). According to preliminary estimates of sufficient length samples h can be 10-15 items permissions. Given that the level of radiation received in the elements of the resolution of the frame is a random variable, estimate the mathematical expectation M1[g], M2[g] and the variance of D1[g], D2[g] radiation received in adjacent samples.

Given the sharp differences in the radiation of the Earth and cosmos can then be applied to the next solution. Among the many pairs of adjacent samples generated in the j-line, distinguish those in which the difference in the estimates of M1[g] compared to M2[g] and D1[g] D2[g] take extreme values. Thereby determine the value of the argument i, which in the j-th row may correspond to the boundary of the cosmos - Earth-Christ."

and to match the border of the "Earth - to-space"

The desired coordinates of the selected elements of the resolution in the j-th row corresponding to the boundaries between space and Earth, determine, for example, as average values of ij(n)=0,5(iM(n)+iD(n))if the values of the extrema is not contradictory, i.e

Here Δipthe threshold criterion, taking into account possible errors in the determination of extrema; n=1 corresponds to the boundary of the space-to-Earth", and n=2 is the boundary of the "Earth-to-space" (see figure 3).

The set of coordinates (j, ij(n)elements of the resolution selected in the set of scan lines of the frame, a handle (16) using the method of ordinary least squares (OLS) and determine the average radius ρ0and the coordinates of the center of j0, i0the observed shape of the Earth in the box frame sensing (17). The average radius ρ0together with small tabulated increments, taking into account the compression of the Earth, adequately represents the shape of the Earth as observed in the frame sensing. This ensures the correct solution to the problem of minimizing the sum of squared OLS-residuals of measurements j, ij(n)coordinate boundaries of the observed shape of the Earth and their RA is an even value, obtained using the parameter estimates ρ0I , j0, i0determining the location in the field of sensing and the size of the observed shape of the Earth.

Compare the coordinates of the j0, i0with coordinates of the center of the Earth j(0)i(0)corresponding to the calculated current SPACECRAFT coordinates αp, δpand determine their difference Δj, Δi. Using the difference Δj, Δi adjust the current values of the angles of orientation of the SPACECRAFT used in the calculation of the coordinates of the object sensing, and determine the values of the corrections to the calculated coordinates KA αp, δpadjust and get (18) the values specified (measured) current SPACECRAFT coordinates αk, δto(19) at time ttobinding information k-frame sensing. Measured (corrected) coordinates KA αto, δtoobtained in the frame-like sensing in the working portion of the n-th round of flight KA, remember, accumulate and form (22) in the array of measured coordinates (α, δ)n(23). In the array (α, δ)ninclude data on the covariance matrices of the errors of the coordinates α, δ.

With a large number of scan lines of the frame sensing in order to reduce the volume of the array (α, δ)nand reduction of consumption of computing resources further processing can b the th completed data compression, which should be included in the array (α, δ)n(23). For this set of coordinates αto, δtoobtained in the frame-like sensing in the working portion of the n-th revolution of the flight SPACECRAFT is divided into subsets corresponding to the zones of smoothing, and the resulting optimal smoothing (20) coordinates αm, δm(21) is formed in the array (α, δ)ninstead of the coordinates αto, δto.

The coordinates of the array (α, δ)nhandle (24) together, using OLS, and define the vector of the current parameters of the orbit R(tnat the time tnthe n-th round of flight KA (25). When processing OLS (24) ensure the minimization of the sum of squares of the weighted residuals of the measured coordinates α, δ array (α, δ)nand their estimated values αp, δpcalculated using the parameters of the vector R(tnand working a stochastic model of the SPACECRAFT motion, accepted for processing. Using the current vector of parameters of the orbit R(tn), form a new initial conditions X(tn) (26). These new parameter values of the initial conditions X(tn) used for subsequent navigation and ballistic calculations before determining the current vector of parameters of the orbit R(tnby results of the information processing apparatus monitoring system on-site the next round of flight KA Thus, update the initial conditions X(tn) perform at each point of the flight SPACECRAFT, which carried out the operation of the apparatus sensing.

