Method of control over spacecraft descent in atmosphere of planets

FIELD: aircraft engineering.

SUBSTANCE: invention relates to spacecraft control in atmosphere of planet by adjusting its aerodynamics. Spacecraft velocity in atmosphere at initial flight part increases (spacecraft flies toward conditional orbit pericentre). Atmosphere density is low yet to cause notable spacecraft deceleration. As spacecraft reaches atmosphere dense layers its velocity decreases to reach atmosphere enter velocity for angle of roll (γ) γ=π to be changed to γ=0. This manoeuvre allows changing the spacecraft to flight part with maximum aerodynamic performances. In flight with γ=0 continuous skip path is maintained whereat spacecraft velocity decreases monotonously. Maximum skip height reached, angle of attack o spacecraft increases, hence, spacecraft intensive deceleration occurs.

EFFECT: decreased final velocity at soft landing system operation, fuel savings.

1 dwg

 

The invention relates to space exploration, in particular to control the descent of a SPACECRAFT (SC) in the atmosphere of the planet, using controlled aerodynamic braking and minimizing the final speed of the spacecraft.

A known method of controlling the descent of a spacecraft in the atmosphere of the planet, using controlled aerodynamic braking and providing the lower end of the speed of the spacecraft, described in the book - Ivanov N. M., Martynov A. I. "the Motion of spacecraft in the atmospheres of the planets. M.: Nauka, 1985, pp. 168-173 - [1]. This method is to control the aerodynamic braking by changing the roll angle γ of the spacecraft. The motion of the spacecraft is carried out with a constant value of angle of attack corresponding to maximum aerodynamic balancing. The method involves one-time switching of the roll angle γ with values equal to π glad to zero, which corresponds to the switching of the effective aerodynamic qualities from the minimum value (-K) to the maximum (+K).

The disadvantage of this method is that it does not control the angle of attack of the spacecraft α. This significantly reduces the possibility of quenching speed, because the ku increasing the angle of attack α leads to an increase in the intensity of the aerodynamic drag of the spacecraft.

The closest to the technical nature of the claimed method of controlling the descent of a spacecraft in the atmosphere of the planet, using controlled aerodynamic braking and providing the lower end of the speed of the spacecraft is described in the book - Ivanov N. M., Martynov A. I. "motion Control of spacecraft in the atmosphere of Mars. Moscow, Nauka, Main editorial Board for physical and mathematical literature, 1977, pages 192-213 - [2], which is selected by the prototype. This method is the two-parameter control roll angles and attack of the spacecraft. The entry of the spacecraft into the atmosphere of the planet is roll angle γ=π rad and angle of attack α, corresponding to the maximum value of the balancing aerodynamic quality. On the initial flight phase switches of the roll angle γ to zero. After reaching the angle of the velocity vector to the local horizon zero value of the roll angle γ is determined from the condition for the provision of space vehicle flight on izolyatsia the plot (the plot with constant height). Then you can switch the roll angle γ to zero, providing the movement of the spacecraft on ricochetnaya trajectory with increasing altitude. On this site happens HC is the increase of angle of attack α from the value corresponding to the maximum aerodynamic lift coefficient, to the value corresponding to the maximum aerodynamic coefficient of drag.

The disadvantages of this method are that it does not fully implemented the provisions of the control spacecraft on the effective damping rate. This is due to the following factors. First, there is no justification for the optimal moment of switching of the roll angle from π glad to zero pleased by the criterion of minimizing the final speed. At this time switch has a significant impact on the dynamics of the braking KA on subsequent phases of flight. Secondly, the use of isonicotinic plots shortens the total duration of the trajectories of descent and thereby reduces the efficiency of inhibition of CA in the atmosphere. Thirdly, control of the angle of attack occurs at relatively short plot: at the beginning of this section, for example, during the descent in the atmosphere of Mars flight speed is about 1 km/s and a height of less than 10 km. however, earlier management of the angle of attack also allows you to increase the intensity of the damping rate due to the increase in aerodynamic drag of the SPACECRAFT.

