RU2659379C1 - Method of the geostationary region survey for the space debris elements and other objects observation from space craft on the semi-diurnal high-elliptic orbit - Google Patents

Abstract

FIELD: astronautics.

SUBSTANCE: invention relates to the space debris and other objects near the geostationary orbit (GSO) detailed images obtaining methods. Survey is performed from the spacecraft (SC) on the semidiurnal high-elliptical orbit (HEO) with apogee A for 200 km below or 500 km above the GSO, and perigee up to 5,000 km, with inclination from 0 to 5°. HEO parameters are selected from the condition of the objects images obtaining for the minimum number of days – in the GSO F₂F₃ controlled areas from the HEO working section, symmetrical to the apogee A. Overlapping sections F₂F₃ cover the entire GSO for the minimum number of the SC turns. According to the transmitted from ground stations tasks, the movement parameters (including the exact angular positions) of objects in the controlled area can be measured from the HEO specially selected sections. SC transmits images during the ground station direct visibility periods, where the observed objects motion parameters are computed.

EFFECT: technical result consists in the achievement of the seamless GSO survey for a minimum time with the given limitation on the objects observation range from the SC.

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Description

The aim of the present invention is to develop a method for obtaining detailed images of space debris and other objects located in a geostationary orbit, as well as measuring the angular coordinates of objects in a geostationary region. The observation range and the size of the observed objects are determined by the characteristics of the spacecraft observation equipment.

Space systems are known that implement methods for viewing the geostationary region to obtain coordinate and non-coordinate information about objects in this region.

1. The spacecraft (SC) of the GSSAP system are located in quasi-geostationary orbits with an inclination of about 0.48 ° and a period of revolution slightly different from stellar days, which ensures constant drift at speeds of about 0.4 ... 0.6 ° / day. relative to the spacecraft in geostationary orbit (GSO). Such a construction provides the possibility of obtaining detailed non-coordinate information on geostationary objects on the span during the drift and obtaining coordinate and photometric information almost constantly (with the exception of restrictions due to lighting conditions) [1, 2].

The disadvantages of this review method are the relatively large times required to access a particular object on the GSO, measured at intervals of several months.

2. The SBSS system began to be deployed in 2010 and consists of spacecraft orbiting in low circular sun-synchronous orbits with an inclination of 97.97 °, a height of perigee 625.0 km, a height of apogee 640.0 km, a period of revolution of 97.42 min Its tasks include an overview of outer space, measuring the coordinates of observed objects and providing information to spacecraft controls. [3].

The disadvantages of this review method include the impossibility of obtaining detailed non-coordinate information on objects in the geostationary region due to the large distances to these objects.

As a prototype, a method for viewing the geostationary orbit by the space system described in [4] was adopted.

To observe objects on the GSO, the authors propose to place spacecraft in semi-daily high elliptical orbits (VEO) with an inclination of about 63.4 ° and an argument of the latitude of perigee in the region of 0 ° or 180 °. Additionally, the requirement was set for the daily repeatability of the spacecraft route, which is ensured (which the authors do not indicate) with a circulation period of 43065 s ... 43067 s. It is indicated that the requirements for observing GSO lead to the decomposition of this task into two components:

- task 1 - survey observation of GSO;

- task 2 - a detailed observation of the detected objects on the GSO.

The main difference between the orbit parameters for solving Problem 2 and the orbit parameters for solving Problem 1 is the increase in perigee height necessary for the apogee to be located in a close vicinity of the GSO.

To solve problem 1, it is proposed to review the entire GSO from two spacecraft of the KA-O type with an observation range of no more than 17,000 km. Both KA-O are located in the same plane, but are separated on the VEO by the time of its passage. They pass the working section of the orbit with a time shift of 6 hours. Throughout the working section, each KA-O is available for observation at the required range of range, the GSO arc section is of the order of 90 °. In this case (which the authors do not indicate) it is necessary that the apogee apnea exceeds the GSO by ~ 4000 km. Thus, in two turns, i.e. during the day, a full review of the GSO can be provided with an observation range of no more than 17,000 ... 18,000 km.