If the accuracy of the current vector of parameters of the orbit R(tn)obtained by the information processing apparatus monitoring system in the working portion of one turn of the flight SPACECRAFT, is inadequate to meet the requirements of monitoring, perform joint processing of data generated by multiple turns of the flight SPACECRAFT. In this case, the arrays of measured coordinates

KA (α, δ)n, (α, δ)n-1,...,(α, δ)n-q+1received on the sliding interval q of coils flight KA form (27) in the aggregate arrays measured SPACECRAFT coordinates {(α, δ)n, (α, δ)n-1,...,(α, δ)n-q+1} (28). Coordinates α and δ from the set of arrays {(α, δ)n, (α, δ)n-1,...,(α, δ)n-q+1} handle together (29), using OLS, and define the vector of the adjusted values of the current parameters of the orbit R(t0) (30). When processing MNCs provide minimization of the sum of squared weighted residuals (differences) of the measured coordinates α, δ from the set of arrays {(α, δ)n, (α, δ)n-1,...,(α, δ)n-q+1} and their estimated values αp, δpcalculated using the parameters of the vector R(t0and working in a stochastic model the motion of the SPACECRAFT. Forms the t new initial conditions X(t n) (26)using the updated values of the parameters of the orbit R(t0). These new parameter values of the initial conditions X(tn) used for calculations of navigation and ballistic security until the following definition of the vector of current values of parameters of the orbit R(t0), corrected according to the results of information processing work area equipment monitoring system the next round of the flight SPACECRAFT. Thus, the update of the initial conditions X(tnin the case of determining adjusted current values of the orbital parameters from measurements of the SPACECRAFT coordinates at several turns his flight performed at each point of the flight SPACECRAFT, which carried out the operation of onboard equipment monitoring system.

If the definition of the adjusted values of the current parameters of the orbit as measured from the several coils of the flight SPACECRAFT could not be performed due to the scarcity of computational resources computational tools for joint processing can be used compressed information, which are the current orbital parameters determined from measurements of one turn of the flight SPACECRAFT. For this form (31) the set of vectors of the current orbital parameters {R(tn), R(tn-1),...,R(tn-q+1)}, obtained from measurements of individual work stations on the sliding interval q of coils flight KA (32) the Parameters of the orbits of R(t n), R(tn-1),...,R(tn-q+1from the set of vectors {R(tn), R(tn-1),...,R(tn-q+1)} filter (33), using OLS, and define the vector of exact values of the parameters of the orbit R(tf) (34). When filtering (33) allow to minimize the sum of squared weighted residuals values of the processed parameters of the orbits of R(tn), R(tn-1),...,

R(tn-q+1) from the set {R(tn), R(tn-1),...,R(tn-q+1)} and the estimated values of Rp(tn), Rp(tn-1),...,Rp(tn-q+1), calculated using the parameters of the vector R(tfand working in a stochastic model the motion of the SPACECRAFT.

Form a new initial conditions X(tn) (26)using the vector of the exact values of the parameters of the orbit R(tf) (34). These new parameter values of the initial conditions X(tn) used for subsequent navigation and ballistic calculations until the following definition of the vector of exact values of the orbital parameters that are performed at each point of the flight SPACECRAFT, which carried out the operation of onboard equipment monitoring system.