The technical result of the proposed method of control the descent of the SPACECRAFT in the atmosphere of the planet is reduced of course the speed when commissioning the soft landing system at the expense of rational management of roll angles and attack. This gives the possibility to reduce fuel consumption on the implementation of the soft-landing a SPACECRAFT on the surface of the planet. The application of the proposed method, depending on the design of the ballistic characteristics of the spacecraft, boundary conditions and parameters of the planet of destination allows to reduce fuel consumption ~ 10-20% compared to using the prototype method.

The invention consists in the rational management of the roll angles of attack which minimizes the final speed of the flight SPACECRAFT. This is achieved by introducing a new episode of the control compared to the prototype. First, the choice of appropriate conditions switching of the roll angle from π to zero pleased: the switching is performed when the speed of the AC becomes less than the speed of its entry into the atmosphere of the planet. As is known, the velocity of the SPACECRAFT at the initial part of flight in the atmosphere increases as the apparatus moves in the direction of the conditional pericenter oscillating orbits, and the density of the atmosphere is still relatively small and has no significant effect on the inhibition of AC. Then, when reaching the KA dense layers of the atmosphere, its speed begins to decrease and at a certain point in time is reduced up to speed entry into the atmosphere. It was at this point it is necessary to perform switching of the roll angle is π glad to zero and transfer the SPACECRAFT into a trajectory with maximum glide. In an earlier switching angle γ (when flight speed KA, greater speed entry into the atmosphere), the resulting lifting force can lead to the departure of the SPACECRAFT from the atmosphere and to the failure of the main tasks of the space mission - planting apparatus in a predetermined area of the surface of the planet. Later switching roll angle γ leads to a decrease in the duration of the descent trajectory and, consequently, to reduce the intensity of the braking KA. Second, a favorable factor for the effectiveness of reducing the final speed is earlier compared to the prototype, the beginning of the control angle of attack α. After putting the SPACECRAFT on flight mode with γ=0 rad implemented long nicoletiidae trajectory, where the velocity of the SPACECRAFT decreases monotonically. At the maximum height of the rebound, for example ~50-60 km in terms of reduction in the atmosphere of Mars and ~200-250 km with a reduction in the atmosphere of Jupiter, there is an increase in the angle of attack and, therefore, more intensive braking KA. Such a condition is the beginning of the control angle α is rational for the following reasons: the earlier the change in the angle α (to complete the set maximum height) decreases the maximum height of rebound reduce the duration of the flight and to reduce the integral effects of atmospheric drag on the SPACECRAFT. When more than the late introduction of the control angle α (at lower altitude KA) device as a rule, no time to repay the speed at the end of the descent trajectory to the lowest possible values.

The essence of the claimed method of controlling the descent of a spacecraft in the atmosphere of the planet lies in its spatial orientation and management of aerodynamic deceleration, stabilization of the spacecraft during entry into the atmosphere of the planet on the roll angle γ equal to about π rad, and angle of attack α for maximum aerodynamic performance of the spacecraft, determining the current values of velocity, density of the atmosphere and the altitude of the spacecraft, the establishment of the roll angle γ is equal to about 0 rad in the process of deceleration of the spacecraft in the atmosphere of the planet, in the implementation of the motion of the spacecraft in the atmosphere of the planet with the subsequent commissioning of the soft landing of a space vehicle, this set the roll angle γ of the spacecraft equal to about 0 rad, providing movement of the spacecraft on ricochetnaya trajectory with increasing altitude, in the process of deceleration of the spacecraft in the atmosphere of the planet when the condition is met:

Vi<VI,

where: Vi- the current value of the speed of the spacecraft during its deceleration in the atmosphere of the planet;

VI- speed WMO is the spacecraft into the atmosphere of the planet,

take further motion of the spacecraft on ricochetnaya trajectory

and when the condition is met:

hi<hmax,

where: hi- the current value of the altitude of the spacecraft in the atmosphere of the planet;

hmax- maximum flight altitude of the spacecraft during its movement on ricochetnaya trajectory

set the value of the angle of attack α of a spacecraft in accordance with the mathematical expression:

