Method of arrangement of stationary earth man-made satellite
FIELD: space engineering.
SUBSTANCE: proposed method consists in using a vessel aboard the satellite with a required constant area of the liquid phase location on the said vessel wall surface. The Earth satellite lengthwise axis, in operating conditions, is constantly directed along the current radius-vector of the orbit mass centre. The aforesaid vessel is placed at a maximum possible distance from the satellite mass centre along the direction parallel to the said lengthwise axis. The vessel should be placed so that the normal at the point of the vessel wall surface opposite the working medium outlet was directed towards the satellite mass centre. The said point should be located at a minimum possible distance from the satellite lengthwise axis.
EFFECT: simpler design and smaller weight of service satellite systems.
The invention relates to space technology, in particular to a stationary artificial satellites of the Earth, and created by the authors in order to perform their tasks.
As part of the spacecraft, such as telecommunication satellites, used capacity, in which during orbital operation is stored a certain stock of a working body:
in the attitude control system and motion control (SOUD) (spacecraft./Under the General editorship Ceptionist. M, Military publishing house, 1983, p.54, 56 ) is the capacity for storing the working fluid, an internal cavity which is initially charged, for example, a xenon under pressure ≈200 kgf/cm2; during the active life of the satellite, the number of xenon gradually decreases and at pressures below the critical pressure (≈59,45 kgf/cm2) and temperatures below the critical temperature (≈289,7K) (see Nbhrhieq. Handbook of thermophysical properties of gases and liquids. M., Nauka, 1972 str ) in the tank is formed of a two-phase environment. From the point of view of the functional purpose of the tank into the pressure reducer, then a jet nozzle for the normal functioning of the xenon must come in the form of vapour, or in the case of supply from the tank the liquid phase, the flow rate of xenon will be significantly overestimated and therefore the conserve mass xenon her to the gearbox must be converted into steam, i.e. in SOUD there is an additional device is quite powerful heater (for example, at a flow rate of xenon ≈2 g/with the power of the heater is equal to ≈120 watts, which is engaged in work on the job SOUD in case of presence of the liquid phase near the holes for the supply of xenon in-line reducer);
in thermal control system (P) (according to the author's certificate of the USSR No. 2117891 ) is a device for maintaining the pressure of the fluid in the circuit P of the SPACECRAFT (SC), the inner cavity of which is filled by a certain number of two-phase working fluid (coolant), for example, ammonia; from the point of view of a functional purpose in terms of orbital operation should be capable of delivering (or receiving) the liquid phase ammonia in the line P and to maintain a certain vapor pressure of ammonia in the gas cavity of the container.
As can be seen from the above, in containers of both satellite systems for normal and optimal functioning of the liquid phase of the working heat in the tanks must be in a certain permanent zones on the inner surface of the vessel wall:
- in capacity SOUD - opposite holes for supplying the vapor of the working fluid in the supply line to the gearbox;
- capacity P - near the outlet for flow of the liquid phase of the capacity is in line P and near heaters.
Currently known technical solutions to ensure the normal functioning of the above systems during orbital operation:
in SOUD provide a heater system for its automatic inclusion in the starting feed reducer liquid phase xenon or, in the absence of electric heaters provide an additional supply of xenon in the vessel with greater capacity to provide health SOUD within a specified period of operation, which complicate the design and increase the weight, the power consumption of the above systems;
- capacity PAGE within its total volume and the area of the outlet set the capillary structure and the heaters complex designs (such a complex and massive structure is provided on the assumption that during orbital operation position of the liquid phase ammonia in the tank undefined).
In the known technical solutions heaters incorporated in the work periodically (for example, 0.5 h included, then at least 2-2 .5 h off) to create (and maintain) a certain operating pressure of the vapor phase of the working fluid in the valid range. When the radiator is boiling of the liquid phase wetting the surface of the vessel wall, in the zone opposite the location of the heaters. In the boiling (vaporization) of the liquid phase, the particles of the liquid and vapor are released into the Central zone of the vessel. After turning off the heater fluid particles according to known physical law (VII. The motion of an artificial satellite about the center of mass. M., Nauka, 1965, page 24, figure 2 on p.25, p.26 (paragraphs 1 and 2 below), p.27 (the formula for the components of the centrifugal force: z and FG, Fu) ) will have a certain direction - they will move parallel to the radius-vector of the center of mass of the satellite in the direction of the center of mass of the satellite.
