Method of determination of space station mass in flight

FIELD: space engineering.

SUBSTANCE: proposed method is based on continuous measurement of parameters of orbital space station motion. In addition to parameters of space station motion, its motion relative to center of mass is measured. Object of known shape and mass is separated from station in section of orbit when engines are disconnected and orientation of station is maintained with the aid of gyrodynes. Simultaneously parameters of motion of object separated from station in orbit and parameters of its rotary motion are measured. Mid-section of separated object is determined by measured parameters of rotary motion. Density of atmosphere is determined by this mid-section and measured parameters of motion of separated object. Then, mid-section area of orbital station is determined by measured parameters of its rotary motion and attitude of its movable parts. Then, mass of orbital station is determined from the respective mathematical expression by the mid-section area and atmosphere density and by measured parameters of motion. Proposed method does not require supply of calibrating and measuring pulses by power plant of cargo spacecraft coupled with station as distinguished from known methods.

EFFECT: saving of working medium; enhanced operational safety of cargo spacecraft.

 

The invention relates to space technology and can be used to determine the mass of a spacecraft, a space station and build spacecraft in orbital flight conditions.

The known method is similar to the determination of the mass of the space station in the process of changing the parameters of its orbit [1]. As space objects in the way are considered transport manned spacecraft "Soyuz" and the cargo ship "Progress" (ha)included in the manned orbital station "Salyut-6". The method includes the measurement of motion parameters of interest (apparent velocity, relative range and so on) when changing orbital parameters at sites far maneuvering, docking and joint flight. Each time the object or Assembly is applied to the impulse propulsion system (PS):

where i is the number of the remote turn - on,

ti- time of the i-th enable,

τi- work time control,

the thrust vector.

Impulse control tells the object impulse maneuver:

where m is the mass of an object;

the apparent velocity vector.

If not to take into account the direction issued by the control pulse, the mass Assembly of objects can be identified by expression the structure:

where- the value of the apparent velocity of the Assembly after the i-th remote turn-on,

R - module of thrust vector control.

The most significant disadvantage of the method-analogue is the uncertainty of the force of propulsion. Taken in the way the assumption that during operation of the propulsion characteristics constant in the interval (t1; t11), and at each i-th activated, they correspond to the nominal values, fails for the case of determining the value of the mass of a space object with an acceptable accuracy. So, if we take as control liquid rocket engine (LRE) cargo spacecraft "Progress-M", it is possible deviation during his work from the average nominal value of thrust is ±10%. Specified LRE is a two-component composition (oxidizer + fuel) and plug the fuel supply system. The value of the above deviation is determined for the installed series DN and depends primarily on the operating parameters of the particular engine and also features of its operation in flight conditions. There is an experimental dependence of the thrust rocket engine from its operating parameters [2], but this dependence allows to determine only approximately the average value of thrust that will not allow you to determine which of the mass of the Assembly space objects with accuracy, sufficient for subsequent ballistic and other scientific tasks.

As the prototype method of the present invention, the authors have chosen the method of determining the mass of the Assembly space objects in space flight [3]. The method is based on changing motion parameters station and continuous measurement of the position of the station in orbit and allows to determine the mass of an Assembly of space objects in the process of changing orbital parameters. For this measure the motion parameters of a joined object of known mass and build on these parameters and the impulse force applied to the Assembly by the mover of the joined object, determine the mass of the Assembly.

Compared with the method similar to the method prototype in the measurement process clarifies the actual thrust propulsion. Indeed, in the measurement process are the time interval in the operation of the thruster, in which the magnitude of the impulse force, determined according to the measured motion parameters of a joined object and the measured parameters of propulsion, have the least difference. For this time interval are fixed values of the operating parameters of the thruster. After Assembly to the regular changes of the orbital parameters of the impulse force implement when the mover is in the range defined above operating parameters. The mass of the Assembly is determined on the same interval p is the magnitude of the impulse force, determined by the measured values of the operating parameters of the engine with the value of the smallest differences of the pulse strength. Prototype method was used to determine the mass of the orbital complex "Mir". As space objects taken docked cargo ship "Progress-M and the orbital complex "Mir". Prototype method was also used to determine the mass of the international space station (ISS). As space objects taken docked cargo ship "Progress-M" and the ISS.