,

where:

;;

αi- the angle of attack α of the spacecraft at time ti;

Vi- the current value of the velocity of the spacecraft at time ti;

ρi- the density of the atmosphere of the planet at time ti;

Δtithe time intervals between subsequent measurements, i=1, 2, 3, ...;

Cx- aerodynamic drag coefficient of the spacecraft;

S - the area of the fuselage mid-section of the spacecraft;

m is the mass of the spacecraft;

β is the logarithmic rate of change of the density of the atmosphere from a height;

l, n - constant coefficients of the approximation of the dependency of the aerodynamic coefficients on the angle of attack of the spacecraft to the analytical mind;

<> a1, a2, a3- constant coefficients, obtained by integration of differential equations of the adjoint variables,

upon reaching the angle of attack α of the spacecraft of α*, corresponding to the maximum value of the aerodynamic coefficient of drag, fly a mission with this value of angle of attack α* before commissioning the soft landing system.

The claimed method of controlling the descent of a spacecraft in the atmosphere of the planet is illustrated by a drawing, which shows the dependence of the velocity V, the height h of the spacecraft, its roll angles γ and attack α from the time the motion in the atmosphere of Mars t while minimizing the final speed.

In addition, the drawing and the text made the following notation: tp- the time of switching of the roll angle with π rad 0 rad.

According to [2], page 194 aerodynamic coefficients of drag and lift forces with a high degree of accuracy can be approximated by the following analytical dependencies:

Cx=Cx0+Asin2(nα+l),

Cy=Cy0+Asin(nα+l)cos(nα+l).

In particular, when using forms spacecraft type carrying case: Cx0=0,2; Cy0=-0,1; A=2,3; n=1,125; l=5,625°.

For other types of forms can be used similar dependencies with other values of the s coefficients - [2], page 194.

The technical result of the invention is to reduce the required energy costs for the implementation of space missions to study the planets of the Solar system and, consequently, increasing the share of the payload in the overall weight balance of the spacecraft.

This technical result is achieved due to the installation aboard the landers management system glide and development of rational management programs roll angles and attack KA, namely due to the fact that in the method of controlling the descent of a spacecraft in the atmosphere of the planet, the selected prototype consists of a spatial orientation of the SPACECRAFT and the management of its aerodynamic deceleration, stabilization of the spacecraft during entry into the atmosphere of the planet on the roll angle γ equal to about π rad, and angle of attack α for maximum aerodynamic performance of the spacecraft, determining the current values of velocity, density of the atmosphere and the altitude of the spacecraft, the establishment of the roll angle γ is equal to about 0 rad in the process of deceleration of the spacecraft in the atmosphere of the planet, in the implementation of the motion of the spacecraft in the atmosphere of the planet with the subsequent commissioning of the soft landing of a spacecraft, in addition establish Hugo the roll γ spacecraft, equal to about 0 rad, providing movement of the spacecraft on ricochetnaya trajectory with increasing altitude, in the process of deceleration of the spacecraft in the atmosphere of the planet when the condition is met:

Vi<VI,

where: Vi- the current value of the speed of the spacecraft during its deceleration in the atmosphere of the planet;

VIthe entrance velocity of the spacecraft in the atmosphere of the planet,

take further motion of the spacecraft on ricochetnaya trajectory

and when the condition is met:

hi<hmax,

where: hi- the current value of the altitude of the spacecraft in the atmosphere of the planet;

hmax- maximum flight altitude of the spacecraft during its movement on ricochetnaya trajectory

set the value of the angle of attack α of a spacecraft in accordance with the mathematical expression:

,

where:

;;

αi- the angle of attack α of the spacecraft at time ti;

Vi- the current value of the velocity of the spacecraft at time ti;

ρi- the density of the atmosphere of the planet at time ti;