Because it is not currently known technical solution, which stipulates the specific placement of the above capacity on the satellite, these fluid particles may move in the direction opposite to the desired direction - this means that the liquid phase will gradually collect on the inner surface of the container opposite from the desired, i.e. in this case, the constructive arrangement of the capacity on the satellite is not conducive to the concentration of the liquid phase in a certain fixed area on the inner surface of the vessel wall, and this technical problem currently requires permission.
Analysis of sources of information - the patent and scientific-technical the Russian literature showed that the closest to the technical nature of the prototype of the proposed technical solution is the way to build capacity SOUD for storing the working fluid - xenon onboard the satellite .
Currently, the requirements specified by the layout of the containers used, for example, on a stationary artificial Earth satellite, not specified (see figure 2, where C is the center of the Earth (center of gravity); the center of mass (center of inertia) of the satellite; Oxyz - right rectangular orbital coordinate system (ACI. The motion of an artificial satellite about the center of mass. M., Nauka, 1965, p.17 (3rd paragraph from the bottom), 23 (3rd paragraph from the top); r is the current radius-vector of the satellite's orbit; A - stationary satellite orbit (the orbit of the center of mass of the satellite); 1 - the payload module of the satellite; 2 - module service systems satellite; 3 - capacity SOUD; 3.1 - hole for exhaust vapor of the working fluid from the tank; 4 - pipe selection the vapors of the working fluid; 5 - pairs of the working fluid; 6 - liquid phase working fluid; 7 - electric heater; 8 - longitudinal axis of the satellite).
As indicated above, a significant disadvantage of the known device is increased weight due to the need to carry an additional supply of the working fluid and power consumption is to turn the liquid phase of the working fluid in the vapor state.
The purpose of accom is amago the authors of the technical solution is to eliminate the above-mentioned significant disadvantages.
This goal is achieved by the arrangement of the satellite in such a way that the specified capacity is set at the greatest possible distance from the center of mass of the satellite in a direction parallel to a specified longitudinal axis with the container so that the normal at the point of the surface of the wall opposite the hole for the exhaust vapor of the working fluid from the tank, was directed toward the center of mass of the satellite, and this point was at the minimum possible distance from the longitudinal axis of the satellite, which is, according to the authors, significant distinctive features proposed by the authors of the technical solution.
The analysis conducted by the authors of the famous patent and scientific literature, the proposed combination of significant distinguishing features of the proposed technical solution in the well-known sources of information are not detected and, therefore, the known technical solutions do not exhibit the same properties as in the present method, the composition of the artificial satellite.
The layout of the satellite according to the authors, the method is as follows (see figure 1, where C is the center of the Earth; center of mass; Oxyz - right rectangular space coordinate system; r is the current radius-vector of the satellite's orbit; A - stationary orbit satellite is; 1 - the payload module, 2 - module service systems; 3 - capacity SOUD; 3.1 - hole for exhaust vapor of the working fluid from the tank; 4 - pipe selection the vapors of the working fluid; 5 - pairs of the working fluid; 6 - liquid phase working fluid; 7 - heater; 8 - longitudinal axis of the satellite):
- conducted requirements analysis for the arrangement of stationary artificial Earth satellite, the longitudinal axis 8 axis of least moment of inertia or the most Central axis of the ellipsoid of inertia) which must always be directed along the current radius vector of the orbit r between the center of the Earth With the center of mass Of the satellite;
- analyze functional purpose of the vessel 3 the given construction, seasoned two-phase working fluid (working fluid), and define the desired constant area location of the liquid phase 6 on the surface of the wall;
- determine the location on the satellite, which as much as possible removed from the center of mass Of the direction parallel to the aforementioned axis 8;
- establish the capacity of the 3 on-Board the satellite in vysovyvanii zone so that the normal n at the point C of the surface of the wall opposite the hole for the exhaust vapor of the working fluid from the tank, was directed toward the center of mass of the satellite, and the specified point C was located at the minimum possible distance about the longitudinal axis 8 of the satellite.