In proceedings of the OLS treatment (24), (29) and (33) as a working model for the motion of the SPACECRAFT for the calculation of the weighted residuals is applied stochastic model for the motion of the SPACECRAFT. To determine the estimated values of αp, δp, the SPACECRAFT coordinates in the treatments (24), (29), as well as ascetic values of R p(tn), Rp(tn-1),...,Rp(tn-q+1parameters of the orbit in the processing of (33) perform calculations deterministic components of the working model of the motion of the SPACECRAFT, which sets the correspondence of the parameters in the source and destination points prediction by numerical integration of the equations of motion of the SPACECRAFT. Using the results of these calculations form the residuals of the processed parameters (coordinates) and their corresponding calculated values. When determining the weight matrix of residuals sum of the covariance matrix of the errors of the SPACECRAFT coordinates α, δ, processed in the calculations (24) and (29), or errors of the parameters of the orbit R(tn), R(tn-1),...,R(tn-q+1)processed in the calculations (34), and the covariance matrix of methodological errors working model of the motion of the SPACECRAFT, which is determined by the stochastic component of the working motion models. The stochastic component of the working motion models is the probabilistic characteristics of its methodological errors due to unaccounted for in the numerical integration of the perturbation of the orbit and its computational errors. Such a truncation error of a working model can occur, for example, in the case of reducing the number of harmonics of the Ground potential, and increase step integration of the equations of motion of the SPACECRAFT that primeniaut reduce consumption of computing resources. The numerical values of the probability characteristics of methodological errors obtained by computational experiments. When computing experiment on the interval τ (not less than d interval treatments(24), (29), (33)), perform parallel calculations using the software package for the numerical integration precision reference model the motion of the SPACECRAFT and software package working model that is used in the treatment of(24), (29), (33). The residual results of calculations of the reference and working models identified on the interval τ, minimize, using the OLS procedure for a corresponding change in the initial conditions of the calculation of the working model. The quadratic form of the residual residuals of these calculations, taking into account the degrees of freedom to solve the system of normal equations OLS is [5] the covariance matrix of methodological errors working model.

Modelling studies performed during the development of the proposed invention, confirm the effectiveness of the proposed method orientation and Autonomous navigation of SPACECRAFT monitoring system of the Earth and near-earth space by using information generated by onboard equipment sensing of the earth's surface for its intended purpose, in the interests of precision orientation and Autonomous navigation of SPACECRAFT operating in an offline match the optical mode at geo and HEO orbits [1]. Shown the possibility of obtaining highly accurate data about the current orientation of the SPACECRAFT and its current coordinates at the working sites geo or HEO orbit. High accuracy and reliability) is achieved by using a very large volume of redundant information, which represent the coordinates of the points of the boundaries between space and Earth, i.e. the contour of the observed shape of the Earth. Applies a set of stages of joint optimal aggregation of this information.

Used sources of information

1. Vlasko-Vlasov KA. From the "Comet" to "Eye". M., "Olga", 2002.

2. Bazhinov I.K., Hawks EAST Navigation in joint flight spacecraft Soyuz and Apollo. M.: Nauka, 1978.

3. Vasilevsky A.S., Zheleznov M.M., Simon AL, Polanski IV Operational gridded video data of Earth remote sensing SPACECRAFT series "meteor-M". M, IKI. Seminar, Russia, Tarusa, 2006. Internet: http : //www.iki.rssi.ru/seminar/tarusa 200606/tarusa_06 pd (prototype) f.

4. Fundamentals of theory of flight spacecraft. Edited Gesneriana and Mccicmservice. M., engineering, 1972.

5. Hudson D. Statistics for physicists. M., Mir, 1970.

1. The way orientation and Autonomous navigation of a SPACECRAFT (SC) monitoring system of the Earth in which to determine the orientation of the SPACECRAFT IP is result measured onboard Astroparticle angles between the construction axes AC and directions assigned to the navigation stars with known coordinates of right ascension α iand declination δiand use the current coordinates of the SPACECRAFT on the celestial sphere αpand δpcalculated using the initial conditions X(tn)intended for navigation and ballistic calculations, characterized in that to determine the current orbital parameters and Refine the calculated data of the current coordinates and orientation of the KA process information generated by onboard equipment sensing in the process of solving problems for its main purpose, and when processing of each individual frame sensing in j-x lines of his sweep allocate the elements of the resolution with coordinates (j, ij(n)that correspond to the boundaries of the "space - to-Earth" (j, ij(1)) and "Earth - to-space" (j, ij(2)), and the coordinates of the elements of the resolution selected in the set of lines in the frame and representing the coordinates of the contour points and the observed shape of the Earth, remember, accumulate, process, least-squares and determine the coordinates of the j0, i0the center of the Earth, observed in the field frame, which are compared to the coordinates of the center of the Earth j(0)i(0)corresponding to the calculated current SPACECRAFT coordinates αp, δpcalculate the difference Δj, Δi specified coordinates and these differences are adjusting estimated the values of the orientation angles CA, used to determine the coordinates of objects sensing, and determine the values of the corrections to the calculated coordinates KA αp, δpreceiving the measured current coordinates KA αto, δtoat time tkbinding information of the k-th frame sensing.