Δtithe time intervals between subsequent of erenee, i=1, 2, 3, ...;

Cx- aerodynamic drag coefficient of the spacecraft;

S - the area of the fuselage mid-section of the spacecraft;

m is the mass of the spacecraft;

β is the logarithmic rate of change of the density of the atmosphere from a height;

l, n - constant coefficients of the approximation of the dependency of the aerodynamic coefficients on the angle of attack of the spacecraft to the analytical mind;

a1, a2, a3- constant coefficients, obtained by integration of differential equations of the adjoint variables at angle of attack α of the spacecraft of α*, corresponding to the maximum value of the aerodynamic coefficient of drag, fly a mission with this value of angle of attack α* before commissioning the soft landing system.

We will show the possibility of carrying out the invention, i.e., the possibility of its industrial applications. A feature of the conduct of space activities in many countries is to enhance the study of the planets of the Solar system. In the framework of the Federal space program 2016-2025, provided the work to create a spacecraft to explore Mars, Venus, Jupiter, mercury, including design landers. However, one of the major problems of having aetsa develop key technologies management providing massively reduce energy costs in all areas of interplanetary flights. The successful solution of this problem is largely provided when placing aboard the landers control systems aerodynamic braking, using the principles of management of roll angles and attack described in the present invention.

As for the technical means to ensure control glide KA, i.e. the management of its roll angles and attack, they are known - see, for example, [1], page 37, [2], page 57, 270, and Navigation flight of the orbital complex "SALYUT-6" - "SOYUZ" - "PROGRESS"" responsible editors B. N. Petrov, I. K. Bazhinov, Moscow, Nauka, 1985, Chapter 1 - [3].

Notes.

1. The applicant was placed in the Annex to the application materials rationale used them (in the description and the claims) of the mathematical expression to calculate the angle of attack of the SPACECRAFT on its final part of the descent in the atmosphere of the planet that do not unnecessarily overload the description of the invention. However, if the expert deems appropriate, the applicant will not argue for its inclusion in the description.

2. According to p. 2.3.1 of the Guidelines for examination of applications for inventions from 25.07.2011, the use in the claims of the sign "about" in describing the values of the number of the s parameters are valid.

3. The applicant in the application materials have used two identical term "switch" the value of the roll angle γ KA (is used to describe analogues) and "set" the value of the roll angle γ KA (in the claims), as, in his opinion, is more preferable. While believing that the unity of terminology in this case is not broken.

App. Refers to the application for the invention "Method of controlling the descent of a spacecraft in the atmosphere of the planet" (managed using aerodynamic braking and minimizing the final velocity of the spacecraft (note the Applicant)).

Conclusion the mathematical dependences for calculation of angle of attack on the final section of the descent of a SPACECRAFT (SC) in the atmosphere of the planet.

The motion of the SPACECRAFT in the atmosphere of the planet according to the works [1, 2] is described by the system of differential equations in the velocity coordinate system with the influence of gravitational, aerodynamic, centrifugal and Coriolis forces in the assumption of the centrality of the gravitational field is:

Here V is the velocity of the SPACECRAFT, θ is the angle of the velocity vector to the local horizon, ε is the heading angle, r is the radius - vector connecting the center of the planet and the position of the KA, λ and φ is the longitude and latitude of subsatellite points KA, respectively, m is the SPACECRAFT mass, t is time, ρ is the energy density is the ability of the atmosphere, Cxand Cy- aerodynamic coefficients of drag and lift forces, respectively, R is the planetary radius, h is the altitude, g is the acceleration of gravity, µ is the product of the constant of gravity on the mass of the planet, S - the area of the fuselage mid-section.

The values of the control parameters α and γ can vary:

0≤α≤αmax, -π≤γ≤π.