Conducted on the basis of  calculations for the capacity of the particular telecommunication stationary satellites on Board the satellite according to the new proposed technical solution, show that the acceleration of fluid particles is not less than 3·10-8m/s2and for ≈40 minutes these particles are the distance between the mutually opposite areas of capacity, i.e. to the time of the next switching of the heater liquid phase will be in the zone capacity, thereby ensuring normal operation.
Thus, as a result of this arrangement the liquid phase 6 of the coolant during orbital operation will always be required in the area: as a result of combined effects of the gravitational and centrifugal forces (which by the magnitude and direction relatively stable in stationary orbit) on fluid particles they will move parallel to the radius-vector r of the center of mass Of the satellite in the direction of the center of mass Of the satellite and will be concentrated on the surface of the vessel 3 in the area in front of her holes 3.1 for removal of vapor of the working fluid from the tank 3 and is pressed to the surface in the zone of the point C, and at the same time under the influence of disturbances (the influence of the moon and Sun, the nonspherical character of the Land, intermittent periodic work SOUD) liquid phase th the media, with slight fluctuations, to be in this zone capacity , thus, by providing the exhaust vapor of the working fluid in the pipeline selection.
It should additionally be noted that in the case of maintaining the temperature of the wall, located in the area opposite said hole, below the temperature of the rest of the aspiration of the liquid phase to the above area will be further enhanced by the high degree of wettability over a cold fluid compared with warm - to do this they need this amount of capacity to insulate insulation with a lower thermal resistance than the rest of the insulation.
Thus, as can be seen from the above, in the layout KA according to the technical solution of the reduced mass SOUD as a result of filling its capacity with the optimal amount of the working fluid, and a heater for converting the liquid phase into vapor is backing element and, consequently, the power consumption of it within the operation of the minimum possible, i.e. thereby achieved the objectives of the invention.
Currently proposed by the authors of the technical solution are reflected in the technical documents of the company for the development of the newly created telecommunications satellites.
The positioning of the stationary artificial Earth satellite, the longitudinal axis of which must be constant the NGOs are directed along the current radius vector of the orbit, connecting the center of the Earth with the center of mass of the satellite, including installation on Board the vessel, filled two-phase working fluid, the desired permanent location area of the liquid phase on the surface of the vessel wall, characterized in that the specified capacity is set at the greatest possible distance from the center of mass of the satellite in a direction parallel to a specified longitudinal axis with the container so that the normal at the point of the surface of the wall opposite the hole for the exhaust vapor of the working fluid from the tank, was directed toward the center of mass of the satellite, and this point was at the minimum possible distance from the longitudinal axis of the satellite.
SUBSTANCE: invention concerns devices for improvement of aerodynamic properties of aircrafts, mostly rocket launchers (RL). The fairing comprises nose cone (1), cylindrical compartment (2), rear extension (3) of the last stage (4) of RL. The fairing features permeable porous rim (5) and insert (6) that damp pressure fluctuation (in the flow separation zones). Biconical nose cone (1) is preferred, and semi-vertex angle of the first cone (7) is 25°- 35°, while semi-vertex angle of the second cone (8) is 13°- 25°. Length of the first cone is 0.2 - 0.25 of the total nose cone (1) length. The rim (5) and insert (6) are both as long as 0.11 of the cylindrical compartment (2) length. Total length of this compartment exceeds its diametre at least by 1.11 times. The extension (3) can have partition ribs made of permeable or pressure pulsation-damping material.
EFFECT: lower aerodynamic load at RL head.
10 cl, 2 dwg
FIELD: physics, space technology.
