Method of controlling orbiting spacecraft

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

 

The invention relates to the field of space technology and can be used to control the motion of the SPACECRAFT).

The SPACECRAFT is equipped with solar batteries (SB), which produce electricity for the functioning of the AC. When you implement flight operations KA activated side apparatus, the elements of which when the work is heated. The heat is used for temperature control of SPACECRAFT, and its excess is discharged into the surrounding SPACECRAFT space through the radiators-emitters. The amount of heat the most effective on shadow earth orbit, during which the entire surface of the radiator-teplocluchenka not exposed to direct solar radiation, and less effective on the sunlit portions of the orbit, when the reset heat comes mainly from those portions of the radiator-teplocluchenka that shaded components KA (Tabor O. N., Kadaner J. C. Questions of heat transfer in space. M.: Higher school, 1972).

Known way to control the orbital SPACECRAFT (Eliseev A. S. Technology of space flight. M: mechanical engineering, 1983), including the reversal SAT in the working position of the Sun and the running of the orbital flight of a SPACECRAFT around a planet in which the reset heat heat sink-teploizolyatsii is in moments of finding KA in the shadow of the planet, and in moments vetovo part of the coil, when the current orientation of the SPACECRAFT design SPACECRAFT obscures the radiator teploizolyatsii from direct sunlight. In this way the reset heat heat sink-teploizolyatsii is due to the natural cooling radiator-teplocluchenka in the moments of its shades of the planet or structures CA. The disadvantage of this method is that it is, in General, does not guarantee the presence of the luminous part of the orbit shading radiator-teplocluchenka design of the SPACECRAFT. For example, when the SPACECRAFT into solar orbit (when the shadow on the orbit the orbit is missing) no shading radiator-teplocluchenka design KA means no shading radiator-teplocluchenka throughout the circuit, which significantly reduces the effectiveness of the radiator-teploizolyatsii its functions.

Known way to control the orbital SPACECRAFT (Malozemov centuries heating mode of the spacecraft. M.: Mashinostroenie, 1980) adopted for the prototype, including the execution of orbital flight SPACECRAFT around the planet, turn SA into position on the Sun and run turn the AC up to Shader radiator-teplocluchenka design of the SPACECRAFT. This method is guaranteed reset heat heat sink-teploizolyatsii due to natural cooling radiator-teplocluchenka at moments of shading design is the Ktsia KA.

Prototype method has a major drawback - to create conditions for natural cooling radiator-teplocluchenka due to shading design KA in this way it is necessary to continuously perform the above special turn of SPACECRAFT that, on the one hand, requires additional energy costs for its implementation, and on the other hand, the implementation of the above special turn of SPACECRAFT in the General case can be contrasted with the construction of the desired target orientation KA-the orientation, which must be KA to solve his targets. Thus, in the process of decision targets KA, which is accompanied by the construction of the desired target orientation of the SPACECRAFT, in the General case, not created conditions for natural cooling radiator-teplocluchenka due to its shading, which impairs the efficiency of the radiator-teplocluchenka.

Task to be solved by the present invention is directed, is the efficiency of the radiator-teplocluchenka installed on the SPACECRAFT, with moveable SAT.

Technical result achieved in the implementation of the present invention is the creation of additional conditions for natural cooling radiator-teplocluchenka due to shading his mobile SA is A.

The technical result is achieved in that in the method of controlling an orbiting SPACECRAFT, including the execution of orbital flight SC posted on the radiator-teploizolyatsii in orbit around the planet and turn SB installed with two degrees of freedom on the SPACECRAFT, in its working position, corresponding to the combination of the normal to the working surface SB with direction to the Sun, additionally build the orbital orientation of the SPACECRAFT during which the rotation plane SAT parallel to the plane of the orbit of the SPACECRAFT and SB is located relative to the plane of the orbit of the Sun, determine the maximum value of the angle between the velocity vector of the SPACECRAFT and perpendicular to the transverse axis of rotation of the SAT, passing through the surface of the radiator-teplocluchenka, determine the angle between the direction of the Sun and the plane of the orbit, determine the height of the orbit, the altitude of the orbit and a particular value of the angle between the direction of the Sun and the plane of the orbit determines the coils of the orbit on which the duration of the illuminated side of the coil exceeds the difference between the orbital period of the AC and the necessary duration of the reset time of the heat radiator-teploizolyatsii on the circuit, and videopreteen orbits the orbit by passing KA illuminated side of the coil do turn SAT around a transverse axis of rotation SAT before crossing the nternet, passing through the sunward surface area of the radiator-teplocluchenka and aimed at the Sun to SAT and turn SAT around the longitudinal axis of rotation of the SAT until the angle between the normal to the working surface SB and the direction to the Sun the minimum value, the above turns SB perform during the total duration of the k·P-T within the time interval, the start and end of which are calculated, respectively, by the formulas:

t1=tS+P2π(η-γ),t2=tS+P2π(η+π+γ),

whereγ=arcsinL2-D2-E2ctg2β2DEctgβ,

k - coefficient characterizing the required duration of the reset time heat radiator teploscat the LEM at each point is equal to the ratio of necessary duration of the reset time of heat on the circuit for the duration of the round,

P is the orbital period of the SPACECRAFT,

T - the duration of the shadow part of the round,

ts- time passing KA sunflower point round,

η is the maximum value of the angle between the velocity vector of the SPACECRAFT and perpendicular to the transverse axis of rotation SB passing through the surface of the radiator-teplocluchenka, the positive direction of the reference angle of the velocity vector in the direction of the radius vector KA,

L - length SAT along the longitudinal axis of its rotation,

D is the distance from the transverse axis of rotation of the SAT to the most remote from the axis point of the surface of the radiator-teplocluchenka,

E - the distance from the plane of rotation of the SAT to the most remote from the plane surfaces of the radiator-teplocluchenka,

β is the angle between the direction of the Sun and the plane of the orbit.

The essence of the invention is illustrated in Fig.1÷5, which shows: Fig.1 - scheme of the mutual position of SB and radiator-teplocluchenka relative to the direction of the Sun, illustrating a view in the plane of the orbit, Fig.2, 3, 4 - scheme of the mutual position of SB and radiator-teplocluchenka relative to the direction of the Sun, illustrating end view of the orbital plane of Fig.5 is a diagram illustrating the definition of the angle between the direction of the Sun and the plane of the orbit, in which the duration of the shadow part of the loop orbit is not the required time duration of the reset heat heat sink-teploizolyatsii on the circuit.

In Fig.1÷5 introduced the notation:

1 - orbit SATELLITES;

2 - longitudinal axis of rotation of the SAT;

3 is transverse the axis of rotation of the SAT;

4 - radiator-teploizolyatsii;

5 is perpendicular to the transverse axis of rotation SB passing through the sunward surface area of the radiator-teplocluchenka,

6 is a plane of rotation of the SAT;

S is the direction vector of the Sun;

Spthe projection direction of the Sun on the plane of the orbit;

About the center of the planet;

β is the angle between the direction of the Sun and the plane of the orbit;

As- the position of the SPACECRAFT in sunflower point orbits;

And1And2position the SPACECRAFT in moments of t1, t2;

AC, A1C1, A2C2, A3C3- the distance from the transverse axis of rotation of the SAT to the most remote from the axis point sunward surface area of the radiator-teplocluchenka;

AB, A1B1And2In2And3In3.- the distance from the plane of rotation of the SAT to the most remote from the plane of the point sunward surface area of the radiator-teplocluchenka;

VM1M1In2M2In3M3- cut longitudinal axis of rotation of the SAT concluded between the beginning and the end of the SAT;

F1F2position the SPACECRAFT at the beginning and end of the shadow area of a coil;

Fs- state the SPACECRAFT at the time of the middle of the shadow area of a coil;

The Z - surface of the planet.

Explain proposed mode of action.

Accept that KA SA established with two degrees of freedom: the panel SAT rotated around a longitudinal axis of rotation and SAT around a transverse axis of rotation of the SAT. And turn SAT around a transverse axis of rotation of the SAT is to rotate the longitudinal axis of rotation SAT around a transverse axis of rotation of the SAT. Thus consider a system of position control of the security Council, in which the transverse axis of rotation SAT directly passes through the longitudinal axis of rotation and SAT perpendicular to it.