2. The method according to claim 1, characterized in that for selecting the resolution of the j-th row scan frame sensing, corresponding to the boundaries of the space - to-Earth and Earth - to-space", is formed in the sliding mode along the line sweep pair of adjacent samples of limited length {gji}1and {gji}2containing data on the radiation parameters of the gjispace and Earth adopted in the separate elements of the resolution, estimate the mathematical expectation M1[g], M2[g] and the variance D1[g], D2[g] the radiation received in the adjacent samples of the set of pairs of adjacent samples that are generated in the j-th row, highlight those differences estimates of M1[g] compared to M2[g] and D1[g] in comparison with the D2[g] take extreme (max, min) values and coordinate values of j and i elements permits separating adjacent sample selected pairs, are used to determine the coordinates of the boundaries of the "space-to-Earth" (j, ij(1)) and "Earth-to-space" (j, ij(2)), taken as the coordinates of the point the contour of the observed shape of the Earth.

3. The method according to claim 1, characterized in that the measured coordinates KA αto, δtoobtained in the frame-like sensing in the working portion of the n-th round of flight KA, remember, accumulate and form the array of coordinates {α, δ}nwhile the coordinates α, δ array {α, δ}nprocess together, using the method of least squares, and define the vector of the current parameters of the orbit R(tnat the time tnthe n-th round of flight KA, minimizing the sum of squared weighted residuals processed coordinates α, δ array {α, δ}nand their estimated values αp, δpcalculated using the parameters of the vector R(tnand working in a stochastic model the motion of the SPACECRAFT, accepted for processing, and when determining the weight matrix of these residuals sum of the covariance matrix of errors processed SPACECRAFT coordinates α, δ and the corresponding methodological errors working model of the motion of the SPACECRAFT, and the parameters of the vector R(tn) are used to form new initial conditions X(tnused in subsequent navigation and ballistic calculations.

4. The method according to claim 3, wherein the set of coordinates αto, δtoobtained in the frame-like sensing in the working portion of the n-th revolution of the flight SPACECRAFT is divided into subsets corresponding to the Onam smoothing and the resulting optimal smoothing coordinates αm, δmform array (α, δ)ninstead of the coordinates αto, δto.

5. The method according to claim 3, characterized in that the arrays of coordinates {α, δ}nreceived on the sliding interval q of coils flight KA, remember, accumulate and form together arrays of coordinates {(α, δ)n, (α, δ)n-1,...,(α, δ)n-q+1}, with coordinates α, δ from the set of arrays {(α, δ)n, (α, δ)n-1,...,(α, δ)n-q+1} handle together, using the method of least squares, and define the vector of the adjusted values of the parameters of the orbit R(t0), while providing the minimization of the sum of squares of the weighted residuals of the processed coordinates α, δ from the set of arrays {(α, δ)n, (α, δ)n-1,...,(α, δ)n-q+1} and their estimated values αp, δpcalculated using the parameters of the vector R(t0and working in a stochastic model the motion of the SPACECRAFT, when determining the weight matrix of residuals sum of the covariance matrix of errors processed SPACECRAFT coordinates α, δ and the corresponding methodological errors working model of the SPACECRAFT motion, the vector R(t0) is used to form the new initial conditions X(tnused in subsequent navigation and ballistic calculations.

