RU2660090C1 - Method of the celestial sphere space surveillance system for approaching from the sun celestial bodies detection and threatening collision with the earth - Google Patents

Abstract

FIELD: astronautics.

SUBSTANCE: invention relates to the satellite systems for the Solar system celestial bodies detection, observation and monitoring, threatening collision with the Earth. Method comprises placing two spacecrafts with telescopes T₁ (SC T₁) and T₂ (SC T₂) on the Earth's orbit (2) around the Sun (1). Both SC are rotated at a constant angular velocity around their longitudinal axes, for example, around the line connecting T₁ and T₂. Telescopes fields of view describe conical surfaces. With the celestial body (3) detection in the telescope field of view (for example, SC T₂), the of the two spacecrafts rotation is terminated, transferring the SC T₂ into the tracking mode, and the radiation receivers is into the signal recording mode, constantly directing the T₂ telescope field of view on the body (3). Orienting the second SC T₁ telescope T₁ field of view to the first SC T₂ and deploying in the plane T₁-(3)-T₂ until the body (3) appearance in the telescope T₁ field of view. Tracking the celestial body (3) by both telescopes, performing measurements to determine its orbit parameters.

EFFECT: reduction in the SC formation and increase in the system operation speed.

4 cl, 3 dwg, 3 tbl

Description

The invention relates to space technology and can be used to create near-Earth space observation systems with space-based devices for detecting, observing, and monitoring the celestial bodies of the solar system, including asteroids and comets approaching from the side of the sun and threatening a collision with the Earth.

The creation of space-based observation and monitoring tools to solve the problem of asteroid-comet safety solely at the expense of space-based devices that review the entire celestial sphere is inexpedient, since the cost of creating space telescopes is two orders of magnitude higher than the cost of ground-based ones [1]. A more rational way is to create space systems that complement the terrestrial telescope system in that part where the observation of dangerous celestial bodies approaching the Earth is not possible from the earth's surface. In this regard, the urgent task of creating effective, least costly and high-speed space warning systems. The use of modern information transmission systems and its processing can not always provide the required speed and efficiency of the space system. The most effective way to increase the speed of the space warning system for asteroid-comet hazard is to reduce the flow of primary information to be processed. In this case, this is possible due to the restriction of the field of view of the celestial sphere by space means only to those that are not available for observation by ground-based means. This is, first of all, the field of view of the possible approach of asteroids and comets from the side of the Sun.

The invention is known - an analogue of “A method for mapping the celestial sphere and a spacecraft for its implementation”, RF patent No. 2014252 IPC5 B64G 1/00, which implements viewing the annular bands of the celestial sphere in the visible range of the spectrum using two observation tubes located at some angle to each other. When one observation tube is oriented towards a known reference star, the spacecraft is rotated relative to this direction until the second observation tube of the object falls into the field of view, the position of which must be entered into the catalog. The method provides a measurement of the angular distances between two stationary luminous objects in the celestial sphere. The number of reference stars and their coordinates are selected in advance, before the launch of the spacecraft.

The disadvantage of this method is that it does not provide the detection and measurement of the parameters of the orbits moving to the Earth of celestial bodies, since there is no way to determine in advance reference points to which the observation tube must be directed when viewing the celestial sphere to detect and track celestial bodies moving to Earth along unknown paths in advance.

There is a known method - an analogue of “Asteroid-Comet Hazard Monitoring Method”, RF patent No. 2573509, B64G 4/00 (MKP 2006.01), 2014, in the implementation of which spacecraft are placed at the Lagrange points L1, L2 and L4 or L5 of the Sun system - Earth.

The disadvantage of this method is the significant remoteness of the spacecraft located at the Lagrange points L4 or L5 from the Earth (150 million km), which creates significant difficulties in real-time information exchange with ground-based points for receiving and processing operational information about asteroid-comet hazard, in particular with a short time for the early detection of previously unknown small celestial bodies approaching from the side of the Sun. The dimensions of such bodies, which constitute a real danger to the Earth’s biosphere, can be 50 m or less, and the time of their approach to the Earth can be no more than 1 day, so the task of timely detection and warning of the approach of such bodies remains very difficult.