SUBSTANCE: plot board for ground surveillance object selection from orbital space vehicle refers to space technology. The plot board for ground surveillance object selection from orbital space vehicle includes flexible tape with the ground map printed on it, and semitransparent plate above the tape. Two halves of the space vehicle's orbit pass are drawn on the plate in such a way that the ascending node in the beginning of the first half of the orbit pass and the descending node in the beginning of the second half of the orbit pass are superimposed. The plot board also comprises a device for the tape map transport along the plate with orbit pass image; the transport device comprising two shafts that are spaced-apart and interconnected in parallel. The map printed on the tape has the equator beginning and ending points superimposed, and the tape is made in the form of a ring and pulled on the shafts so as to move along the map equator line. The distance between the shaft axes and size of the plate with the orbit pass image along the equator direction are equal to (L-d)/2, where L - is the map equator length; d - is absolute distance between passes measures in linear units along the map equator; shaft radius equals to d/(2π). The stated above gives the effect of decreased plot board size.
EFFECT: decreased size of plot board for ground surveillance object selection from orbital space vehicle.
FIELD: space engineering; large-sized high-precision transformable structures.
SUBSTANCE: proposed method consists in monitoring acceleration and respective deviations of geometric parameters from theoretical magnitudes at check points. Monitoring is carried out continuously in real time. Check points are made for all shape-forming members of antenna structure. Device proposed for realization of this method has geometry and acceleration monitoring system which is made in form of combined spatial position and acceleration sensors. Sensors are made in form of miniature three-axis units of gyroscopes-accelerometers which are electrically connected with onboard information-and-measurement control system via analog-to-digital converters. Onboard information-and-measurement control system is electrically connected via respective power amplifiers with actuating members of extended structural members of spacecraft.
EFFECT: enhanced reliability of parameters under test; improved characteristics of antenna.
3 cl, 3 dwg
FIELD: space engineering; manufacture of artificial satellites and other spacecraft.
SUBSTANCE: proposed instrumentation module has body made from honeycomb panels. Thermostatted plate of payload is mounted on body with the aid of brackets. Equipment of instrumentation module is installed on honeycomb panels, mainly inside module. Honeycomb panels are provided with labyrinth-type vent holes and technological holes for introduction and removal of structural members: rods, pipe lines, cable bundles, etc. Clearances between honeycomb panels of body of instrumentation module, between instrumentation module and thermostatted plate, between edges of technological holes and said structural members are shielded by means of optically opaque member which is made from material having electrically conducting layer ensuring electrical tightness of instrumentation module. Shielding member is also provided with labyrinth-type vent holes.
EFFECT: enhanced protection of equipment against electromagnetic radiation, spurious currents and action of charged particles of natural or technogenous nature (for example, magneto-sphere plasma and jets of electric rocket engines).
3 cl, 4 dwg
FIELD: rocketry and space engineering; mounting the device on spacecraft external surface and subsequent separation of ultra-red target in form of inflatable thin-film envelopes with dark coat.
SUBSTANCE: proposed device has cassette with inflatable envelopes and doors, nitrogen generators with pyro retarders, starting device with spring mechanism for ejection of cassette from starting device and electric connector. Door axles are provided with torsion springs. Nitrogen generators provided with check valves are located on one side from thin-film envelopes. Height of ejection mechanism in compressed state is comparable with diameter of each nitrogen generator. Electric cable of electric connector has free loop whose length is equal to path of motion of cassette till moment of steady ignition of pyro retarders. Circuit supplying the electric pulse to igniter of each pyro retarder is provided with terminal closing this circuit at the beginning of motion of cassette.
EFFECT: reduced overall dimensions of device; enhanced reliability; simplified construction.
FIELD: rocketry and space engineering; separable rocket nose cones.
SUBSTANCE: proposed rocket nose cone has body made in form of envelope which is closed at one end and is provided with frame on other end which is not closed; rocket nose cone is provided with tubular force exciters. Body has weakened section in plane of separation. Body forms inner closed chambers in plane of separation where tubular force exciters of directive action are located. Chambers are formed by body envelope and by its longitudinal fins. Walls of these fins are directed to plane of separation and are parallel relative to it; they are strengthened by transversal structural members fastened to body envelope. Detachable covers are secured on longitudinal fins of chambers by means of fasteners; they have weakened sections in plane of separation of nose cone which are similar to those made in body; their areas are equivalent.