To solve problem 2, that is, detailed control of objects at the GSO, a different type of spacecraft KA-DK is used with a range of observation range from 100 km to 400 km, located on a wind turbine, which has a climax in the close vicinity of the GSO. By the condition of the daily repeatability of the route, the apogee of the same (even or odd) turns has an unchanged geographical longitude with a distance of 180 ° between the even and odd turns. As the corresponding calculation shows, the length of the monitored section of GSO longitudes with a maximum observation range of 400 km at each turn is approximately ± 440 km relative to the apogee longitude. To maintain space objects throughout the entire GSO, it is possible to provide a certain daily discrete displacement (drift) of the KA-DK working section along the GSO in longitude. Such a drift is carried out by appropriate correction of the orbit's orbital period.

The specified prototype has the following significant disadvantages, predetermined mainly by the described way of reviewing the GSO.

1. A method for survey observation of GSO and a method for detailed observation of detected objects at GSO is implemented by two different types of spacecraft — KA-O and KA-DK, and the KA-O and KA-DK orbits differ in altitude of apogee and perigee.

2. The task of fully covering the entire GSO with detailed control areas for a limited number of days is not even set.

3. The long term for the full implementation of the GSO by the detailed control sections in the regime of a fixed drift speed, which ensures the adjoining of the control sections on the adjacent identical (even or odd) turns. According to the corresponding calculations, this requires at least 100 days with a maximum range of detailed observation of 400 km.

The technical result of the invention consists in a new way to view the geostationary region from one spacecraft in a semi-daily highly elliptical orbit. By semidiurnal VEO in the following we mean VEO with a period T less than half a stellar day (43082 s) by up to 800 ... 1000 s. The proposed method provides both measuring the exact coordinates of the object in a given region of near-Earth space, and obtaining detailed images of the object at any point of its standing on the GSO with the shortest possible time for a full review of the GSO. Upon receipt of detailed images, the maximum observation range L _max , determined by the characteristics of the observation equipment, is, for example, 400 km, and the permissible minimum observation distance is also determined by the characteristics of the observation equipment.

The technical result of the invention also lies in the rational choice of the parameters of a semi-daily highly elliptical orbit - the period of revolution, the radius of the orbit at the peak, the argument of the latitude of perigee, inclination. With these parameters, at the working site of the orbit, the observation range of the object on the GSO is ensured, not exceeding the permissible value of 400 km, at which the required resolution is observed. The coordinates of the objects in the geostationary region are measured at selected areas of the HEO at a range whose maximum value is determined by the penetrating power of the BAO.

The invention aims to minimize the time of a full non-throughput survey of the entire geostationary orbit when obtaining detailed images of objects in this orbit.

The invention includes a method for selecting parameters of a semidiurnal highly elliptical orbit. For such an orbit, it is proposed to use the inclination i, which is in the range from 0 ° to 5 °. The radius of the orbit at the apogee should exceed the radius of the GSO by an amount slightly less than the maximum permissible observation range L _max . From the apogee section of this orbit, the spacecraft provides a detailed image of each object with a standing point at any GSO longitude for a minimum number of days. Moreover, the range of detailed observation does not exceed a given value L _max .

In FIG. Figure 1 shows the configuration of GSO and VEO in the equatorial plane and the formation of a controlled area on the GSO. The drawing shows the apogee A of a highly elliptical orbit and the VEO working section between points B and C. The controlled section on the GSO is limited to points F ₂ and F ₃ , which are removed from the extreme points of the working section B and C by L _max . In this diagram, the motion of the spacecraft in a highly elliptical orbit and of objects in the GSO, which changes the size of the observed portion of the GSO, is not taken into account, which is explained in detail and shown below.