The disadvantage of this method is that in the vicinity of points L4 and L5 there are the remains of a gas-dust cloud of the era of formation of the planets of the Solar system, which can create real interference and form false targets when viewing the celestial sphere when solving the problem of detecting small celestial bodies approaching the Earth from the side The Sun (http://fiz.1september.ru/view_article.php?ID=201000910). In 2010, a sufficiently large asteroid was discovered at the fourth point of Lagrange. No large space objects have been found at the fifth point of Lagrange at the moment, but recent studies confirm that there is a large accumulation of interplanetary dust.

The famous "International Aerospace Global Monitoring System (MAXM)" is a prototype. Patent No. 2465729 B64G 99/00 (MKP 2009.01), 2010. The space segment of the system is implemented in such a way that two spacecraft with infrared telescopes on board are placed in orbit of the Earth ahead of and behind it at distances of ~ 0.1 AU . and 0.7 AU, respectively, direct the axis of rotation of the spacecraft to the Earth, rotate both spacecraft around their own axes at a constant angular velocity, the telescopes are placed at certain angles to the axis of rotation of the spacecraft and thereby create a barrier field of view around the Earth and when a celestial body moving toward the Earth is detected, information about its motion is transmitted to a spacecraft with a long-focus infrared telescope on board located at the Lagrange point L1 of the Sun-Earth system, the lengths are sent nofocus telescope on the detected celestial body and determine its angular position relative to all three telescopes, while telescopes use charge-coupled devices (CCD) in the mode of time delay and accumulation (WZN) of charge as radiation receiver.

The disadvantages of this space system is the need for an infrared telescope at the Lagrange point L1, which significantly reduces the active life of a specialized space system for operational monitoring of asteroid and comet hazard and depends on the initial volume of refrigerant for cooling the elements of the telescope optical circuit. The disadvantage is that during the operation of infrared telescopes in rotation mode to create a barrier zone, the sun is inevitably exposed to telescope elements located in front of the entrance pupil of the telescope optical circuit, for example, parts of the inner surfaces of the blends of infrared telescopes, which causes them to heat up and cause background illumination of the optical path of the infrared telescope in its working range of the spectrum. Most exposure affects the functioning of a telescope located at a distance of 0.1 AU from the earth. This leads to a decrease in the signal-to-noise ratio at the radiation receiver and, accordingly, to a deterioration of the threshold sensitivity of the infrared telescope. The most effective way to eliminate this drawback is to orient the infrared telescope in a plane perpendicular to the direction to the Sun and in the direction from the Earth, but in this case it is impossible to create a barrier zone around the Earth using infrared telescopes. So, for example, the WISE telescope is in a sun-synchronous orbit, and its optical axis is always directed in the opposite direction relative to the direction to the center of the Earth and is 90 ° away from the direction to the center of the solar disk (https://ru.wikipedia.org/ wiki / Wide-Field_Infrared_Survey_Explorer). This is what ensures its effective functioning in the infrared range due to the lack of exposure from the sun and optimal conditions for the economical use of the refrigerant stored at startup.

At the same time, in view of the considerable remoteness of the infrared telescope at the Lagrange point L1, there are significant difficulties in real-time information exchange with ground-based points for receiving and processing operational information, since a specialized relay space vehicle is probably necessary.

The aim of the invention is a method of constructing a space system for viewing the celestial sphere to detect celestial bodies approaching from the side of the Sun and threatening a collision with the Earth, is to reduce the grouping of spacecraft due to the lack of the need to place an additional spacecraft at the Lagrange point L1 and increase the speed of the space system due to reducing the review of near-Earth outer space only in the space between the Earth and the Sun (Fig. 1).