6. The method according to claim 3, otlichayushiesya, the vectors of the current parameters of the orbit R(tnreceived by the sliding interval q of coils flight KA, form the set of vectors {R(tn), R(tn-1),...,R(tn-q+1)}, the parameters of the orbit R(tn), R(tn-1),...,R(tn-q+1from the set of vectors {R(tn), R(tn-1),...,R(tnq+1)} filter using the least-squares method, and define the vector of exact values of the parameters of the orbit R(tf), providing the minimum sum of the squares of the weighted residuals of the processed parameters of the orbit R(tn), R(tn-1),...,R(tn-q+1from the set of vectors {R(tn), R(tn-1),...,R(tn-q+1)} and the estimated values of Rp(tn), Rp(tn-1),...,Rp(tn-q+1), calculated using the parameter values of the vector R(tfand working in a stochastic model the motion of the SPACECRAFT, and when determining the weight matrix of residuals sum of the covariance matrix of the errors of the processed parameters of the orbit R(tn), R(tn-i),...,R(tn-q+1) and the corresponding methodological errors working model of the motion of the SPACECRAFT, and the calculated parameters of the orbit R(tf) is used to form the new initial conditions X(tnused in subsequent navigation and ballistic calculations.



 

Same patents:

Astro-finder device // 2319109

FIELD: astronautics, in particular, systems for astro-correction of launch azimuth of carrier rockets.

SUBSTANCE: claimed device is used for launch azimuth correction of position of gyro-stabilized platform of carrier rocket control system and contains optical objective and photo-sensitive section positioned on aforementioned platform. The platform consists of at least two photo-detectors with different spectral characteristics and a scanning element, made in form of piezo-element with rounded reflecting end. When scanning element oscillation is low, any beam after the objective is reflected from rounded surface and received on a single photo-detector. That moment depends on position of a star relatively to finder axis of objective and on sweep path. When sweep is stable, the moment of appearance of signal from photo-detector inside the sweep period is rigidly connected to position of aforementioned beam, and position of finder axis is fixed when device is positioned on gyro-stabilized platform. When two photo-detectors are used with different spectral characteristics, it becomes possible to correct position of gyro-stabilized platform on basis of two spectrally different navigational stars. It is possible to install the device directly on the gyro-stabilized platform of control system of carrier rocket and to precisely aim the rocket at active flight section after comparatively rough aiming based on launch azimuth during pre-launch preparation, which results in simplified and cheapened ground-based aiming system, and allows to remove errors in launch azimuth caused by time difference between launch of engine block of first step of rocket and its release from launching device.

EFFECT: reduced mass and size of astro-finder device, increased reliability of its operation.

2 dwg

FIELD: space engineering; forming satellite systems for positioning objects on earth surface.

SUBSTANCE: proposed method includes injection of N artificial satellites into circular or other orbits which work in "n" planes (where "n" is integer which is more than 2) by mi (i=1, ...n) satellites (where n is integer) in each plane. Satellites are positioned in orbits of datum plane and in planes located symmetrically and in pairs relative to datum plane. These planes of orbits are positioned irregularly along terrestrial equator relative to datum plane at angles a priori not equal to 360°/n. Artificial earth satellites in orbits are positioned irregularly and symmetrically in pairs relative to base satellite.

EFFECT: reduced number of artificial earth satellites in navigational system with no impairment of system parameters at positioning of ground objects.

1 dwg, 1 tbl

FIELD: information satellite systems; forming global radio-navigational field for sea-going ships, ground, air and space vehicles.

SUBSTANCE: proposed system includes many low-orbit spacecraft whose number depends on conditions of global covering of access areas of users. Each spacecraft contains communication unit in addition to navigational equipment for communication of this spacecraft with two other spacecraft in its orbital plane and two spacecraft from adjacent orbital planes. Communication is performed in millimetric wave band absorbed by Earth atmosphere. At least one spacecraft is provided with high-accuracy synch generator. Thus spacecraft group is formed which is provided with noise immunity system of relaying and measuring radio lines connecting all spacecraft groups and navigational radio line covering upper hemisphere.

EFFECT: enhanced reliability and accuracy; enhanced noise immunity of data fed to users of satellite system.

1 dwg

The invention relates to space technology and, in particular, to methods and means for securing the binding time of registration of the observed phenomena on Board the spacecraft (SC) to local time on the Earth

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Navigation system // 2168703
The invention relates to aircraft instrumentation and can be used in the airborne equipment aircraft, providing management and guidance

The invention relates to aircraft instrumentation and can be used as part of the on-Board aircraft equipment, ensuring the fulfilment of the tasks of navigation and targeting

The invention relates to aircraft instrumentation and can be used in the airborne equipment aircraft for navigation, control and guidance

FIELD: aerospace engineering.