Transform the original equation (1) given the introduction of the assumptions previously used in a number of domestic and foreign operations, in particular in the works [1, 2]:

h<<R, ρ=ρ0exp(-βh), Fto+FC<<Fg<<Fand,

where ρ0- the density of the atmosphere on the surface of Mars, β is the logarithmic rate of change of the density of the atmosphere with height, FtoFCFgFand- Coriolis, centrifugal, gravitational and aerodynamic forces, respectively.

We will consider only the final section of the descending SPACECRAFT, beginning with the time the device maximum height after the flight on ricochetnaya trajectory and ending on the date of entry into force of the soft landing system.

Using these assumptions, considering the motion of the SPACECRAFT in the plane of entry into the atmosphere and taking into account that the target area is descent with zero roll angle, convert the system of equations to the form:

,,,

where M is a piecewise constant function, according to the works [1, 2].

The solution of the problem of finding the optimal SPACECRAFT control with the minimum final speed was performed using the Pontryagin maximum principle. Let us write the Hamiltonian H and the conjugate variables Ψi:

,

,

,.

Comparing the equations for the functions H, Ψ1, Ψ3convert formulas for the adjoint variables as follows:

,.

From the transversality conditions at the end point the SPACECRAFT trajectory, it follows that

Given that the Hamiltonian does not depend explicitly on time of flight, it is lawful to record the equation:

H≡0.

It allows to represent the dependencies for the calculation of the adjoint variables in the form:

,.

Integrating these equations, taking into account formulas (2), we obtain:

,

, Ψ3(t)=a3=const.

Provided continuous measurement of the current values of flight velocity Viand the density of the atmosphere ρipair the by variables with a high degree of accuracy at the time of measurement t ican be calculated by the formula:

where Δtithe time intervals between subsequent measurements.

The dependency analysis for the calculation of conjugate variables taking into account the equality to zero of the Hamiltonian has shown that Ψ1(t) is negative monotonically increasing function, reaching the end point of the trajectory values of -1; Ψ2(t) is a positive monotone decreasing function, reaching the end point of the trajectory values of zero; Ψ3(t) is a constant function with a negative value.

Define the optimal control law the angle of attack of the conditions for the achievement of the extremum of the Hamiltonian:

.

Solving this equation, we obtain:

Given the described character changes adjoint variables Ψ1and Ψ2we come to the conclusion that the expression ∂Cx/∂Cyhas a negative value on all final flight. This corresponds to a monotonic increase of angle of attack: in this case, ∂Cx>0, and ∂Cy<0. Moreover, the intensity change of the angle of attack is increased monotone decreasing the speed of flight of the SPACECRAFT.

According to the works [1, 2] aerodynamic coefficients windshield resisting film to prevent the means and the lifting force with a high degree of accuracy can be approximated by the following analytical dependencies:

Cx=Cx0+Asin2(nα+l),

Cy=Cy0+Asin(nα+l)cos(nα+l).

For landers-type bearing housing Cx0=0,2; Cy0=-0,1; A=2,3; n=1,125; l=5,625°.

Taking into account these dependencies formula (5) is converted to the following:

.

Then the equation for determining the current values of angles of attack in moments of measurement parameters KA can be written as follows:

,

where the variables Ψ1iand Ψ2iare calculated by the formulas (3), (4). The analysis of this equation showed that the angle of attack α in the area of flight KA monotonically increases from α≈45÷50° to α≈70÷85°, which corresponds to the maximum value of the aerodynamic coefficient of drag.

Sources of information

1. Ivanov N. M., Martynov A. I. "the Motion of spacecraft in the atmospheres of the planets. M.: Nauka, 1985, pp. 168-173.

2. N. M. Ivanov, A. I. Martynov, "motion Control of spacecraft in the atmosphere of Mars. Moscow, Nauka, Main editorial Board for physical and mathematical literature, 1977, pp. 159-169.