This goal is achieved by the fact that the axis of rotation of the spacecraft T ₁ and T ₂ (Fig. 2) located in orbit of the Earth ahead of and behind, are in the range from 0.2 to 0.5 astronomical units and sent to the calculated points K ₁ and K ₂ located also in the Earth’s orbit in such a way that the telescopes' fields of view, located at certain angles to the axis of rotation of the spacecraft, describe conical surfaces in the space between the Earth and the Sun, and when a celestial body is detected in the telescope field of view of one of the space their spacecraft, the rotation of both spacecraft is stopped, the spacecraft that has detected the celestial body is in tracking mode, and the field of view of the telescope of this spacecraft is constantly directed to the detected celestial body, the telescope of the second spacecraft is also put into tracking mode of the detected celestial body, while orient the field of vision of the telescope to the first spacecraft and carry out a turn in a plane containing both spacecraft and the detected sky the other body before it appears in the field of view of the second telescope, and in the mode of joint tracking of the detected celestial body with both telescopes, they perform triangulation measurements taking into account the distance D between the spacecraft to determine the orbit parameters of the detected celestial body.

If the axis of rotation of the spacecraft T ₁ and T ₂ for implementing the method, directed to the points K ₁ and K ₂ in the orbit of the Earth, which coincide with the locations of the spacecraft T ₂ and T _1, respectively, then the axis of rotation of the spacecraft are directed along the line connecting these spacecraft.

The distance D between telescopes located in the Earth’s orbit for the implementation of the method is determined by the ratio

where ΔA is the required accuracy in determining the position of a celestial body in near-Earth space;

L is the distance to the celestial body;

Δα is the angular resolution of the telescope.

Two spacecraft T ₁ and T ₂ (Fig. 2) with telescopes on board, placed in the orbit of the Earth's rotation around the Sun at a distance D from each other, which is 10 ³ - 10 ⁴ times greater than the possible distances between ground-based telescopes. From here a priori follow the advantages of synchronous basis observations made by space telescopes in the tracking mode of a celestial body detected by one of the telescopes.

Rhombus A ₁ A ₂ A ₃ A ₄ is a possible zone of uncertainty in the position of the detected celestial body.

и

- минимальное и максимальное значения квазидальностей от телескопа Т₂ до точек А₁ и А₃, а величины L₁ ⁽¹⁾ и L₃ ⁽¹⁾ - минимальное и максимальное значения квазидальностей от телескопа T₁ до точек А₁ и А₃ ромба неопределенности положения обнаруженного небесного тела; ϕ₁ и ϕ₂ - максимальное и минимальное значения угла между квазидальностями от небесного тела до телескопов T₁ и Т₂; α⁽¹⁾ и α⁽²⁾ - углы между базой D и минимальными квазидальностями L₁ ⁽¹⁾, L₁ ⁽²⁾; ρ - расстояние между серединой базы D и А_х - средним возможным положением небесного тела, являющимся центром ромба.Quantities

and

- the minimum and maximum values of quasidalities from the telescope T ₂ to points A ₁ and A ₃ , and the values L ₁ ⁽¹⁾ and L ₃ ⁽¹⁾ - the minimum and maximum values of quasidalities from the telescope T ₁ to points A ₁ and A _{3 of the} uncertainty rhombus the position of the discovered celestial body; ϕ ₁ and ϕ ₂ - the maximum and minimum values of the angle between quasidalities from the celestial body to the telescopes T ₁ and T ₂ ; α ⁽¹⁾ and α ⁽²⁾ are the angles between the base D and the minimum quasidalities L ₁ ⁽¹⁾ , L ₁ ⁽²⁾ ; ρ is the distance between the middle of the base D and A _x is the average possible position of the celestial body, which is the center of the rhombus.

As the standard error of determining the range, take half the maximum size of the rhombus of the possible positions of the celestial body, that is, the distance ΔA to be determined. The error ΔА depends on the distance D between the telescopes, the angles α ⁽¹⁾ and α ⁽²⁾ and the angular measurement error Δα = Δα ⁽¹⁾ = Δα ⁽²⁾ .

