Method of thermoregulation of spacecraft and device for its drying

FIELD: heating.

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

 

The present invention relates to space technology and can be used in systems management (P) automatic SPACECRAFT) with the instrument containers from sotapanna with the use of heat pipes (TT) or regulated heat pipes (PTT).

Each time in the process of creating a new CA we are searching for the best solutions of its composition and thermal control system (P), which allow the operating temperature of the devices in perhaps a more narrow range, as this increases the reliability of their operation, and CA in General. It is known that a decrease or increase of the working temperature tracking devices on their given medium temperature for 10°With the intensity of their failures is increased by 25% (see Gnidelines. Heat and mass transfer in electronic equipment. Moscow, Higher school", 1984, p.7).

Known platform KA S B-44 (Jean J.Dechezelles, Dietric E Koelle gn and application of the AS/MBB Spasebus Famile AJAA 11 Sattelite Communication systems, March 17-20, 1986, pp.688-696. R j 41, 1986, Ref-10.41.126). Platform KA contains unpressurized instrument container in the form of a parallelepiped with sitopaladi made in the form of radiators-radiators on the Northern and southern sides, the inner sides of which include on-Board devices. In the internal structure of the radiators-radiators built TT, through which excess heat from titsa from devices, distributed on the surfaces (skins) radiators-radiators, from which it is radiated into space.

In this analogue implemented method of thermal control of SPACECRAFT, in which the transfer of excess heat from devices that are installed on the inner surfaces of the radiators-radiators, evaporators unregulated TT and forth with their capacitors excess heat is removed for radiators-radiators with subsequent radiation in open space. Because the devices are mounted directly on radiators-radiators, partially excessive heat from devices directly allocated to unregulated radiators-radiators due to the heat its structural elements.

With this method of control when the SPACECRAFT is at the minimum internal and external thermal loads, the temperature of the devices is reduced to the maximum of the minimum values, which is not provided sufficiently reliable to work them, and the maximum heat loads on the SPACECRAFT operating temperatures of the devices are increased to maximum values at which the reliability of the devices is not high enough. This is because the principle of this device is based on the use of unregulated TT, and devices directly and without the use of the effect of the main means of control are related directly to the radiators (emitters).

As in the standard orientation of the SPACECRAFT for a more stable thermal control total difference in thermal stress between the appliances installed on the southern radiator-the radiator, and appliances installed on the Northern radiator-the radiator is inversely proportional to the difference external heat loads on the said radiators-radiators, in the event of staffing changes the orientation of the SPACECRAFT, for example, when the correction of the orbit North of the radiator-emitter can be illuminated by the Sun, and the South will be placed in a shaded position, it also leads to the exclusion of the possibility of providing temperature devices in range for their reliable operation.

Known thermal control system according to the invention "spacecraft modular design (patent RU No. 2092398, CL 64G 1/10, priority 24.10.1995).

That decision was implemented method of thermal control of SPACECRAFT, which by means of TT is given the excess heat from devices that are installed on the inner sides of the radiators-radiators N - and P-shaped stopplng blocks, as well as from devices installed on secondary sotapannas N - and P-shaped stopplng blocks using unregulated and l-shaped adjustable diode TT.

The disadvantage of this solution is that it does not provide a sufficiently high efficiency termoreg the modeling of devices and as a consequence, reduced the reliability of their work and the work of the AC as a whole. The reason for this is that the TT is made U-shaped and, therefore, unregulated, and l-shaped adjustable TT is built into the structure of sotapanna with parallel arrangement of their evaporators with a pitch (distance) 200 mm As a consequence, there is a large temperature difference between the TT, which degrades the accuracy of temperature compensation in the places of installation of the devices.

In addition, for maximum heat loads on the SPACECRAFT disadvantage is the inefficient use of contact thermal coupling between the capacitors TT average setpanel with casings radiators-radiators H-shaped stopplng block. This leads to a significant temperature rise of the devices relative to a set for them rated temperature level, and hence to further reduce the reliability of the instruments and the SPACECRAFT as a whole.

The decrease in the efficiency of thermal control devices are also connected with the fact that the devices installed on the radiators-radiators with their inner sides, so part of the excess heat from devices permanently allocated to unregulated radiators-radiators due to the heat and their temperature is strongly influenced by the temperature of the radiators-radiators and changes in NR is snih thermal loads on the radiators-radiators. Therefore, when the operation modes of the AC with the maximum and minimum heat load temperature devices are provided in a wide range, which reduces the reliability of their work.