SUBSTANCE: proposed platform comprises service hardware module representing a rectangular parallelepiped formed by end face plate (1) and four lateral boards (2, 3, 4, 5). Two intermediate boards (6, 7) are fitted there inside to divide said module into three compartments for service hardware. Storage batteries are arranged between boards (4, 7), while between boards (5, 6) onboard control complex, power supply, orientation and stabilisation systems electronic devices are arranged. Lateral board (5) accommodates control devices of orientation and stabilisation systems (11) and antenna (12). One of the service module boards, parallel to service boards, accommodates assemblies to couple with separation system. Engine plant is mounted between intermediate boards (6,7) in the area of assumed space ship center of gravity so that thrust vector of micro engine (19) is parallel to intermediate boards and perpendicular to end face board (1). Engine plant can displace in two mutually perpendicular directions thanks to appropriate adjusting elements (threaded posts with nuts, slots and "screw-nut"-type mechanisms). Solar battery panels (9, 10) are arranged on brackets extending beyond module edges. Useful load module mounting assemblies are arranged on free end faces of module lateral boards and extending brackets. Note here that useful load module hardware are arranged in space between solar batteries (9,10) and free zone of the module on the side of its exposed part. To protect instrumentation and engine plant, detachable covers (33, 34, 35) are provided. Engine plant incorporates fuelling coupling (36).

EFFECT: improved compactness of space vehicle, higher performances.

7 dwg

FIELD: space-system engineering.

SUBSTANCE: inventions relate to means and methods of conducting predominatly long-term experiments on low-earth orbit in field of materials production under conditions of ultra-high vacuum. Spacecraft (SC) consists of pressurised section and lock chamber with access door and escape door, mechanical arm, molecular beam epitaxy (MBE) set and molecular protective screen. Lock chamber escape door hatch is made detachable, protective screen is firmly fixed to this hatch on its outside and is equipped with attachment point to the mechanical arm end link. MBE set is connected to the mentioned hatch on its inside. SC is equipped with device for attaching escape door hatch to lock chamber. According to the present invention, method also provides for boosting MBE set together with screen out of the lock chamber, detaching them from SC with mechanical arm, orienting screen to position of protecting MBE set from outside atmosphere flow. Hereafter researches in screen wake sphere are conducted and MBE set is brought back to lock chamber. At this stage MBE set and screen are fixed on the mentioned hatch. When bringing back MBE set to lock chamber using mentioned attachment device, angular and linear errors of mechanical arm overlapping escape door hatch and lock chamber axes are corrected. Mentioned hatch is engaged from mechanical arm and is axially moved till it is completely attached to the chamber with all the necessary docking interface sealing forcing provided. Before starting experiments in screen wake, screen can be oriented in a way to secure maximal illumination of MBE set equipment and screen for its degasification.

EFFECT: increasing data rate and quality of conducted MBE researches and reducing expenses on their realisation.

8 cl, 8 dwg

FIELD: transportation.

SUBSTANCE: invention is related to satellite systems for realisation of communication and monitoring tasks, comprising groups of spacecrafts (SC) brought to orbits of different height. System is intended for servicing of vast geographic region: land territory with adjacent sea and ocean areas. System comprises SC on highly elliptical orbits (101-105), including two SC (1, 2) for meteorological and heliogeophysical monitoring and specialised SC (3, 4, 5) for communication, and also SC (6, 7) at low orbits (106, 107) for radio-locating monitoring. System comprises ground complex (29) for SC control, ground complex (24) for reception, processing and distribution of space data and ground communication segment (25). At that possibility is provided to transfer information from SC (1, 2) directly to ground complexes (24, 29), and also to transfer data by components of ground complex (24) and ground communication segment (25) with application of space retranslation channels via SC (1-5).

EFFECT: improved reliability and expansion of system functions in combination with simplification of its service.

9 cl, 4 dwg

FIELD: aircraft industry.