The main disadvantage of the prototype method is the necessity of issuing tiraumea pulse propulsion joined the space object on the area of Autonomous flight of this object, as well as the necessity of issuing a pulse that changes the parameters of the orbit Assembly of space objects.

For calibration when determining using the prototype method the mass of the ISS used a regular pulse is issued on the second day after the launch vehicle to lift it into orbit of the ISS. For calibration at the site of Autonomous flight thrust the engines of the cargo ship, which later are changing the parameters of the orbit of the station, it is necessary to avoid a reaction of the other engines. Meanwhile, there are two regular mode maintain the desired orientation of the cargo ship at the time made what I maneuver to raise its orbit. In the first case the motors, which receive a pulse operate in pulsed mode, i.e. when leaving the corners in the process of issuing impulse to create moments that offset the resulting moments of withdrawal, are turned on or off corresponding pairs of engines. In the second case, the engine, which is transmitted pulse, operate in a continuous mode, and the moments of turning of the vehicle are compensated on-off additional pairs of engines. The pulse mode of engine operation in the first case and the availability of additional sources of perturbation in the second does not allow one to adequately evaluate the traction motors, which adversely affects subsequent measurement of the mass of the station.

There is a third option for maneuver on a stand-alone site, which is the disabling circuit orientation of the cargo ship and the maneuver is carried out without maintaining a given orientation. This option maneuvering leads, firstly, to a significant skid ship at the corners. Cargo ship loses the desired orientation. The standard he should be able to twist on the Sun to provide the necessary energoprivod on its solar panels. To restore the orientation of the voltage in the power supply system of the vehicle may fall to a critical value, which will rubbed the vehicle and a breakdown of program execution flights. In case of successful outcome of the operation will require additional adjustment ("MOP-up") of the pulse, which leads to additional fuel consumption.

Thus, carrying out calibration of the engine thrust in the first two cases does not allow to obtain an estimate of the engine thrust with acceptable accuracy, while in the third case unsafe from the point of view of operation of the vehicle, and is also associated with excessive consumption of the working fluid.

In addition, as was installed on the data processing results on a stand-alone plot maneuvering, the value of the relative error random error in the determination of the impulse on the whole interval of the issuance of the pulse is almost an order of magnitude higher than for a single interval. Thus, selection of sites, the smallest differences according to the method prototype is not possible to estimate the mass of the space station with an error of less than 2.3% [3].

The problems to be solved by the proposed method are the savings of the working fluid and ensuring the safe operation of the cargo ship, as the proposed technical solution, there is no need of issuing tiraumea and measuring pulses.

The technical result is achieved in that in the method of determining the mass of the space station in-flight, based on continuous measurement of motion parameters of the station, unlike the t known it as a parameter of the motion stations are also measured motion parameters relative to the center of mass, produce separation from the station of an object of known shape and weight, in the area of the orbit on which the engine is off, make maintaining orientation with the help of gyrodines, simultaneously measure the motion parameters detachable object in its orbit and the parameters of its rotational movement, of the measured parameters of the rotational movement of the detachable object determines the midsection detachable object for a particular midsection detachable object and the measured parameters of its movement determines the density of the atmosphere, and then on the measured parameters of the rotational motion of the orbital station and angular position of its moving parts define the midsection orbital station for some particular midsection orbital and density of the atmosphere and the measured motion parameters orbital station determines its mass according to the expression:

where Cx- the drag coefficient of the station,

Sx- the cross-sectional area of the station (the area of the midsection),

- geocentric radius vector of the center of mass of the station, the point marked differentiation by time t,

- tension gravit the operating field of the Earth,

- speed station Greenwich coordinate system,

ρ - density incident on the station aerodynamic flow.