The method of controlling the descent of a spacecraft in the atmosphere of the planet, lies in its spatial orientation and management of aerodynamic deceleration, stabilization of the spacecraft during entry into the atmosphere of the planet at the Lou roll γ, equal to about π rad, and angle of attack α for maximum aerodynamic performance of the spacecraft, determining the current values of velocity, density of the atmosphere and the altitude of the spacecraft, the establishment of the roll angle γ is equal to about 0 rad in the process of deceleration of the spacecraft in the atmosphere of the planet, in the implementation of the motion of the spacecraft in the atmosphere of the planet with the subsequent commissioning of the soft landing of a spacecraft, wherein installing the roll angle γ of the spacecraft is equal to about 0 rad, providing movement of the spacecraft on ricochetnaya trajectory with increasing altitude, in the process of deceleration of the spacecraft in the atmosphere of the planet when the condition is met:
Vi<VI,
where: Vi- the current value of the speed of the spacecraft during its deceleration in the atmosphere of the planet,
VIthe entrance velocity of the spacecraft in the atmosphere of the planet,
take further motion of the spacecraft on ricochetnaya trajectory
and when the condition is met:
hi<hmax
where: hi- the current value of the altitude of the spacecraft in the atmosphere of the planet,
hmax- maximum flight altitude of the spacecraft during its movement on R is cochairwoman path,
set the value of the angle of attack α of a spacecraft in accordance with the mathematical expression:
,
where:
;,
αi- the angle of attack α of the spacecraft at time ti,
Vi- the current value of the velocity of the spacecraft at time ti,
ρi- the density of the atmosphere of the planet at time ti,
Δtithe time intervals between subsequent measurements, i=1, 2, 3, ...,
Cx- aerodynamic drag coefficient of the spacecraft,
S - the area of the fuselage mid-section of the spacecraft,
m is the mass of the spacecraft;
β is the logarithmic rate of change of the density of the atmosphere with height,
l, n - constant coefficients of the approximation of the dependency of the aerodynamic coefficients on the angle of attack of the spacecraft in analytical form,
a1, a2, a3- constant coefficients, obtained by integration of differential equations of the adjoint variables,
and upon reaching the angle of attack α of the spacecraft of α*, corresponding to the maximum value of the aerodynamic coefficient of drag, fly a mission with this value of angle of attack α* before commissioning the system to me is some planting.



 

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2 dwg

FIELD: physics; control.

SUBSTANCE: invention relates to controlling movement of a spacecraft fitted with a heat radiator and a solar panel. The method includes flying the spacecraft on an orbit around a planet and turning the solar panel in a position corresponding to the alignment of the normal to the working surface of the solar panel with the direction towards the Sun; performing orbital orientation of the spacecraft, where the plane of rotation of the solar panel is parallel to the plane of the orbit of the spacecraft and the solar panel is located relative to the plane of the orbit on the side of the Sun; determining the maximum value of the angle between the velocity vector of the spacecraft and the perpendicular to the transverse axis of rotation of the solar panel, passing through the surface of the radiator; determining the orbital altitude of the spacecraft and the angle between the direction towards the Sun and the plane of the orbit of the spacecraft; based on the orbital altitude and the angle, determining the orbit passes where the duration of the illuminated part of the pass exceeds the difference between the orbiting period of the spacecraft and the required duration of the heat release by the radiator on the pass; on the said orbit passes, when the spacecraft passes through the illuminated part of the pass, the solar panel is turned around the transverse axis of rotation until the intersection of the line passing through the region of the surface of the radiator facing the Sun and directed towards the Sun with the solar panel; turning the solar panel around the longitudinal axis of rotation until the angle between the normal to the working surface of the solar panel and the direction towards the Sun assumes a minimum value. The said solar panel rotations are performed within a calculated time interval.

EFFECT: high efficiency of the radiator by creating conditions for natural cooling thereof during eclipse of the solar panel for any altitude of an almost circular orbit of the spacecraft.

5 dwg

FIELD: transport.