From FIG. 1 follows the relationship:

From the triangle T ₂ A ₁ A ₃ we get

From the triangle T ₂ A ₁ T ₁ we have

From the triangle T ₂ A ₃ T ₁ we get

We believe that the error ΔА in order of magnitude is equal to half the maximum difference of the quasidalities L ₃ ⁽²⁾ and L ₁ ⁽²⁾ , since the angle Δα is very small. Then

This expression can be coded as

Due to the small angle Δα, the expression (cosΔα - cos2 Δα) ~ 3/2 Δα ² and is equal to 0.

Using the representations for sin ϕ ₁ = sin (α ⁽¹⁾ + α ⁽²⁾ ) and for cosϕ ₁ = -cos (α ⁽¹⁾ + α ⁽²⁾ ), we obtain the formula

By the cosine theorem, we find the angles α ⁽¹⁾ and α ⁽²⁾

The standard error of determining the distance to the celestial body ΔA is determined by the known values of D, the error Δα and the quantities L ₁ ⁽²⁾ and L ₁ ⁽¹⁾ .

Given the initial data for the case α ⁽¹⁾ = α ⁽²⁾ = α, ΔA has the form

Considering that if the angles α ⁽¹⁾ and α ⁽²⁾ are equal, the relation L⋅cos α = D / 2 is satisfied, we obtain:

If α ⁽¹⁾ = α ⁽²⁾ = α and the celestial body is at a distance L = L ₁ ⁽ 2 ⁾ = L ₁ ⁽¹⁾ from the observers, which is orders of magnitude greater than the base D between the telescopes, then α ~ π / 2 then

To increase the noise immunity of the space system according to the claimed method, the spectral range of the optical path of space telescopes is the visible range, since radiation outside the field of view can be effectively eliminated with a lens hood, and heating of the internal parts of the lens hood from solar exposure produces radiation in the range outside the spectral sensitivity of the space telescope.

To determine the possible negative consequences of the fall of a celestial body on the earth's surface, it is important to determine the physico-chemical characteristics of the detected celestial body, threatening a collision with the Earth. For this, when implementing the claimed method, on board the spacecraft, in addition to telescopes, measuring equipment is installed to determine the physicochemical characteristics of the celestial body.

и

между космическими аппаратами и Землей зависит от затрат характеристической скорости для вывода космических аппаратов на орбиту Земли

и

и требуемого запаса характеристической скорости V_x1 и V_x2 для поддержания неизменным расстояние между Землей и космическими аппаратами T₁ и Т₂.Optimal distance selection

and

between spacecraft and the Earth depends on the cost of the characteristic speed for putting spacecraft into orbit of the Earth

and

and the required margin of characteristic velocity V _x1 and V _x2 to maintain constant the distance between the Earth and spacecraft T ₁ and T ₂ .

a.e. характеристическая скорость для вывода космического аппарата на орбиту Земли составляет:The most energy-consuming is the launch of the T1 spacecraft ahead of the Earth [2] and depending on the distance

ae the characteristic speed for the launch of the spacecraft into orbit of the Earth is:

For a spacecraft with a lag from the Earth, the costs of the characteristic velocity are:

For distances from 0.2 AU up to 0.35 au the costs of the characteristic speed for launching spacecraft into the Earth’s orbit do not increase significantly.

и

на них действует возмущающее ускорение Земли

. При небольших расстояниях

и

возмущающее ускорение практически перпендикулярно вектору центростремительного ускорения

от Солнца к Земле, где V_З - скорость обращения Земли вокруг Солнца, V_З = 29,85 км/с, радиус орбиты Земли R_oЗ = 1 а.е. = 1,5⋅10⁸ км. Ускорение q_c = 6⋅10^-3 м/с².When the spacecraft operates in the Earth’s orbit at distances

and

they are affected by the disturbing acceleration of the Earth

. At short distances

and

perturbing acceleration is almost perpendicular to the centripetal acceleration vector

from the Sun to the Earth, where V _З is the velocity of the Earth around the Sun, V _З = 29.85 km / s, the radius of the Earth’s orbit R _oЗ = 1 _a.u. = 1.5⋅10 ⁸ km. Acceleration q _c = 6⋅10 ^-3 m / s ² .