In using a method similar to and of the prototype method is the law of conservation of momentum. If the body is exerted by a force, then the amount of motion of a body is proportional to the current at him strength. Thus, if you try to clarify the magnitude acting on the body forces, then, using the law of conservation of momentum, we can determine the mass of this body. In the method prototype, a method of refinement of this force, in this case, the calibration of the force of propulsion. In this invention it is proposed to refuse the use of active propulsion. In flight to the space station the force of aerodynamic resistance. Continuous exposure of the drag force on the station leads to its inhibition. Using the law of conservation of momentum, we can determine the mass of the station. Thus, in the proposed method actually use natural braking force station in the atmosphere. This force is much smaller than the forces created by the operation of the thruster, but currently used measurement tools allow you to measure with high accuracy that the AET us to implement this method.

In the literature the problem of describing the motion of a space object is broken into two parts. To describe the motion of an object in orbit is considered the movement of the center of mass of the object, in other words, is considered the movement of an object as a material point. To describe the angular position of the object is considered its rotational movement or, in other words, its motion relative to the center of mass. In the proposed method, the measurement of rotational motion of objects allows you to define their Miceli, or square projections of the objects on the plane perpendicular to tangents to the trajectories of these objects. The measurement of motion parameters of the objects allows to assess the effect of the drag force on these objects and, ultimately, determine the mass of the station.

Consider the essence of the proposed method, taking as space objects the ISS and detachable satellite "Nano".

The orbital motion of the station is described by the equation:

where- the radius vector of the center of mass of the station, the point marked differentiation by time t,

the intensity of the Earth's gravitational field (gravitational effect on the moon and other Solar System bodies is neglected),

m is the mass of the station,

- the main vector of non-gravitational forces of nature acting on the station.

From non-gravitational forces acting on the station, we will consider only the force of aerodynamic resistance, which approximate the formula:

where- speed station Greenwich coordinate system,

b - the ballistic coefficient of the station,

ρ - density incident on the station aerodynamic flow.

The elements of the motion of the center of mass of the station are according to the trajectory measurements.

There is an expression to evaluate ballistic coefficient:

where Cx- the drag coefficient of the station,

Sx- the cross-sectional area of the station (the area of the midsection), which is calculated on the basis of available telemetry data on the rotational motion of the station.

Using the expression (7), we Express the mass of the station through the ballistic coefficient:

The ballistic coefficient of the station according to expressions (5), (6) can be represented as:

Substituting equation (10) in (9), we obtain:

In order of vopolop shall be the expression (11), it is necessary to clarify the actual density of the atmosphere. It takes detachable space object, which may be a satellite with known mass-inertia characteristics. The only restriction imposed on the satellite, is the need to control rotational motion of the satellite. The satellite can either be equipped with equipment to monitor its movement about the center of mass, or to be oriented in a special way without the help of law enforcement agencies job orientation, which is achieved, for example, by using a strong magnet. The satellite with a magnet, like a compass needle will always oriented along the lines of the magnetic field of the Earth.

As the ballistic coefficient of the satellite when it is known the midsection can easily be determined by the formula (7), specifying for satellite expressions (5) and (6) and expressing the density of the atmosphere through the ballistic coefficient of the satellite, we get the following expression to determine the density of the atmosphere:

where- the radius vector of the center of mass of the satellite,

- the speed of the satellite relative to the Greenwich coordinate system,

bs- the ballistic coefficient of the satellite.

The elements of the motion of the center of mass of the satellite also will the be according to the trajectory measurements.

The application of the described method involves the separation of the satellite. On Board the ISS occasionally delivered scientific, educational, commercial satellites, the removal of which the Board is carried out by the crew during the regular outputs into space. The separation of such satellites can be used to determine the mass of the station this way.

Currently available measurement tools allow you to define the parameters of motion of the centers of mass of the space objects (in this case, the parameters of motion of the centers of mass of the ISS and detachable satellite "Nano") with a high degree of accuracy. Can use data as the Russian system of ASN (Autonomous Satellite Navigation), and the American system GPS (Global Position System). The use of GPS allows you to determine the radius of the center of mass of the ISS with accuracy not worseand the velocity vector of the center of mass of the ISS with accuracy not worse. In the case of the satellite have the same error:,.

The error in the determination of the midsection of the object caused by the error in the determination of the rotational motion of the object, in other words, the error in determining the actual orientation of the object in the current time.