EFFECT: facilitated procedure of manufacture; avoidance of dynamic action on rocket in the course of separation; avoidance of collision in the course of and after separation.
4 cl, 4 dwg
FIELD: space engineering; other industries; production of the stable-sized platforms.
SUBSTANCE: the invention is pertaining to the load-carrying structures made out of the laminated polymeric composite materials and may be used in the high-precision space and ground equipment, for example, as the support for the optical instruments, the antenna systems and the measuring systems. The presented platform is made in the form of the plane annular or circular centrally-symmetric board and contains the encasings made out of the layers of the fibrous material impregnated with the polymeric binding, the cell-type filler placed between the casings and the attachment points disposed with the equal angle pitch. Each layer of the casings consists of the sectors docked among themselves with the equal central angle. Quantity of the sectors in each layer is equal or multiple to the number of the attachment points. In each sector of one layer the filaments are oriented at the equal angle concerning the central axis of the sector. T sectors of each subsequent layer are shifted concerning the sectors of the previous layer at the angle equal to the half of the central angle of the sector. In each sector of the same layer the filaments may be oriented at the angle of 90° to the central axis of the sector. There may be present the layers, where the filaments are oriented at the angle of 0° to this axis. The technical result of the invention is to ensure the control over the thermal deformation of the platform with the purpose to reach the given accuracy of positioning of the attachment points on it at fulfillment of the strength and rigidity requirements to its design.
EFFECT: the invention ensures the control over the thermal deformation of the platform with the purpose to reach the given accuracy of positioning of the attachment points on it at fulfillment of the strength and rigidity requirements to its design.
3 cl, 3 dwg
FIELD: centrifugal frameless structures formed in space and used for deployment of solar batteries, reflectors and other large-sized systems.
SUBSTANCE: proposed method includes placing the flexible sectors on carrier, rotating the carrier in plane corresponding to working position of frameless centrifugal structure and deploying the sectors from carrier under action of centrifugal forces. Sectors are interconnected by side edges forming single working surface in the course of their deployment and preliminarily when necessary. Additional deploying force is applied to sectors along joint areas from periphery of centrifugal frameless structure to its center. Device proposed for realization of this method has carrier for placing the sectors, its rotation drive, as well as drive and mechanism for extension of sectors. Articulated on bearing part of extension mechanism are brackets provided with pairs of hold-down and drive rollers. One sector joint area is passed through each pair of roller. Drive roller is provided with drive which is kinematically aligned with sector extension drive, thus forming additional deploying force. Side edges of sectors may be connected at points of application of this force (or near them) with the aid of connecting elements, such as zipper, as well as by welding, bonding, sewing, etc. Carrier may be made in form of common drum or in form of reels separate for each sector. Device forming the centrifugal frameless structure has surface smoothly stretched in two axes.
EFFECT: enhanced reliability; facilitated procedure.
12 cl,, 11 dwg
FIELD: space power engineering; film-type solar batteries on base of amorphous silicon.
SUBSTANCE: proposed solar battery has central power member. Solar battery consists of two sections. Each section is formed from standard film-type trihedral prisms on base of inflatable tubular skeleton. Outer surface of this skeleton is coated with compound which gets hardened under action of ultraviolet and visible solar radiations. Solar battery deployment system includes two electric motors of central power member and additional electric motor. Inputs of these electric motors are connected with outputs of pitch, yaw and roll channels of solar battery control unit. Solar battery is additionally provided with additional position electric motors which are used for discrete turn of each trihedral film-type prism through angle of 0o to 360o at pitch of 120o. Specification gives description of solar battery modification which includes reserve film-type panels increasing active life of solar battery. Total power of proposed solar battery is about 120 kW.
EFFECT: enhanced reliability of film-type panel tension.
3 cl, 3 dwg
FIELD: specialized spacecraft for refilling self-contained spacecraft with cryogenic agents (liquid nitrogen, liquid helium) and propellant components (liquid oxygen, liquefied methane, hydrazine).