The proposed method achieves a rational choice of the parameters of a semi-daily highly elliptical orbit depending on the characteristics of the observation equipment used. With these parameters, the length of the GSO longitude interval is estimated for one revolution, on which, when the spacecraft passes the apogee section of the HEO, the range of obtaining a detailed image is ensured, not exceeding the maximum permissible. The VEO parameters are chosen so that the indicated longitude intervals for adjacent identical (even or odd) turns have some overlap and do not leave uncontrolled intervals on the GSO. At the same time, the number of days for which the GSO is completely covered by observation intervals is the minimum possible.

The coordinates of objects in the geostationary region are measured according to tasks transmitted from ground control points. Detailed or coordinate information received at intervals of the line of sight of the spacecraft from ground control points and receiving information is transmitted to these points during the observation process. If detailed or coordinate information is received outside the line of sight intervals, this information is stored and transmitted from the spacecraft to ground-based information receiving and processing points during periods of its direct line of sight.

According to the proposed method for viewing the geostationary region, it is possible to obtain a detailed image of an object moving in a geostationary region in an orbit with a non-zero inclination, that is, not located on the GSO, but periodically approaching it. To do this, make accurate measurements of the observed coordinates of the object and on the ground command and control point calculate the parameters of its orbit. They make a prediction of the object’s movement for a certain number of turns, determine a number of positions when the object passes near the GSO, and record the exact values of its coordinates in three-dimensional linear space. From these positions, the optimal one is chosen and the time instant and the vector of application of such a satellite orbit correction pulse are calculated, after which the spacecraft will be at the selected point at the appointed time.

The configuration of the spatial position of the geostationary and semidiurnal highly elliptical orbits is shown in FIG. 2. The drawing shows the center of the Earth O, the polar axis ON, the perigee P, the radius vector of the spacecraft at the apogee OA, the current position of the spacecraft at point K and the radius of the GSO r _gco (segment OA _GSO ). A highly elliptical orbit with a nonzero positive small inclination i and an argument of the latitude of the perigee ω _π = 0 is considered (however, the mirror-symmetric picture with an inclination of 180 ° - i is shown in the drawing for clarity).

It is taken into account that in order to achieve the maximum length of the longitude interval on the GSO, on which detailed images are obtained, the argument of the latitude of the perigee ω _π should be equal to 0, and the radius of the apogee of the HEO (point A) must exceed the radius of the GSO. When choosing the magnitude of the apogee radius r _ar take into account that the height of the apogee above the GSO (segment AA _GSO ) should be slightly less than the maximum range of detailed observation L _max .

When determining the length of the GSO longitude interval, at which it is possible to obtain detailed images of objects at a range not exceeding L _max , a value of λ _wu is assigned at which the apogee of this turn has a geographical longitude D = 0.

To specify an elliptical orbit, use the following parameters:

- period of circulation T;

- the radius of apogee r _ar ;

- geographical longitude of the initial ascending node λ _wu ;

- the argument of the latitude of perigee ω _π ;

- the inclination i.

Based on these parameters, for any current point in time, the coordinates of the spacecraft — the radius vector r, geographical longitude D and latitude W — are calculated. Next, the geometric parameters shown in FIG. 2: the distance h from the spacecraft to the plane of the equator, the minimum distance D _q from the spacecraft to the GSO, the projection of the radius vector of the spacecraft on the equator plane (segment OK _e ) and the distance b in the plane of the equator from point K _e to point S located on the GSO on maximum range of observation L _max .

On certain three sides of the oblique triangle OK _e S in the plane of the equator, the angle Δ Д of this triangle is calculated. As mentioned above, when rationally setting the value of λ _{wu, the} apogee of this turn has a longitude equal to zero. In this case, the angular distance in longitude D _Σ from the apogee to the point S on the GSO, achievable for observation but maximally distant from the apogee at a given time, is determined by the expression D _Σ = D + ΔD.

According to the results of calculations with a small time step in the interval of spacecraft motion on the VEO from perigee to apogee, the maximum value of D _{Σmax of the} value of D _Σ at one turn for this orbit is revealed.

In the calculations, it is taken into account that the oblique triangle OK _З S, in which the angle of blood pressure is determined, exists only when considering points on that portion of the HEO in the apogee region at which the minimum distance Dq from the spacecraft to the nearest GSO point does not exceed the maximum permissible observation range L _max For the rest of the orbit, point S does not exist.