The ISS is equipped with sensor and orientation - sun and star sensors, and magnetometers, which taking into account errors introduced by errors in the telemetry channels, allow us to determine the orientation station with an accuracy of 20″. The maximum error in determining the midsection station, caused by a specified tolerance of orientation, will not be more than.

Satellite "Nano" is equipped with a strong magnet that allows it to be permanently oriented position according to the lines of the magnetic field of the Earth. In addition, the satellite is equipped with solar sensor, which allows to determine its actual orientation in space with an accuracy better than 1°. This error causes the error definition midsection satellite.

Assuming that these errors are random in nature, can determine the random error in the determination of the mass stations of the proposed method by the formula:

In the expression (13) is not taken into account the error in determining the mass of the satellite due to the fact that this error in comparison with the above error is very small. Thus, the random error in the determination of the mass of the ISS in the described manner will be:

Modern space station become the CE more complex and cumbersome. The Mir station was operated for many years, but its weight was about 120 tons. The international space station is operated only a few years, and her weight is about 180 tons. Even now design and evaluation of its weight give an uncertainty of about 2 tons, which is 1.11%. The error in determining the mass of the ISS proposed method is significantly less than the project uncertainties. Thus, the proposed method will allow to clarify the mass of the ISS.

Sources of information

1. Navigation obschenie flight of the orbital complex "Salyut-6" -"Soyuz" - "Progress". - M.: Nauka, 1985.

2. Senarav D.B, Dobrovolsky MV Liquid rocket engines. - M.: Oboronprom, 1957.

3. Y.P. Semenov and other means of determining the mass of the Assembly space objects in the process of changing orbital parameters. EN 2098326 C1 IPC 6: 64G 1/26, 10.13.97 prototype.

The method of determining the mass of the space station in-flight, based on continuous measurement of motion parameters of the station, characterized in that as the motion parameters of the station are also measured the parameters of its motion relative to the center of mass, produce separation from the station of an object of known shape and weight, in the area of the orbit on which the engine is off, maintain the orientation of the station with the help of gyrodines, simultaneously measure the parameters of movement of the detachable object in its orbit and the parameters of its rotational movement, on the measured parameters of the rotational movement of the detachable object determine its midsection, and in particular midsection and measured motion parameters detachable object determines the density of the atmosphere, and then on the measured parameters of the rotational motion of the orbital station and angular position of its moving parts determine the area of the midsection of the orbital station, on a certain area of the midsection and the density of the atmosphere, as well as on the measured motion parameters orbital station determines its mass according to the expression

where Cx- the drag coefficient of the orbital station;

Sx- the area of the midship orbital station;

- geocentric radius vector of the center of mass of the orbital station, and the dot indicates time differentiation t;

- the strength of gravitational field of the Earth;

- speed station Greenwich coordinate system;

ρ - density incident on the station aerodynamic flow.



 

Same patents:

FIELD: information transmission systems; satellite communication systems; spacecraft control systems.

SUBSTANCE: proposed system includes spacecraft, mobile control member and global satellite navigational system 9GLONASS) including navigational satellites which are provided with onboard command-and-measuring system and inter-satellite communication system; spacecraft includes User's navigational equipment which is connected with onboard control system; User's navigational equipment operates in navigational fields of GLONASS system or GLONASS and GPS systems; onboard command-and-measuring system is used for connection with mobile control member which includes User's navigational equipment operating in navigational fields of GLONASS or GLONASS and GPS systems; it is connected with operation station performing function of spacecraft command post and flight control center; operation station is connected with command-and-measuring system establishing communication both with spacecraft and navigational satellites.

EFFECT: extended functional capabilities of system; enhanced stability of spacecraft control.

1 dwg

The invention relates to spacecraft

The invention relates to space technology and can be used in large-sized high-precision transformable structures, for example, mirror antenna space radio telescopes

The invention relates to a space and can be used for processing trajectory measurements in order to accurately determine the parameters of the near-circular orbit of the spacecraft

The invention relates to radar technology and can be used when designing information systems for solving tasks such as automatic berthing and docking of spacecraft, providing the best way of ships in the Bay, landing on unprepared sites, etc

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

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

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

1 dwg

FIELD: highly accurate ground and space systems.