SUBSTANCE: invention refers namely to telecommunication satellites with energy consumption power of 1-2.5 kV. According to the invention, space vehicle (satellite) is made of two modules: payload and support systems. Instruments are installed on inner skins of their radiators - honeycomb panels. Evaporation zones of horizontal straight and uncontrolled L-shaped heat tubes are built into those panels opposite the instrument location area. Those zones are attached with their flanges to inner skin of panels in that area. Condensation zones of the above tubes are arranged in the panel areas free of instruments. Radiators are located in planes perpendicular to the axes corresponding to northern and southern sides of the vehicle. At that, opposite located radiator panels are arranged at a minimum possible distance from each other, which is determined beforehand based on instrument arrangement conditions. More heat-stressed instruments are located in the lower part of panels, in which there used are uncontrolled L-shaped heat tubes. Condensation zone flanges of the above heat tubes are made so that they face outer skins, and are attached thereto. The above zones are located in extreme zones of panels, which are free from instruments.

EFFECT: decreasing mass, and reaching the acceptable configuration of the above satellites.

4 dwg

FIELD: space engineering.

SUBSTANCE: proposed invention relates to space engineering and can be used in development of space vehicles intended for comprehensive investigation of soil of celestial bodies and delivery of effective loads to the massifs of Mars, Moon, asteroids and other planets and celestial bodies of solar system. The device to deliver effective cargo into celestial body massifs (versions) comprises a hollow primary structure having a front and cylindrical tail part accommodating a ballast with a mean density exceeding that of the primary structure and effective cargo. The length of cylindrical tail parts makes 8 to 15 its diameters, the center of masses being located at the distance equal to 0.4 to 0.5 of the primary structure length starting from the front part head. In compliance with the first version the front part, starting from the head, represents the first truncated cone or, simply, cone of the cylinder behind the head, the said truncated cone lager base abutting on the cylinder tail part of the other truncated cone. In compliance with the second version the front part hast the holes communicating with the primary structure space wherein ballast and effective cargo are located, while the ballast or a part of it are made from materials can penetrate, due to inertial forces, through the said holes into ambient medium.

EFFECT: deeper penetration, smaller area in contact with medium, rectilinearity and predictability of trajectory of motion in medium, reduced factor of friction with medium in contact.

5 cl, 3 dwg, 1 tbl

FIELD: geophysics.

SUBSTANCE: invention is related to procedure of global geophysical events monitoring and prediction of emergence and development of natural and anthropogenic disasters on Earth. System comprises space segment and surface segments. Space segment consists of three orbit groups. In orbit group of small spacecrafts (SSC), which are located on geostationary orbit, SSC that are combined into two orbit groups of three satellites on tops of two triangular planes, create constellation of six tops. For orbit group that consists of 3-4 SSC on sun-synchronous orbits with height of 600-700 km, orbit planes are evenly distributed along longitude of ascending node. In orbit group of 50 microsatellites (MSC - micro-spacecrafts), the latter are located mainly on sun-synchronous orbits and partially on geostationary orbits. Highly sensitive equipment is installed on SSC and MSC with complex of instruments for measurement of foreshocks and sensors of operational control and prediction of natural and anthropogenic disasters.

EFFECT: system that provides automated aerospace monitoring of global geophysical natural and anthropogenic disasters, makes it possible to obtain operative short-term forecast - warning hours and days before.

10 cl, 9 dwg

FIELD: space engineering.

SUBSTANCE: proposed method consists in that, on receiving the spacecraft in distress alarm signal, the Doppler frequency of the signal above is measured by an interrogatorless method to locate the spacecraft at the moment when the said frequency equals zero. At this very moment, the angle between the spacecraft receiving antenna axis and that of the horizon pickup is measured to calculate the coordinates of the point below the satellite. The aforesaid measurements are made two times and the coordinates of the points below the satellites along the measured angles allow determining the coordinates of the spacecraft in distress on the Earth surface. To measure the Doppler frequency, two signal processing channels are used wherein the received signal frequency is converted using the onboard master oscillator. In the first processing channel, the oscillator voltage is phase-shifted by 90°, the frequency difference voltages are isolated, amplified and amplitude limited to be converted into rectangular clipped voltages. The first channel voltages of this type are converted into a series of short positive pulses with their time position corresponding to the moments of the voltage passing through the zero level with a positive derivative. Adjacent clipped positive voltages of the second channel are inverted in phase by 180° and quantised by the said short positive pulses of the first channel. The quantised pulses digital form allows determining the Doppler frequency. Given its zero value, corresponding to the spacecraft passing through the beam point, the control pulse is generated to allow the further processing of the received signal.