Accept that SB made opaque: SAT delay coming to them, the flow of solar energy and can shade yourself from the Sun the external surface of the SPACECRAFT.

Accept that SB have an elongated rectangular shape, and the length of the SAT measured along the longitudinal axis of rotation of the SAT. The width SB is not less than the value of the linear dimension of the surface of the radiator-teplocluchenka.

In the proposed method performs orbiting SPACECRAFT to host radiator-teploizolyatsii around the planet in near-circular orbit.

Perform the reversal SAT in working position, corresponding to the combination of the normal to the working surface SB with direction to the Sun. In this orientation SB provides the maximum advent of electricity is.

Build the orbital orientation of the SPACECRAFT during which the rotation plane SAT parallel to the plane of the orbit of the SPACECRAFT and is located relative to the plane of the orbit from the Sun. This corresponds to a transverse axis of rotation SAT perpendicular to the plane of the satellite orbit, and SB is located relative to the plane of the orbit of the Sun.

After building the orbital orientation perform its maintenance and determine the maximum value of the angle η between the velocity vector of the SPACECRAFT and perpendicular to the transverse axis of rotation SB passing through the surface of the radiator-teplocluchenka, the positive direction of the reference angle η of the velocity vector in the direction of the radius vector of the SPACECRAFT.

Determine the angle between the direction of the Sun and the plane of the orbit of β (we assume that is always β≥0).

Determine the height of the orbit N.

For certain values of the angle between the direction of the Sun and the plane of the orbit of β and the orbit altitude H determine (select) the coils of the orbit on which the duration of the illuminated side of the coil exceeds the difference between the orbital period of the AC and the necessary duration of the reset time of the heat radiator-teploizolyatsii on the circuit (data coils are not able to ensure the required duration of natural cooling radiator-teplocluchenka in the shadow of the planet).

Definition (selection of such coils is for example, as follows.

When the current altitude of the orbit of the SPACECRAFT in the altitude range [H1H2]:

H1HH2,(1)

H1=Rsin(arccos(cos(kπ)cosβmax))-R,(2)

H2=Rsin(kπ)-R,(3)

βmax=min{|i|+ε,π2},the (4)

where k is the coefficient characterizing the required duration of the reset time of the heat radiator-teploizolyatsii at each point is equal to the ratio of necessary duration of the reset time of heat on the circuit for the duration of the round,

R is the radius of the planet,

βmax_the maximum value that can take the angle between the direction of the Sun and the plane of the orbit,

i is the inclination angle of the orbit KA,

ε is the inclination angle of the Ecliptic (ε ~ 23°26'),

select only the coils in which the current value of the angle between the direction of the Sun and the plane of the orbit more β values β*, in which the duration of the shadow part of the loop orbit is equal to the required duration of the reset time of the heat radiator-teploizolyatsii on stage:

β>β*,(5)

β*=arccoscosθcosλ,(6)

sinθ=RR+H.(7)

λ=T22πP,(8)

T=kP,(10)

where θ is the angular polarstar visible with the AC drive of the planet,

λ is the angular polarstar shadow part of the round orbit, measured from the center of the planet,

P is the orbital period of the SPACECRAFT,

T - the duration of the shadow part of the loop.

Condition (5) corresponds to the fact that the length of the shadow part of these orbits the orbit is less than the required duration of the reset time of the heat radiator-teploizolyatsii on the circuit. If the condition (5) shadow on the circuit or missing entirely, or its duration is less than the required time duration of discharge of heat, R is Diatron-teploizolyatsii on the circuit. If the condition (5) is not performed (when (β≤β*), then at this stage the length of the shadow part of the coil is greater than or equal to the required duration of the reset time of the heat radiator-teploizolyatsii on the circuit.

At the current height of the orbit is greater than the height H2:

H2<H,(10)

select all the coils, because when running (10) shadow on all the turns or missing, or its duration is less than the required duration of the reset time of the heat radiator-teploizolyatsii on the circuit.

At the current height of the orbit is less than the height H1(if N<N1), at any stage there is a shadow part and its duration is always more necessary duration of the reset time of the heat radiator-teploizolyatsii on the circuit.