SUBSTANCE: invention relates to frame-type stable-size bearing structures made of laminated polymeric composite material. Proposed integral framed construction made of laminated polymeric composite material consists of right angle ribs and their connecting units forming, together with ribs, a monolithic load-bearing skeleton made of layers of fibrous material impregnated with polymeric binder lying in plane of frame. Each rib and each unit has at least one layer of fibrous material fibers in which are orientated along longitudinal axis of rib, and layers of fibrous material fibers in which are orientated in directions corresponding to direction of longitudinal axes of other ribs.

EFFECT: increased stability and accuracy of positioning of frame units, reduced variations of thermomechical properties along length of ribs, provision of high accuracy of dimensions of articles.

2 cl, 2 dwg

FIELD: satellites of small mass and methods of mounting them on carriers.

SUBSTANCE: proposed mini-satellite has body in form of parallelepiped, solar battery panels secured on its side plates and units for connection with separation system which are located on one of side plates and on end plate. Each panel is made in form of tip and root parts articulated together. Root parts of panel are articulated on side plate of mini-satellite body where connection units are mounted. Mechanical locks mounted on opposite side plate are used for interconnecting the tip parts of panels and for connecting them with body. Articulation units are provided with drives for turning of parts. Articulation units are located above mechanical locks relative to plate on which these locks are mounted. Novelty of invention consists in reduction of area of end part of mini-satellite by 35%, reduction of its height in center of cross section by 23%, reduction of mass by 6-7% and increase of density of arrangement by 17-18%.

EFFECT: enhanced efficiency; increased number of mini-satellites carried on adapter.

7 dwg

FIELD: spacecraft inter-planetary flights with the aid of cruise jet engines, mainly electrical rocket engines.

SUBSTANCE: proposed method includes injection of spacecraft into heliocentric trajectory at distance of spacecraft from Sun followed by its approach to Sun. Active motion of spacecraft is realized in part of this trajectory behind Earth's orbit during operation of jet engines. Then, spacecraft returns to Earth at velocity increment and increases its heliocentric velocity in the course of gravitational maneuver near Earth. After spacecraft has crossed Earth's orbit in section of its approach to Sun and before entry into Earth's gravisphere, spacecraft is accelerated by repeated switching-on of cruise jet engines. At the moment of fly-by over Earth when gravitational maneuver is performed, angular motion of Earth and spacecraft relative to Sun are equalized. During fly-by over Earth, spacecraft is subjected to its gravitational field changing the vector of spacecraft heliocentric velocity, thus ensuring further acceleration of spacecraft and forming inter-planetary trajectory of flight to target.

EFFECT: reduction of time required for organization of gravitational maneuver in Earth's gravity field at injection of spacecraft into required inter-planetary trajectory.

2 dwg

FIELD: control of group of satellites in one and the same orbit or in crossing longitude and latitude ranges of geostationary orbit.

SUBSTANCE: proposed method consists in measurement of parameters of satellite orbits, determination of orbital elements, comparison of them with required ones and performing of correcting maneuvers with the aid of thrusters. Satellite inclination vectors are brought to circular areas of their permissible change which are spaced apart so that angle between line connecting the end of vector with center of its circular area and direction to Sun should exceed right ascension of Sun by 180°. According to first version, vectors of satellite eccentricity are shifted to similar circular areas so that similar line lags behind direction to the Sun by half angular displacement of vector over circumference of its natural drift within circular area. Then, distances between satellites are changed within required limits compensating for quasi-secular increment of inclination vector and correcting eccentricity vector so that at passing the center of interval between point of circumference entry of its natural drift to its circular area and point of exit from this area, line connecting the center of this circumference and center of circular area coincide with direction to the Sun. In case circular area of permissible change of each eccentricity vector is close to circumference of its natural drift (second version), said line for this area is matched with direction to the Sun and no correction is made in this case.

EFFECT: saving of propellant for correction; protracted flight of satellites at safe distance.

3 cl, 13 dwg

FIELD: global satellite information systems.