EFFECT: higher accuracy of measuring minor values of the Doppler frequency and fixation of its zero value.

3 dwg

FIELD: space engineering.

SUBSTANCE: proposed method consists in using a vessel aboard the satellite with a required constant area of the liquid phase location on the said vessel wall surface. The Earth satellite lengthwise axis, in operating conditions, is constantly directed along the current radius-vector of the orbit mass centre. The aforesaid vessel is placed at a maximum possible distance from the satellite mass centre along the direction parallel to the said lengthwise axis. The vessel should be placed so that the normal at the point of the vessel wall surface opposite the working medium outlet was directed towards the satellite mass centre. The said point should be located at a minimum possible distance from the satellite lengthwise axis.

EFFECT: simpler design and smaller weight of service satellite systems.

2 dwg

FIELD: space-system engineering.

SUBSTANCE: system can be used on the earth satellite vehicle in orbit except geostationary centers which are stabilised by rotation along the vertical axis, and the ground reception centers. The system contains an emergency object transmitter, spacecraft equipment and ground reception center equipment. Spacecraft equipment includes a horizon sensor, receiving antenna, comparison device, receiver, Doppler frequency meter, blocking-generator, two gating circuits, two valves, pulse generator, pulse counter, plug board chart, and magnetic storage device, transmitter, transferring antenna, modulation code generator, high-frequency generator and power amplifier. The ground reception center equipment includes reception antenna, high-frequency amplifier, two mixers, calibration frequency block, phase doublet, three narrowband filter, scale-of-two phase circuit, phase detector, Doppler frequency meter, computing unit and registration block.

EFFECT: it enlarges functional possibilities of the system.

5 dwg

FIELD: production methods.

SUBSTANCE: suggested method includes emission and retranslation of primary and final radio signals between spaceship, basic and additional land determinations stations. At the same it is additionally retranslated the final radio signal from spaceship to the basic land station and it is admitted at this station. The radio connection with the radio signal of the spaceship with one or more additional stations, admission of the primary signal to the additional station, its transformation to the final signal and admitting of it to the spaceship. About the distance between spaceship and main determination station is judged by the interval between the moment of emission and moment of admitting the primary signal at this station. About the distance between spaceship and additional determination station is judged by the interval between the moment of emission and moment of admitting the primary signal at the main station. It is measured additionally the moving of the frequency of the final signal, admitted on the main determination station, regarding the frequency of the primary movement, emissed from the same station. The distance between spaceship and additional station is determined with measuring of the Doppler drift.

EFFECT: it is reduced the time and increased the accuracy of distance determination between spaceship and determination stations.

5 cl, 4 dwg

FIELD: rocketry and space engineering; scientific and commercial fields.

SUBSTANCE: proposed method includes placing payloads on injection facility, launching the launch vehicle, separation of injection facility from launch vehicle and injection of injection facility into geocentric orbit where said payloads are separated from injection facility. Main payload is placed on injection facility directly of body of accompanying payload; this body combines its functions with functions of main load-bearing member of adapter system for placing the main payload. After separation of injection facility from launch vehicle, additional acceleration of injection facility is performed and injection facility is injected into reference orbit and then it is shifted to geocentric orbit where main and accompanying payloads are separated. Accompanying payload is separated from injection facility after main payload is at safe distance without waiting for complete turn of main payload. Spacecraft in facility injecting the artificial satellites into geocentric orbit are placed in succession on injection facility beginning with lower one. Main payload in form of one or several spacecraft is placed on body of lower spacecraft through separation device. Body of lower spacecraft combines its functions with functions of adapter load-bearing member for placing the main payload.

EFFECT: increased mass ratio of launch vehicle and injection facility; extended functional capabilities.

3 cl, 2 dwg

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