SUBSTANCE: invention relates to space cable systems (SCS) and can be used for the transfer of SCS to a spinning mode in the orbit plane without the application of jet engines. SCS development is executed from its initial compact state in the circular orbit by the repulsion of objects at a low relative speed. SCS end weights are connected by a cable, its length being varied by a cable feed-haul-in device arranged on one of the end objects. The objects are separated by a vector of local peripheral speed, for example, by a pusher. The objects are driven by a start pulse to separate the objects in practically free paths at the free feed of the cable. The cable development is terminated by the SCS transfer to a stable mode of associated pendulum motion at the stretched preset-length cable. At a definite range of angular phases of this mode the SCS objects are stretched by hauling in the cable at a definite constant speed. This results in changing the SCS into the spinning mode at a preset power integral and fixed final end of the cable.

EFFECT: relaxed weight-size constrictions of SCS, enhanced performances.

8 dwg

FIELD: physics, navigation.

SUBSTANCE: group of the inventions relates to control of angular motion of space vehicle (SV). The method includes additional generation of signals for assessment of orientation angle and angular velocity of rotation of space vehicle. Also the reference signals of the orientation angle, angular velocity and control assessment signal are generated. For the named orientation angle and angular velocity their differences with their assessed signals, and also the difference with their reference values are determined. The difference of control signal and its assessed value and, at last, the signal of correction of the signal of assignment of mathematical model and the signal of assessment of external noises using the respective formulas are determined. On this base the signals of assessment of orientation angle and angular velocity of space vehicle are determined, which are used for space vehicle control. The device in addition contains the reference model of the basic circuit of orientation of space vehicle and other necessary devices and connections.

EFFECT: improvement of orientation accuracy and operational reliability in case of failures of orientation angle sensor and sensor of angular velocity of space vehicle rotation.

2 cl, 2 dwg

FIELD: physics.

SUBSTANCE: invention relates to controlling orientation of an artificial earth satellite with solar panels. The disclosed method includes performing necessary turning of the artificial earth satellite along with solar panels and, separately, the solar panels about a first and a second axis. The antenna of the artificial earth satellite is directed towards the earth and the normal to the solar panels is directed towards the sun. Independent programmed turns about the first and second axes of the artificial earth satellite are performed in intervals of uncertainty of orientation of the artificial earth satellite on shadow orbits. In different versions of said turns, after the first turn, the artificial earth satellite is held in an intermediate position and normal orientation of the artificial earth satellite is then restored. This improves the accuracy of predicting movement of the artificial earth satellite on shadow orbits and accuracy of measuring the range to the artificial earth satellite.

EFFECT: high accuracy of determining navigation-time data on navigation artificial earth satellites by consumers.

4 cl, 12 dwg

FIELD: aircraft engineering.

SUBSTANCE: invention relates to in-flight control over spacecraft equipped with heat radiator and solar battery. Proposed process comprises spacecraft flight in orbit around the planet with solar battery turn to position corresponding to normal to solar battery working surface directed to the Sun. Spacecraft orbital orientation is constructed whereat solar battery spinning plate is parallel with spacecraft orbit plane while solar battery is located on the Sun side relative to orbit plane. Spacecraft orbit altitude and angle between direction to the Sun and spacecraft orbit plane are defined. Magnitude of said angle (β*) is defined whereat duration of turn shadow section equals the necessary time of radiator heat release in said turn. Orbit turns are defined wherein current magnitude of said angle is larger than β*. In said turns, solar battery is turned around crosswise and lengthwise rotation axes unless shadowing of solar battery radiator. Note here that minimum departure of orientation of solar battery working surface to the Sun. Spacecraft orbital flight is conducted in near-circle orbit at altitude not exceeding a definite design value.

EFFECT: higher efficiency of radiator with solar battery shadowed at whatever position of spacecraft on orbit turn.

3 dwg

FIELD: transport.