SUBSTANCE: proposed system and method of organization of communication with the aid of this system includes injection of satellites into inclined elliptical orbits ensuring simplified tracking of satellites by means of ground tracking stations. Satellite orbits form pair of repeated routes (130, 140) embracing the earth's globe in projection on ground surface. Satellites are activated on each of these routes only on active arcs located considerably higher or lower relative to equator, thus emulating some essential characteristics of geostationary satellites. Parameters of satellite orbits are so set that final points of active arcs of two routes coincide; point at which active arc terminates in one route coincides with point where active arc starts on other route. Satellites placed on such active arcs are accepted by ground station located in satellite servicing zone as satellites slowly moving in one direction at rather large elevation angle. Their trajectory in celestial sphere has shape of closed teardrop line.

EFFECT: increased capacity of global satellite communication system with no interference in operation of geostationary satellites; simplifies procedure of tracking satellites.

39 cl, 15 dwg, 1 tbl

FIELD: illumination of separate sections of planet surface at night.

SUBSTANCE: proposed method consists in illuminating the night surface of planet by sunlight. One section is formed as circle and other sections are formed as circular rings of large diameter. All sections are lighted-up by rays scanning at rate no less than several revolutions per second setting-up these sections in order of increase of their diameters. Largest circular ring whose outer diameter is equal to size of surface being illuminated is set-up last.

EFFECT: increased area of illuminated sections at one and the same sizes of solar reflectors.

1 dwg

FIELD: space engineering; temperature control systems of automatic spacecraft flying in near-earth orbits.

SUBSTANCE: proposed method includes removal of excessive heat from instruments through two first and two second evaporators interconnected in longitudinal direction; said evaporators are made in form of L-shaped adjustable thermal tubes. Removal of heat from condensers of these tubes is effected to first and to second radiators-emitters of U-shaped heat-conducting honeycomb unit located orthogonally relative to first ones. Inner surfaces of side radiators-emitters are provided with heat insulation and side radiators-emitters have edges projecting beyond boundaries of instrument container. Their inner surfaces are provided with heat-controlled coat. Temperature control system is provided with two instrument containers interconnected by their center honeycomb panels. Side radiators-emitters are located in parallel or orthogonal planes. Built in structure of each U-shaped honeycomb units are L-shaped thermal tubes in such position that their condensers are located in side radiators-emitters and evaporators are located in center honeycomb panel. System provides for narrow range of control of seats of instruments mounted on center honeycomb panels.

EFFECT: enhanced efficiency of temperature control; enhanced reliability of spacecraft; extended field of application.

6 cl, 6 dwg

Micro-satellite // 2268205

FIELD: rocketry and space engineering; development of new satellites and updating of present artificial satellites, 20 to 100 kg in mass.

SUBSTANCE: some parts of onboard equipment in proposed satellite are secured on radiation surface on inner side of solar battery framework. Other parts are located in pressure container; number of parts is dictated by condition of retaining thermal mode of equipment in pressure container. Solar sensors are mounted on end face of pressure container coupled with micro-satellite separation system. Sensors for operation with ground stations and sensors of scientific equipment are located on other end face of pressure container. They are mounted on rectangular plate fitted with bracket in form of parallelepiped whose side faces carry scientific sensor equipment. Mounted on opposite surface of plate are sensors for operation with ground stations and extensible gravitational boom. Mounted on end cargo of boom are optical sensors. Four L-shaped antennae are located in parallel relative to faces of plate.

EFFECT: reduction of satellite mass (by 10-12%) and its length (by 14-15%); enhanced reliability.

5 dwg

FIELD: illumination of separate sections of night surface of planet by reflected sunlight.

SUBSTANCE: proposed satellite includes body and reflector disk. Axis of rotation of satellite body and central part of reflector disk intersect in center of mass of reflector disk. In orbit, satellite is twisted around axis of rotation of its body. Sun ray reflecting from revolving reflector disk forms light spot bypassing the circular section of surface being illuminated.

EFFECT: increased area of illuminated sections of planet surface at one and the same sizes of reflector.

2 dwg

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

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

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

3 cl, 2 dwg

FIELD: space engineering; spacecraft flying in geostationary or high-altitude elliptical orbits.