Due to the symmetry of the geometric situation, with the further movement of the spacecraft from the apogee to the next perigee at ω _l = 0, the similar extreme point is also removed from the point A _rc to the other side by the value of DΣ _max . Thus, at one turn, the length of the entire longitude interval, at which the observation distances of objects on the GSO do not exceed L _max , is 2D _Σmax .

In the above calculation of the orbital motion of the spacecraft, both the proper motion of the spacecraft in an elliptical orbit and the rotation of the Earth with stationary points of objects on the GSO are taken into account. The length of the monitored section 2D _Σmax is _thus obtained less than that shown in FIG. 1. At the same time, disturbing elliptical motion factors are not taken into account: the off-center nature of the Earth's gravitational field, the attraction of the Moon, the Sun, etc.

As an example, for visual construction of the spacecraft motion path shown in FIG. 3, relative to the fixed-longitude GSO points, the following current spacecraft coordinates are calculated with a small time step:

- geographical longitude D, degrees;

- latitude W, degrees;

- geocentric distance (radius) r, km.

Due to the fact that the inclination of the orbit is small, the projection of the spacecraft with these coordinates on the equatorial plane can be considered with acceptable accuracy as the path of the spacecraft relative to the fixed points of the geostationary circle.

The given example shows the spacecraft motion path constructed on these coordinates at five turns relative to the geographic points of the geostationary orbit with overlapping control sections. In this example, the motion of the spacecraft in a strictly elliptical orbit (excluding disturbing factors) with a period of 41977 s, a height of perigee 3344 km, an inclination of 3.3 ° is considered. With these parameters, with an acceptable observation range L _max = 400 km, the full coverage of the GSO with control sections for 39 turns, that is, for 19 days, is ensured.

Due to the fact that in the vicinity of the HEO apogee, the speed of the spacecraft is about 1.8 km / s, and the speed of the GSO points is about 3 km / s, the spacecraft on the HEO in the vicinity of the apogee shifts westward relative to the GSO points, and the length of the controlled GSO section decreases by compared to that shown in FIG. one.

In the general case, when calculating the VEO parameters, the non-zero inclination value i is set in the interval from 0 ° to 5 ° and the value of the argument of latitude perigee ω _π = 0. The initial value of the number of days n is taken deliberately small and the period of the highly elliptical orbit T is calculated for it according to the formula T = T _rco ⋅n / (2n + 1).

Strictly speaking, the integer n is the number of stellar days with a period T _gco = 86164 s, and not true days with a duration of 86400 s. With this value of the period T, the highly elliptical orbit is multiply synchronous with respect to the geostationary orbit, which has the period T _GSO , since the duration of n GSO periods is equal to the duration of 2n + 1 HEO periods.

The parameters T and r _ar completely determine the shape of the highly elliptical orbit and make it possible to choose a value of the geographic longitude of the ascending node λ _wu at which the spacecraft longitude at the apogee is zero. _Using these parameters, in the manner described above, the length 2D _{Σmax of the} controlled GSO section at one turn in an angular measure and their total length over 2n + 1 turns are determined. If the total length of the controlled sections does not exceed 360 °, then increase the number of days n by one and repeat the calculations until the desired number n is obtained, at which the minimum excess of 360 ° begins with the total length of the controlled sections of the GSO.

After choosing the HEO parameters described above in a way that minimizes the number of turns when the GSO is completely covered by control sections, they proceed to refine the period T of the highly elliptical orbit. For this, the motion of the spacecraft is predicted taking into account disturbing factors by 2n + 2 turns, the specified geographical longitude of the apogee is fixed on this turn of D _ut and compared with the previous value of D _ap obtained for the motion of the spacecraft in a strictly elliptical orbit. According to these data, the correction of the circulation period ΔT is calculated by the formula ΔT = (D _ut- D _ap ) / (ω _s ⋅ (2n + 1)). Here, ω _s = 0.004178 ° / s is the angular velocity of the Earth's rotation. Then, this correction is made to the previous value T _pr of the revolution period calculated for a strictly elliptical orbit, and the final value of the multiple-synchronous revolution period T = T _pr + ΔT is obtained.