практически нормально к q_c Provided that

almost normal to q _c

- расстояние между Землей и КА на ее орбите в км.where: q _З = 9.81 m / s ² , R _З - radius of the Earth, 6.37⋅10 ³ km,

- the distance between the Earth and the spacecraft in its orbit in km.

и

необходимо производить коррекцию орбит космических аппаратов T₁ и Т₂. Характеристическая скорость, необходимая для компенсации возмущения в течение срока активного функционирования космического ппарата, равного τ_ас ~ 10 лет, равнаTo compensate for disturbances that tend to reduce distances

and

it is necessary to correct the orbits of the spacecraft T ₁ and T ₂ . The characteristic speed necessary to compensate for the disturbance during the period of active operation of the spacecraft equal to τ _as ~ 10 years is

Dimension τ _as in seconds, a V _x in m / s.

наилучшим образом аппроксимируется на отрезке

полиномиальным уравнением 6-й степени следующего видаFunction of characteristic speed V _x from distance

best approximated on a segment

6th degree polynomial equation of the following form

The reliability of the approximation R ² = 1.

представлен на фиг. 3.Dependency graph

shown in FIG. 3.

а.е. требуемый запас характеристической скорости на поддержание орбиты равен 22 м/с. При

а.е. требует увеличения запаса характеристической скорости на 117 м/с до 139 м/с, что является приемлемой величиной. Расположение космического аппарата ближе 0,2 а.е. к Земле требует значительного увеличения запаса характеристической скорости и энергозатрат на поддержание космического аппарата на заданном расстоянии от Земли на ее орбите. Так приближение космического аппарата к Земле на 0,05 а.е. от 0,2 а.е. до 0,15 а.е. требует увеличения запаса характеристической скорости на 109 м/с (от 139 м/с до 248 м/с.From the above calculations it follows that for

a.e. the required margin of characteristic velocity to maintain the orbit is 22 m / s. At

a.e. requires an increase in the characteristic velocity margin of 117 m / s to 139 m / s, which is an acceptable value. The location of the spacecraft is closer than 0.2 AU to the Earth requires a significant increase in the reserve of characteristic speed and energy consumption for maintaining the spacecraft at a given distance from the Earth in its orbit. So the approach of the spacecraft to the Earth at 0.05 AU from 0.2 AU up to 0.15 au requires an increase in the characteristic velocity margin of 109 m / s (from 139 m / s to 248 m / s.

Removing the spacecraft by more than 0.5 AU It requires the availability of spacecraft-transponders, leads to communication difficulties with ground-based information processing points and reduces the speed of the space-based operational warning system about a possible collision of a celestial body with the Earth.

Analyzing the data presented and the calculations, we can conclude that the most rational in terms of energy consumption and maintaining communication with the spacecraft is to place the spacecraft in orbit of the Earth at distances of 0.2 ÷ 0.5 a.u. from her.

FIG. 2. Scheme of measurements of motion parameters of a detected celestial body

T ₁ - telescope 1;

T ₂ - telescope 2;

D is the distance between the telescopes;

A ₁ A ₂ A ₃ A ₄ - the region of uncertainty of the position of the celestial body;

And _x is the possible position of the celestial body in the region of uncertainty;

ρ is the distance from the point A _X to the middle of D;

- расстояния от телескопов до граничных точек А₁ и А₃ области неопределенности;

- the distance from the telescopes to the boundary points A ₁ and A ₃ areas of uncertainty;

α ⁽¹⁾ is the angle between the directions T ₁ T ₂ and T ₁ A ₁ ;

α ⁽²⁾ is the angle between the directions of T ₂ T ₁ and T ₂ A ₁ ;

ϕ ₁ is the angle between the directions A ₁ T ₁ and A ₁ T ₂ ;

ϕ ₂ is the angle between the directions A ₃ T ₁ and A ₃ T ₂ ;

Δα is the angular error of a single measurement of the NT position.