SUBSTANCE: proposed spacecraft has module case with projecting members. Two opposite faces of each module perform function of radiators with built-in thermal tubes. Arranged in modules are engine unit and some heat-loaded units and onboard devices (number n). Other units (number k) , for example metal-hydrogen storage batteries are secured to engine unit and are heat-insulated from first units. Units and devices are secured to engine unit by means of brackets through heat-insulating gaskets. Unit is made in form of three-layer honeycomb panel where thermal tubes with heaters are laid. Each of k-units has thermal contact with axial U-shaped thermal tube embracing them. This thermal tube is brought in contact with evaporator of loop thermal tube connected with radiator by means of vapor line which is communicated with loop thermal tube and its evaporator through condensate lines. Radiators are mounted beyond boundaries of projecting parts shading zones on side of extravehicular space.

EFFECT: increased cooling effect of spacecraft temperature control system; reduction of mass of this system.

2 cl, 4 dwg

FIELD: illumination of separate sections of night surface of planet by reflected sunlight.

SUBSTANCE: proposed satellite includes body and reflector disk. Axis of rotation of satellite body and central part of reflector disk intersect in center of mass of reflector disk. In orbit, satellite is twisted around axis of rotation of its body. Sun ray reflecting from revolving reflector disk forms light spot bypassing the circular section of surface being illuminated.

EFFECT: increased area of illuminated sections of planet surface at one and the same sizes of reflector.

2 dwg

Micro-satellite // 2268205

FIELD: rocketry and space engineering; development of new satellites and updating of present artificial satellites, 20 to 100 kg in mass.

SUBSTANCE: some parts of onboard equipment in proposed satellite are secured on radiation surface on inner side of solar battery framework. Other parts are located in pressure container; number of parts is dictated by condition of retaining thermal mode of equipment in pressure container. Solar sensors are mounted on end face of pressure container coupled with micro-satellite separation system. Sensors for operation with ground stations and sensors of scientific equipment are located on other end face of pressure container. They are mounted on rectangular plate fitted with bracket in form of parallelepiped whose side faces carry scientific sensor equipment. Mounted on opposite surface of plate are sensors for operation with ground stations and extensible gravitational boom. Mounted on end cargo of boom are optical sensors. Four L-shaped antennae are located in parallel relative to faces of plate.

EFFECT: reduction of satellite mass (by 10-12%) and its length (by 14-15%); enhanced reliability.

5 dwg

FIELD: space engineering; temperature control systems of automatic spacecraft flying in near-earth orbits.

SUBSTANCE: proposed method includes removal of excessive heat from instruments through two first and two second evaporators interconnected in longitudinal direction; said evaporators are made in form of L-shaped adjustable thermal tubes. Removal of heat from condensers of these tubes is effected to first and to second radiators-emitters of U-shaped heat-conducting honeycomb unit located orthogonally relative to first ones. Inner surfaces of side radiators-emitters are provided with heat insulation and side radiators-emitters have edges projecting beyond boundaries of instrument container. Their inner surfaces are provided with heat-controlled coat. Temperature control system is provided with two instrument containers interconnected by their center honeycomb panels. Side radiators-emitters are located in parallel or orthogonal planes. Built in structure of each U-shaped honeycomb units are L-shaped thermal tubes in such position that their condensers are located in side radiators-emitters and evaporators are located in center honeycomb panel. System provides for narrow range of control of seats of instruments mounted on center honeycomb panels.

EFFECT: enhanced efficiency of temperature control; enhanced reliability of spacecraft; extended field of application.

6 cl, 6 dwg

FIELD: illumination of separate sections of planet surface at night.

SUBSTANCE: proposed method consists in illuminating the night surface of planet by sunlight. One section is formed as circle and other sections are formed as circular rings of large diameter. All sections are lighted-up by rays scanning at rate no less than several revolutions per second setting-up these sections in order of increase of their diameters. Largest circular ring whose outer diameter is equal to size of surface being illuminated is set-up last.

EFFECT: increased area of illuminated sections at one and the same sizes of solar reflectors.

1 dwg

Up!