SUBSTANCE: proposed filling module has load-bearing skeleton in form of hexahedral prism and peripheral coupling units, movable truss structure equipped with additional power electric drives and quick-detachable locking devices. Truss has open inner space where changeable cryogenic and propellant reservoirs are located for storage and transportation of cryogenic agents and propellant components. Availability of locking devices and movable truss structure makes it possible to perform repeated operations of replacing empty reservoirs of self-contained spacecraft with filled ones.
EFFECT: extended functional capabilities of orbital filling module; considerable reduction of losses of components (up to 5-7% of total mass).
FIELD: space-system engineering.
SUBSTANCE: system can be used on the earth satellite vehicle in orbit except geostationary centers which are stabilised by rotation along the vertical axis, and the ground reception centers. The system contains an emergency object transmitter, spacecraft equipment and ground reception center equipment. Spacecraft equipment includes a horizon sensor, receiving antenna, comparison device, receiver, Doppler frequency meter, blocking-generator, two gating circuits, two valves, pulse generator, pulse counter, plug board chart, and magnetic storage device, transmitter, transferring antenna, modulation code generator, high-frequency generator and power amplifier. The ground reception center equipment includes reception antenna, high-frequency amplifier, two mixers, calibration frequency block, phase doublet, three narrowband filter, scale-of-two phase circuit, phase detector, Doppler frequency meter, computing unit and registration block.
EFFECT: it enlarges functional possibilities of the system.
FIELD: production methods.
SUBSTANCE: suggested method includes emission and retranslation of primary and final radio signals between spaceship, basic and additional land determinations stations. At the same it is additionally retranslated the final radio signal from spaceship to the basic land station and it is admitted at this station. The radio connection with the radio signal of the spaceship with one or more additional stations, admission of the primary signal to the additional station, its transformation to the final signal and admitting of it to the spaceship. About the distance between spaceship and main determination station is judged by the interval between the moment of emission and moment of admitting the primary signal at this station. About the distance between spaceship and additional determination station is judged by the interval between the moment of emission and moment of admitting the primary signal at the main station. It is measured additionally the moving of the frequency of the final signal, admitted on the main determination station, regarding the frequency of the primary movement, emissed from the same station. The distance between spaceship and additional station is determined with measuring of the Doppler drift.
EFFECT: it is reduced the time and increased the accuracy of distance determination between spaceship and determination stations.
5 cl, 4 dwg
SUBSTANCE: suggested method includes the conduction of heat gain from instruments, installed on middle heat-conducting cellulite panel to the side radiator-oscillator by L-shaped regulating heating tubes (HT). At the same the instrument container of the spacecraft is created by union of two flat-topped heat-conducting cellulite panel blocs, and in each of them it is installed the L-shaped HT. The vaporizers of this HT are made by pairwise connection in the longitudinal direction in the middle cellulite panel. The condensers of HT are made in the side cellulite panel of flat-topped block, which is the side radiator-oscillator. Besides L-shaped HT, for additional heat-conducting from one internal surfaces of the radiator-oscillator and transmission of it to the others radiator-oscillator it is used parallel flat-topped HT. The vaporizers and condensers of the HT are built into additional cellulite panel layer of the proper side radiator-oscillator and orthogonal situated relatively the condensers of L-shaped HT.
EFFECT: it is increased the exactness and safety of the work and it is exceeded the possibility of using the system of thermoregulation.
2 cl, 4 dwg
FIELD: space engineering; forming satellite systems for positioning objects on earth surface.
SUBSTANCE: proposed method includes injection of N artificial satellites into circular or other orbits which work in "n" planes (where "n" is integer which is more than 2) by mi (i=1, ...n) satellites (where n is integer) in each plane. Satellites are positioned in orbits of datum plane and in planes located symmetrically and in pairs relative to datum plane. These planes of orbits are positioned irregularly along terrestrial equator relative to datum plane at angles a priori not equal to 360°/n. Artificial earth satellites in orbits are positioned irregularly and symmetrically in pairs relative to base satellite.