On the selected coils of the orbit by passing KA illuminated side of the coil do turn SAT around a transverse axis of rotation of the SAT to the intersection of the line passing through the sunward surface area of the radiator-teplocluchenka and aimed at the Sun to SAT and turn SAT around the longitudinal axis of rotation of the SAT until the angle between the normal to the working surface SB and the direction is by giving the Sun the minimum value. In this orientation, the panel SAT shadows facing the Sun, the surface area of the radiator-teplocluchenka. In this case, rotation SAT around the longitudinal axis of rotation of the SAT until the angle between the normal to the working surface SB and the direction to the Sun the minimum value is used both to increase the area, shaded SAT, and to maximize the generation of electricity generation of electricity depends on the angle of incidence of solar radiation on the surface of the SAT). Data turns SB perform during the total duration Δ:

Δ=kP-T,(11)

and within time interval [t1, t2], the moments of the beginning and end of which are calculated by the formulas:

t1=tS+P2π(η-γ),(12)

t2=tS +P2π(η+π+γ),(13)

whereγ=arcsinL2-D2-E2ctg2β2DEctgβ,(14)

ts- time passing KA sunflower point round,

L - length SAT along the longitudinal axis of rotation of the SAT, measured from the transverse axis of rotation SAT,

D is the distance from the transverse axis of rotation of the SAT to the most remote from the axis point of the surface of the radiator-teplocluchenka,

E - the distance from the plane of rotation of the SAT to the most remote from the plane surfaces of the radiator-teplocluchenka,

β is the angle between the direction of the Sun and the plane of the orbit.

The distance E can be counted along the transverse axis of rotation of the SAT. In this case, this distance can be defined as the distance between the longitudinal axis of the treatment of the security Council and perpendicular to the transverse axis of rotation of the SAT, passing through the most remote from the plane of rotation of the SAT point of the surface of the radiator-teplocluchenka.

In the absence of a shadow on the circuit side in equation (11) the duration of the shadow part of the loop is equal to zero (T=0).

The interval [t1, t2] obtained in such a way that at any moment this time interval SB can be rotated to a position in which a straight line directed from the sunward surface area of the radiator-teplocluchenka towards the Sun, crosses the panel SB - i.e., the SAT will overshadow the surface area of the radiator-teplocluchenka. Outside the time interval [t1, t2] length SB L, in General, not enough to the surface of the radiator-teplocluchenka could be shaded SAT.

Thus, the result of performing the above steps, the total on the circuit, taking into account the duration of the shadow part of round T, the radiator teploizolyatsii will be shaded during the time Δ+T=k·R, which is the required duration of the reset time of heat on the circuit.

Explain the formulas used.

Ratio(12), (13), (14) obtained from the relation:

L=D+Etg(π2-β).the (15)

Explain the relation (15). To do this, let us consider the rotated position of the longitudinal axis of rotation of the SAT, which line passes through the point sunward surface area of the radiator-teplocluchenka and aimed at the Sun, intersects the longitudinal axis of rotation of the SAT.

Equation (15) determines in turn orbits a1And2such that when the angle between the direction to the Sun and the plane of the orbit, is equal to the value of β, and in such rotated position of the longitudinal axis of rotation of the SB line passing through the point sunward surface area of the radiator-teplocluchenka and aimed at the Sun, passes through the end of SAT. This means that the length SB L corresponds to the length which is necessary and sufficient for shading the surface of the radiator-teplocluchenka panel SAT at points round And1And2(Fig.2).

Geometrically the position of data points And1And2described by the following angles: a1And2distance from point Asin the course of the orbital motion of the SPACECRAFT at the corners, respectively, η γ, η+π+γ (Fig.1). Where there is a formula(12), (13).

At all points of a coil located along the orbital flight of KA between points a1And2(for example, at point a in Fig.1, figs.3), for shading the surface radiation is ora-teplocluchenka panel SAT quite smaller length SAT, than the length of the security Council, are necessary and sufficient to shade the surface of the radiator-teplocluchenka at points a1And2.

Thus, at all points of a coil located along the orbital flight of KA between points between points a1And2defined by the relation (15), length SB L will be enough to shade radiator-teplocluchenka panel SAT.