FIG. 1. Scheme of the near-Earth space review.

1 - the sun;

2 - Earth;

3 - an asteroid;

4 - the meeting point of the asteroid with the Earth;

T ₁ - spacecraft with a telescope 1;

T ₂ - spacecraft with a telescope 2;

ω is the angular velocity of rotation of the fields of view of telescopes 1 and 2;

α ₁ and α ₂ are the angles between the axes of rotation of the spacecraft T ₁ and T ₂ and the axes of the telescopes mounted on them;

Q is the angle of the cone of invisibility by ground-based telescopes;

K ₁ and K ₂ are the calculated points in the Earth’s orbit for the orientation of the rotation axes of the spacecraft T ₁ and T ₂ ;

β ₁ , β ₂ are the angles of the field of view of telescopes 1 and 2.

для поддержания космического аппарата на орбите в зависимости от дальности

до Земли.FIG. 3 Required margin of characteristic speed

to maintain a spacecraft in orbit depending on range

to Earth.

Literature

1. K.N Sviridov. Optical location of space debris. - M: Knowledge, 2006, pp. 413-437.

2. V.A. Emelyanov, S.S. Klimov. The use of space telescopes to detect dangerous celestial bodies and determine the parameters of their orbits. Space and rocket science, No. 2 (63), 2011, pp. 100-105.

Claims (8)

1. A method of constructing a space system for observing the celestial sphere, in which two spacecraft T ₁ and T ₂ are placed in orbit around the sun around the sun, with one spacecraft ahead of the earth and the second behind it, a telescope is installed at each spacecraft at a certain angle to the longitudinal axis of rotation of the spacecraft, while in the search and search mode of celestial bodies approaching the Earth from the side of the Sun in near-Earth space, both spacecraft rotate simultaneously with a constant angular velocity around their own longitudinal axes, in telescopes as radiation detectors used charge coupled devices in the mode of time delay and integration charge, characterized in that the spacecraft axis of rotation T ₁ and T ₂ is sent to the points K ₁ and K ₂ are located on the orbit Earth in such a way that the telescopes' fields of view describe conical surfaces in the space between the Earth and the Sun, and when a celestial body is detected in the telescope's field of view of one of the spacecraft, the rotation of both spacecraft ceases they’ll take the spacecraft that detected the celestial body into tracking mode, the radiation receivers are put into the frame recording mode of the signal, provide a constant direction of the field of view of the telescope that detected the celestial body of the spacecraft to the detected celestial body, and the field of view of the telescope of the second spacecraft is oriented to the first spacecraft and carry out a turn in a plane containing both spacecraft and a detected celestial body, until this celestial body appears in the field of sp Nia second spacecraft telescope, and when the joint accompanied both telescopes detected celestial body operate triangulation measurement given the distance D between the spacecraft orbital parameters for determining the detected celestial body. 2. The method according to claim 1, characterized in that the spacecraft are located at a distance D depending on the required accuracy of determining the position of a celestial body in near-Earth space, while

, where ΔA is the required accuracy in determining the position of a celestial body in near-Earth space; L is the distance to the celestial body; Δα is the angular resolution of the telescope on board the spacecraft. 3. The method according to claim 2, characterized in that the spacecraft are placed in orbit of the Earth at distances of 0.2 ÷ 0.5 AU from her. 4. The method according to claim 1, characterized in that the points K ₁ and K ₂ in the orbit of the Earth are combined with the locations of the spacecraft T ₂ and T _1, respectively, while the axis of rotation of the spacecraft is directed along the line connecting these spacecraft.

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