SUBSTANCE: invention relates to control over spacecraft, particularly, to holding of geosynchronous spacecraft in preset are of stay and collocation with the other geostationary spacecraft. Proposed method comprises determination and correction of initial inclinations and longitude of injection orbit ascending node with allowance for epoch of spacecraft placing in orbit and term of its active existence. Note here that the time of beginning of operation in geostationary orbit when spacecraft orbit inclination reaches maximum permissible value iper. area. The latter corresponds to permissible reach in latitude at the boundary of nominal spacecraft stay area in altitude. Stable and minimum eccentricity magnitudes are defined. Eccentricity vector is corrected so that it equals the nominal value for spacecraft collocation and spacecraft orbit apse line is aligned with that of nodes. Spacecraft active collocation is executed at changing the inclination from 0 to iper without interaction with adjacent spacecraft control centres. At inclination larger than iper , eccentricity is increased to minimum with setting of Laplace vector in direction from the Sun. Note here that eccentricity vector is not corrected unless the end of spacecraft active existence term termination. At inclinations larger than iper, eccentricity vector equals modulo and is spaced apart relative to eccentricity vectors of the other spacecraft.

EFFECT: decreased power consumption for stay area and collocation of geostationary spacecraft.

9 dwg

FIELD: instrumentation.

SUBSTANCE: invention relates to control of movement of space vehicle (SV). According to the proposed method, thrusts of correction engines (CE) (control accelerations) are determined as per total changes in a period of SV revolution from correction to correction. The latter are performed with one and the same CE and a thrust level of those CE is estimated. For reliable understanding of thrusts of a pair of mutually opposite located CE, series control activations of the same pair are performed from time to time with equal pulses. Discrepancy as to total pulse of thrust is entered in equal parts with an opposite sign to the implemented pulses. As a result, reliable thrust levels of CE in operation are obtained.

EFFECT: reduction of costs and improvement of CE thrust determination accuracy as per the data of trajectory measurements, as well as improvement of SV orbit correction accuracy.

FIELD: transport.

SUBSTANCE: invention relates to spacecraft (SC) motion control using solar radiation pressure forces distributed over SC working zones. The latter are formed as flat parallel optically transparent droplet flows. Distance between droplets of R radius in each flow in its lengthwise direction (Sx) and frontal-lateral direction (Sy) is divisible by 2R. Number of flows is n=(Sx/2R)1. By mutual bias of flows in direction of their motion for 2R distance droplet mist flows are generated in number of m=(Sy/2R)1. Each of the mentioned flows is biased relative to previous flow for 2R distance in frontal-lateral direction. Thus opacity in frontal-lateral direction and transparency in direction of plane perpendicular to a flow is created. Unit distributed light pressure force is regulated by changing radius and number of droplets coming to point of it application in unit time. Total action value is regulated by changing number of droplet jets.

EFFECT: increase of efficiency of light pressure distributed external forces usage by means of decreasing their disturbing effect on relative SC motion.

3 dwg, 1 tbl

FIELD: chemistry.

SUBSTANCE: invention can be used to move space debris from working orbits to recycling orbits. The method includes taking towing spacecraft and a self-contained docking module into the region of an orbit from which space debris is to be cleared. The sequence of removing space debris is selected by comparing a criterion, for example the probability of the space debris colliding with other space objects, for each space debris. Compensation for accumulated errors of motion parameters of the towing spacecraft during previous manoeuvres, as well as the pointing system is distributed between correcting pulses of the towing spacecraft at the long-range guidance step and of the self-contained docking module at the self-guidance portion.

EFFECT: high efficiency of removing space debris from working orbits.

FIELD: space engineering; designing spacecraft motion control systems.

SUBSTANCE: proposed method is performed by information of orientation unit to Sun by introducing the orbit parameters into on-board computer followed by calculating the Sun position in observation field of orientation unit for each point of orbit for orientation of axes in orbital coordinate system; search angular velocity is set for spacecraft to ensure capture of Sun by observation field of orientation unit, after which angular velocity is decreased to zero ensuring position of Sun in observation field of orientation unit. Then spacecraft is turned in such way that Sun should move to required initial point; turning the spacecraft to preset points is continued for each orbital point.

EFFECT: reduced mass; simplified construction of spacecraft due to reduced number of instruments and units; extended field of application.

3 dwg

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