On the other hand, at the points of a coil located along the orbital flight of KA between points a2And1(for example, at point a3in Fig.1, figs.4), for shading the surface of the radiator-teplocluchenka panel SAT requires a large length of the SAT than the length of the security Council, are necessary and sufficient to shade the surface of the radiator-teplocluchenka at points a1And2- i.e. length SB L, in General, not enough to the surface of the radiator-teplocluchenka could be shaded SAT between points a2-A3-A1.

Note that in Fig.1÷4 presents illustrations, in which the distance from the transverse axis of rotation of the SAT to the most remote from the axis point of the radiator-teplocluchenka D is equal to AC, A1C1And2With2And3With3and the distance from the plane of rotation of the SAT to the most remote from the plane of the point of the radiator-teplocluchenka E is the AB1In1And2B2And3 In3. In the General case D≥AC, A1C And2With2And3With3and E≥AB, A1B1And2In2And3In3.

Ratio (5)÷(9) are illustrated by the scheme shown in Fig.5, the ratio (9) corresponds to the equality of the length of the shadow part of the revolution necessary duration of the reset time of the heat radiator-teploizolyatsii on the circuit.

The relation (2), (3) follow from (6)÷(9), respectively,

β*=βmaxand β*=0.

Usually on KA post some SS and some radiators-emitters. For example, SA can be installed in pairs, with each pair of the longitudinal axis of rotation SAT in opposite directions. A few (for example, not less than four radiators-emitters, each of which has a flat shape, can be placed on different sides of the external surface of the SPACECRAFT. In this case, the validity of the proposed method is applied to all sorts of different combinations and SAT radiators-emitters.

Describe the technical effect of the invention.

The proposed solution improves the efficiency of the functioning of the radiator-teplocluchenka placed on with moveable SB KA, by creating additional conditions for natural cooling radiator-teplocluchenka by Satan is that its SB KA at any height near-circular orbit.

The achievement of the technical result is achieved through:

- complete construction of the proposed orbital orientation of the SPACECRAFT in which the plane of rotation SB focus suggested by the way,

the proposed definition of angles and height of the orbit on which the proposed method determines the coils of the orbit, which violates the condition for achieving the desired duration of natural cooling radiator-teplocluchenka in the shadow of the planet, and the proposed manner is determined by the time interval within which are proposed twists SAT,

- perform the proposed orbits the orbit of the proposed SB turns over the proposed duration of time and within the proposed time interval.

Consequently, the proposed action and the proposed terms of their implementation, it is possible to implement shading radiator-teplocluchenka rotating security Council, which creates conditions for natural cooling radiator-teplocluchenka (in moments of lack of lighting-radiator-teplocluchenka the Sun). Evaluated the effectiveness of the present invention on the international space station (ISS) has shown that its use will improve the efficiency of the radiators-emitters placed on the modules Ross is isogo segment of the ISS.

An industrial implementation of essential features that characterize the invention, is not difficult and can be performed by known technologies.

The method of controlling the orbital spacecraft, including the execution of orbital flight of a spacecraft placed on it by the radiator-teploizolyatsii in orbit around a planet and the reversal of the solar panels mounted with two degrees of freedom on the spacecraft, in its working position, corresponding to the combination of the normal to the working surface of the solar battery with the direction to the Sun, characterized in that build orbital orientation of the spacecraft in which the plane of rotation of the solar panels parallel to the plane of the orbit of the spacecraft and the solar panel is located relative to the plane of the orbit of the Sun, determine the maximum value of the angle between the velocity vector of the spacecraft and perpendicular to the transverse axis of rotation of the solar battery, passing through the surface of the radiator-teplocluchenka, determine the angle between the direction of the sun and the orbit plane of the spacecraft, determine the height of the orbit of the spacecraft, at a certain altitude and a particular value of the angle between the direction of the Sun and the plane of the orbit to the body of low unit determines the coils of the orbit, where the duration of the illuminated side of the coil exceeds the difference between the orbital period of the spacecraft and the necessary duration of the reset time of the heat radiator-teploizolyatsii on the circuit, and videopreteen orbits the orbit by passing spacecraft illuminated side of the coil perform the rotation of the solar battery around a transverse axis of rotation of the solar battery to the intersection of the line passing through the sunward surface area of the radiator-teplocluchenka and aimed at the Sun, with solar battery and rotating solar panels around the longitudinal axis of rotation of the solar panels until the angle between the normal to the working surface of the solar battery and the direction to the Sun the minimum value, the above turns solar panels perform during the total duration of the k·P-T within the time interval, the start and end of which are calculated, respectively, by the formulas:

where
k - coefficient characterizing the required duration of the reset time of the heat radiator-teploizolyatsii at each point is equal to the ratio of necessary duration of the reset time of heat on the circuit for the duration of the round,
P is the orbital period of the spacecraft,
T - glutelin is here the shadow part of the revolution,
ts- the time passing by the spacecraft seed point round,
η is the maximum value of the angle between the velocity vector of the spacecraft and perpendicular to the transverse axis of rotation of the solar battery, passing through the surface of the radiator-teplocluchenka,
L is the length of the solar battery along the longitudinal axis of its rotation,
D is the distance from the transverse axis of rotation of the solar panels to the most remote from the axis point of the surface of the radiator-teplocluchenka,
E - the distance from the plane of rotation of the solar panels to the most remote from the plane surfaces of the radiator-teplocluchenka,
β is the angle between the direction of the sun and the orbit plane of the spacecraft.



 

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2 dwg

FIELD: transport.

SUBSTANCE: invention relates to thermal control systems of, mainly, long-term operation telecommunication satellites. TCS circuit with two-phase heat carrier (ammonia) comprises fluid pump, manifolds of metre and radiator panels and accumulator. Accumulator case has zones of vapours of said heat carrier and heat carrier liquid phase. The latter zone is connected with circuit directed to fluid pump intake. Said circuit is connected via connection pipe via controlled throttle with accumulator case. Said throttle serves to control heat carrier pressure and temperature in accumulator case. About 10% of liquid hear carrier flow is fed therefrom into the case central zone. To separate liquid phase form undissolved gas bubbles (if any) at the outlet of said pipe shaped to half-loop of somewhat radius. Cross-section of said section features rectangular shape with larger side located in the plane perpendicular to heat carrier flow.

EFFECT: reduced heat carrier losses, hence smaller onboard store of heat carrier.

2 dwg

FIELD: transport.

SUBSTANCE: invention relates to thermal control systems of, mainly, telecommunication satellites. Proposed system comprises closed heat carrier circulation circuit. Said circuit is composed of fluid circuits of electrically driven pump unit, manifolds of radiator panels, metre panels and connection pipes. Length of said fluid circulation circuit has two parallel different-length branches. Smaller-length branch features smaller ID. Total length of said sections is calculated by definite mathematical formula.

EFFECT: decreased sizes, higher reliability.

3 dwg

FIELD: transport.

SUBSTANCE: invention relates to thermal control systems of, mainly, telecommunication satellites. Thermal control system comprises closed heat carrier circulation circuit. The latter includes electrically driven pump unit, pressure accumulator, manifolds of control panels and those of radiators. Said elements are interconnected by lengths of connection pipelines with inlet and outlet flow sections complying with those of said elements. Portion of lengths of said connection pipelines feature identical equivalent ID smaller than diameters of the other parts and total length that satisfies the definite relationship.

EFFECT: decreased sizes, higher reliability.

1 dwg

FIELD: transport.

SUBSTANCE: set of invention relates to spacecraft thermal control systems and can be used in spacecraft preparation for flights as well as in other fields. In compliance with this method, prior to filling evacuated hydraulic line with working fluid hydropneumatic compensator (HPC) gas chamber maximum volume (VHPSmeas) measured. Said chamber is filled with gas at pressure higher than that of displacement gas above working fluid surface in refueller tank. Said line filled with working fluid, its weight-average temperature is measured. Initial gas pressure in HPC has chamber is set defined by measured pressure (VHPSmeas) and design pressure (VHPSw), design working pressure (pW) in said gas chamber and working fluid column height from the point of HPC fluid chamber connection to said line to top points of said line. Then, HPC fluid chamber is filled with working fluid at control over current gas precure in HPC gas chamber. HPC gas chamber pressure reaching pW filling it is terminated. Proposed device comprises refueller with filling and draining tanks, vacuum unit, gas pressure source, appropriate filling draining and controlling equipment with appropriate valve and accessories. Inventions allow exclude jobs of gaged discharge and associated processes of neutralisation and recover of discharged working fluid.