EFFECT: reduced number of artificial earth satellites in navigational system with no impairment of system parameters at positioning of ground objects.
1 dwg, 1 tbl
FIELD: space engineering; operation of spacecraft flying in orbit of artificial earth satellite, but for geostationary orbit, which are stabilized by rotation along vertical axis, as well as ground reception points.
SUBSTANCE: system used for realization of this method includes emergency object transmitter, onboard equipment of spacecraft and ground equipment of reception point. Onboard equipment of spacecraft includes horizon sensor, receiving antenna, comparison unit, receiver, Doppler frequency meter, blocking oscillator, two AND gates, two rectifiers, pulse generator, pulse counter, switching circuit, magnetic memory, transmitter, transmitting antenna, modulating code shaper, RF generator and power amplifier. Ground equipment of reception point includes receiving antenna, RF amplifier, two mixers, standard frequency unit, phase doubler, three narrow-band filters, phase scale-of-two circuit, phase detector, Doppler frequency meter, computer and recording unit. Proposed method consists in search of such space position of space object by receiving antenna when Doppler frequency of received signal is equal to zero. Measurement at this moment of angle between mechanical axle of receiving antenna and horizon axis is carried out referring to onboard receiving unit.
EFFECT: extended functional capabilities; enhanced accuracy of determination of spacecraft orbit elements; reduction of time required for search of emergency object.
FIELD: radio engineering.
SUBSTANCE: device has bed unit mounted on lateral surface and pivotally connected to device casing. Bed unit rotation axis is arranged in parallel to casing end face part. Mechanisms for rotating and fixing the bed unit relative to the casing (optionally of screw-nut type) are mounted on both sides with respect to the axis. The mechanisms provide supporting bed unit surface arrangement at an angle less than 90°. Solar battery board is rigidly mounted on the supporting bed unit surface. Camera for taking Earth surface pictures is mounted in nanosputnik casing end face on the opposite side with respect to nanosputnik units for connecting it to separation system. The camera is arranged in plane passing through longitudinal nanosputnik axis arranged in perpendicular to the solar battery board.
EFFECT: reduced device weight; increased effective solar battery board area.
FIELD: space engineering; spacecraft flying in earth artificial satellite orbit, but for geostationary orbit stabilized by rotation along vertical axis.
SUBSTANCE: system used for realization of this method includes spacecraft case, infra-red horizon pulse sensor, receiving antenna, comparison unit, receiver, Doppler frequency meter, biased blocking oscillator, two AND gates, two rectifiers, pulse generator, pulse counter, switching circuit, magnetic storage, transmitter, transmitting antenna, onboard timing device, onboard master oscillator and emergency object transmitter. Doppler frequency meter includes 90-deg phase shifter, two mixers, two difference frequency amplifiers, 180-deg phase inverter, two AND gates and reversible counter. Frequency of received oscillations is preliminarily reduced in two processing channels.
EFFECT: enhanced accuracy of determination of coordinates due to accurate measurement of minor magnitudes of Doppler frequency and recording its zero magnitude.
FIELD: power supply systems for high-orbit and geostationary orbit communication satellites whose orbits are corrected by means of electric jet engines.
SUBSTANCE: proposed method consists in determination of power requirements for each items of onboard using equipment for all spacecraft of cluster, electric jet engines inclusive. In case some items of onboard using equipment are not provided with electric energy at interval of dynamic mode with the aid of electric jet engines, this mode is changed-over for another permissible interval. In case of absence of permissible intervals, duplicate spacecraft of equivalent payload are selected. Items of equivalent using equipment of main spacecraft which are not provided with electric power are changed-over to duplicate spacecraft which are provided with required electric power. Spacecraft equipment items are changed-over till restoration of power supply on board main spacecraft (upon completion of dynamic modes of these spacecraft). Then control of power supply is performed for spacecraft of orbital cluster of later performance of dynamic modes. Main spacecraft are used as duplicate spacecraft.
EFFECT: reduced power requirements of cluster spacecraft; possibility of supplying power for additional items of using equipment.
1 dwg, 1 tbl
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.
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.
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