To obtain detailed images with the highest resolution when observing an object located at any point in the GSO, the KA observer is brought closer to it at the minimum permissible distance determined by the characteristics of the observation equipment. In this case, the period of revolution is corrected and transferred to a highly elliptical orbit with a period T different from the multiple-synchronous one and given by the formula

T = T _gco ⋅n / (2n + 1) + Δt.

Here Δt is the addition to the above-described period of the multiple-synchronous orbit, calculated in a certain number of seconds and causing a slow displacement of the path of this orbit in geographical longitude. With this choice of the VEO revolution period, after a certain time, with a slight correction of the period of revolution of the KA observer, it will come closer to the observed object at the minimum allowable range. Keep in mind that such a rapprochement may take a long time.

According to the proposed method for reviewing the geostationary region with respect to obtaining detailed images of objects at the GSO from a multiple-synchronous orbit, calculations were performed for two options for the maximum permissible range for obtaining detailed images: L _max = 100 km and L _max = 400 km.

The calculations showed that the proposed method for the review of the geostationary orbit can reduce the time for a full review of the GSO in 6-10 times compared with the prototype.

It follows from the foregoing that the proposed survey method has an advantage over the known methods for surveying the geostationary region and obtaining detailed images of objects in the geostationary orbit from spacecraft in near-earth geocentric orbits.

Information sources

1. I. Black. High Flight Spy. Cosmonautics News. No. 5. 2014.S. 62-63.

2. I. Lisov. "Angels" and "guards" for the geostationary. Cosmonautics News. No. 9. 2014.S. 54-57.

3. I. Black. Launched a satellite monitoring the space situation. Cosmonautics News. No. 11. 2010.S. 34-36.

4. N.N. Klimenko, A.E. Nazarov. A promising space system for observing a geostationary orbit. Bulletin of the NGO named after S.A. Lavochkina. Number 4. 2015.S. 16-22.

Claims (2)

1. A method for reviewing the geostationary region for observing space debris and other objects from a spacecraft in a semi-diurnal, highly elliptical orbit, characterized in that the spacecraft (SC) is placed in a semi-diary highly elliptical orbit with an apogee, the distance from which to the geostationary orbit does not exceed the maximum permissible range upon receipt of detailed images, with a perigee height of up to 5000 km, with an inclination in the range from 0 to 5 °, moreover, the specific values of these parameters of the spacecraft orbit are chosen within the indicated limits, depending on the characteristics of the observation equipment used to obtain detailed images of objects in the geostationary orbit, and the revolution period T of the semidiurnal highly elliptical orbit is preliminarily set by the multiple synchronous period of the geostationary orbit T _GSO according to the formula T = T _GSO ⋅n / (2n + 1), where n is a variable number of days, for each value of n determine the length of the monitored section in the geostationary orbit, in which the observation range does not exceed the maximum permissible upon receipt of detailed and of images, find the minimum number of days n necessary for a complete continuous coverage of the geostationary orbit in controlled areas, take into account disturbing factors, make 2n + 2 turns of the orbital motion of the spacecraft with a preliminary value of the period T and specify the magnitude of the period T so that the geographical longitude of the apogee on the turn with serial number 2n + 2 coincided with the longitude of the apogee of the first turn, while the coordinates of the objects observed in the geostationary region are measured by on-board observation equipment with spacecraft selected areas of the highly elliptical orbit according to the tasks that are transmitted from ground control points during periods of direct visibility of the spacecraft or according to programs installed on the spacecraft, and the information received and stored in the on-board storage device is transmitted to ground-based points of reception and processing information from the spacecraft during periods of its direct visibility. 2. The method according to claim 1, characterized in that the Δt value is added to the period T, the value of which is selected so that any point on the GSO is observed from the spacecraft after a certain number of days at the minimum allowable distance determined by the characteristics of the observation equipment.

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