EFFECT: higher safety and reliability of HPC filling.

2 cl, 1 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 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: electricity.

SUBSTANCE: result is achieved by increasing strength of connection of shunting diodes and solar elements, increased repeatability of the process of manufacturing of the solar battery of spacecrafts due to optimisation of the technology of manufacturing of shunting diodes and solar elements of the solar battery, and also switching buses that connect the solar elements and shunting diodes, which are made as multi-layer. The solar battery for small-size spacecrafts comprises the following: panels with modules with solar elements (SE) adhered to them, a shunting diode; switching buses that connect the face and reverse sides of the shunting diodes with solar elements, at the same time the shunting diode is installed in the cut in the corner of the solar element, at the same time switching buses are made as multi-layer, made of molybdenum foil, at two sides of which there are serial layers of vanadium or titanium, a layer of nickel and a layer of silver, accordingly.

EFFECT: increased resistance of solar batteries to thermal shocks, to impact of mechanical and thermomechanical loads, increased manufacturability of design, extended active life of a solar battery of spacecrafts, increased functional capabilities due to expansion of temperature range of functioning and optimisation of design of a solar battery, simplification of a switching system.

7 cl, 4 dwg, 3 tbl

FIELD: transport.

SUBSTANCE: invention relates to spacecraft electric power supply with the help of solar batteries. Proposed method comprises definition of preset angle of solar battery orientation to the Sun by measured angular position of normal to battery working surface and computation of design angle relative thereto. Solar battery is spinned in direction of decrease in mismatch between preset and design angles. Solar battery acceleration angle (αAC) and deceleration angle (αDEC) are defined. Design angle is corrected when angle transducer readings vary by discrete sector of solar battery turn. Threshold of operation and drop-away (αT) and (αD) are set to terminate battery spinning if mismatch between preset angle and current angle increases but not over αT. Solar battery angular velocity is set or the order and larger than maximum angular velocity of spacecraft revolution around the Earth while discrete sector magnitude is set to smaller than αT. Solar battery working angle (αW) is set provided that αT < αW < (α"ГОР" - 2·(αAC + αDEC)). Angular position of closest beam of angle αW is assigned to preset angle if direction to the Sun in projection to the plane of spinning of said normal is located outside of αW. Is angular position of said normal is outside αW to vary in direction of increase of angle relative to nearest beam of angle αW, failure warning is generated to terminate control over solar battery.

EFFECT: ruled out jamming and breakage of solar battery panels or spacecraft onboard hardware at turns from 90° to 180°.

3 dwg

FIELD: transport.

SUBSTANCE: invention relates to spacecraft electric power supply with the help of solar batteries. Proposed method comprises definition of preset and current angles of solar battery orientation and solar battery angular velocity (ωSB). Design angle is computed to assign measured angle magnitude thereto and memorised prior to start of control over solar battery. Solar battery is spinned in direction of decrease in mismatch between preset and design angles. Defined are angles of solar battery acceleration and deceleration (tAC, αDEC) and threshold (tTHR, αTHR) and maximum tolerable angle of its deflection (αMAX) proceeding from minimum tolerable currents of solar battery. Said angles are used to set operation threshold (αS). The latter exceeded, said mismatch is generated. The latter is not taken into account if lower than drop-away threshold (αDROP). The latter reached, solar battery spinning is terminated. Solar battery design angle is corrected with the limits of one discrete sector of solar battery spinning circle. Discrete sector magnitude depends of angles αAC, αTHR and αS. Depending upon αS and ωSB threshold of the interval of control over continuous variation of data on solar battery angular position is set. Count of said interval is made if current measured angle differs from memorised one by more than one discrete sector and is terminated otherwise. Threshold of the time of control over solar battery spinning is set depending upon tAC, tTHR, αMAX, ωSB and discrete sector magnitude. This time is counted at zero time of control over continuity is sign of mismatch between measured and memorised angles dose not satisfy the solar battery preset direction of spinning. Otherwise, count is terminated to zero the time of control of spinning direction. Note here that when measured angle varies by one discrete sector, angular angle of boundary between discrete sectors is taken to be design angle to assign new measured angle to memorised angle. In case the time of control over continuity or that over spinning direction exceeds its threshold, failure signal is generated to terminate control over solar battery.

EFFECT: higher survivability and efficiency.

3 dwg

FIELD: transport.

SUBSTANCE: invention relates to spacecraft electric power supply with the help of solar batteries. Proposed method consists in definition of preset angle of solar battery, measurement of its current angle and computation of design angle by angular velocity and spinning time. Solar battery acceleration angle (αAC) and deceleration angle (αDEC) are defined. Solar battery is spinned to threshold of drop-away (αDROP ≈ (αDEC) when mismatch between said preset and design angles is terminated. Before start of the control, preset angle is memorised to take initial preset angle as valid actual current angle. Mismatch threshold (αTHR) of said angles proceeding from angles αAC and αDEC, as well as minimum tolerable currents of solar battery. Angle transducer circle is divided into equal discrete sectors by magnitude a given the condition αAC + αDEC < σ < αTHR. Discrete sector bisectors are taken to be measured magnitudes. Period of valid current angle definition is set to the order and exceeding maximum duration of transducer data fault and smaller than minimum interval of faults train. Said interval is divided to four equal interval while analysis of measured and memorised magnitudes in said intervals are reset to generate validity signal. In the latter case, solar battery is spinned to mismatch between design and preset angles αDROP to set new preset angle magnitude.

EFFECT: higher survivability and efficiency.

4 dwg

Solar battery strut // 2499751

FIELD: aircraft engineering.

SUBSTANCE: solar battery strut comprises two-link mechanism with common axle supporting torsion spring with cocking devices. One link is arranged at solar battery frame while another one is mounted at craft body. Spring-loaded rod to lock the link at end position is arranged at said link perpendicular to axis. Rocker is arranged at spring-loaded rod end to turn thereat. Antifriction bearings are rigidly secured at rod both ends to interact with cam taper grooves, said cams being rigidly mounted at the link opposite spring-loaded rod. Links of aforesaid mechanism have openings to link retainers threaded therein.

EFFECT: higher reliability, simplified installation of solar battery at craft body.

13 dwg

FIELD: transport.

SUBSTANCE: set of intentions relates to space power engineering and may be used for transmission of electric power in the form of laser radiation to Earth surface and for high-accuracy measurements in space, data transfer, etc. Proposed station comprises base module 1, system of mirrors 2, laser radiation summator 3 directed to system 2 and photo converter panel 4 arranged outside of module 1. Every panel 4 consists of two types: photoelectric panels 5 and independent photo emitting panels 6. The latter are connected in chain for self-opening and arranging in closed flat zigzag-like figure. Panels 5 are mounted at the start of chain 5, 6. Note here that the first panel is connected with base module 1. Said module 1 comprises the following systems: control system 8, cooling system 11 and supply system 12. Every panel 5 is connected with supply system 12. Every independent panel 6 is composed of a carcass with Fresnel lenses are carcass end with photo converters (not shown) aligned therewith and located there above. Carcass bottom part base accommodates power accumulators, control unit of panel 6 and fiber lasers with pumping units and laser radiation summator. Aforesaid photo converters of panel 6 are electrically connected via power accumulators with pumping and control units. Summators of independent panels 6 are connected to aforesaid summator via FO 30.

EFFECT: higher efficiency and reliability, longer life and expanded operating performances.

7 cl, 13 dwg

Solar cell battery // 2485026

FIELD: transport.

SUBSTANCE: invention relates to space engineering and may be used in designing external structures of spacecraft, primarily, solar cell batteries. Solar cell battery comprises frame, articulated top and bottom flaps with torsions fitted on hinge pins. Opposite ends of torsions support brackets wherein fitted are torsion resetting mechanisms. Said brackets are secured at torsions and set to initial position, in symmetry about torsion axis. Note here that one of said brackets is fitted on top flap while second bracket is mounted at bottom flap to allow resetting mechanism to twist torsions in one direction.

EFFECT: higher reliability, ruled out jam of flaps.

4 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

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