The application of that decision drop-down using an active Electromechanical actuators thermal curtain system zackowski and rusckowski, as a forced measure the efficiency of thermal control of SPACECRAFT, on the one hand, has a positive effect, on the other hand, brings a negative effect, as applied active Electromechanical actuators and system zackowski and rusckowski have the reliability is less than one, and hence this reduces the reliability of the SPACECRAFT.

As the prototype is set to "Method of thermal control of the spacecraft and the device for its implementation" (Patent of Russia №2268207, IPC B64G 1/50; B64G 1/10).

That decision was implemented method of thermal control of SPACECRAFT, including the removal of excess heat from each device installed on the conductive setpanel through evaporators and condensers built-in honeycomb l-shaped adjustable heat pipes on the side radiators-radiators, in which the surplus heat removal is performed via pairs directly connected to each other in the longitudinal direction of the first two and second two evaporator of these l-shaped re wiremac heat pipes, and the heat from condensers these heat pipes with the specified first and second connected to each other evaporators operate respectively on the first side radiators-radiators and located orthogonal them second lateral radiators-radiators.

thermal control system of the spacecraft for the implementation of this method, including the instrument container with exterior insulation formed by combining two U-shaped heat-conducting stopplng blocks, each of which is designed with built-in inner structure with l-shaped parallel-arranged adjustable heat pipes with their capacitors in the side sotapannas made in the form of lateral radiators-radiators and evaporators in the middle of setpanel, on the inner side of which installed devices are made so that the evaporators of these l-shaped adjustable heat pipes each U-shaped heat-conducting stopplng unit is made directly in pairs connected to each with each other in the longitudinal direction, and the capacitors of each specified pair of heat pipes are made respectively in its lateral radiators-radiators; on the inner surface of the side radiators-radiators installed internal insulation; it is made with two similar to the instrument containers, joined by their planes one medium sotapanna with the location of the lateral radiators-radiators in parallel or orthogonal planes.

The disadvantage of the prototype is its limited capacity to ensure the accuracy of the temperature control medium setpanel and devices installed on it, and thus to ensure a sufficiently high reliability SPACECRAFT, provided its application in terms of the increased difference between the maximum and minimum heat loads on the SPACECRAFT. This is due, on the one hand, with the limited possibilities of increasing the heat radiating surfaces of P with regard to the requirements that it is compact for placement under the fairing of the launch vehicle, and on the other hand, with her inability to adjust the emissivity of the radiator-emitters during operation of SPACECRAFT in various conditions of heat stress. This also limits the possibilities of its application to a wider class of SPACECRAFT.

The purpose of the proposed technical solutions - improving the accuracy and reliability management, expanding application possibilities.

This objective is achieved in that in the method of thermal control of the spacecraft, including the removal of excess heat from each device installed on the conductive sotapanna and, via built-in honeycomb pairs directly connected to each other in the longitudinal direction of the first two and second two evaporator l-shaped adjustable heat pipes, and heat from the condensers of these heat pipes with the specified first and second connected to each other evaporators operate respectively on the first side radiators-radiators and located orthogonal them second lateral radiators-radiators, perform additional removal of excess heat from the inner surfaces of one of the radiators-radiators of the first lateral radiators-radiators and located orthogonal them second lateral radiators-radiators on the orthogonal located on their condensers evaporators U-shaped heat pipes and the removal of excess heat from the capacitors above the U-shaped heat pipes operate respectively on the inner surface of the other radiators-radiators, respectively, of the first lateral radiators-radiators and located orthogonal them second lateral radiators-radiators with mutually orthogonal arrangement of their capacitors relative to the capacitors of the U-shaped heat pipes; thermal control system of the spacecraft, including the instrument container with exterior insulation formed of p is the Union of two U-shaped heat-conducting stopplng blocks, each of which is designed with built-in inner structure with l-shaped parallel-arranged adjustable heat pipes, vaporizers which made pairs directly connected to each other in the longitudinal direction in the middle of setpanel, on the inner side of which is installed devices, and capacitors each specified pair of heat pipes are made respectively in its lateral sotapannas made in the form of lateral radiators-radiators, each U-shaped heat-conducting sotopaneley the unit is made with parallel spaced U-shaped heat pipes, evaporators and condensers which are embedded in the inner side of the additional sotopaneley layers of the respective lateral radiators-radiators and orthogonal position relative to the capacitors of these radiators-radiators.

It is possible to increase the accuracy of the temperature control medium setpanel with devices by reducing the difference of temperatures between the side radiators (emitters) of each U-shaped heat-conducting stopplng unit by providing thermal connection through the U-shaped TT, to improve the reliability of the instruments and the SPACECRAFT as a whole, to extend the application PAGE for KA various internal and external heat loads, and takita account the possibility of installing devices on the lateral radiators-radiators to ensure its operating temperature in a narrower range.

Analysis of the known technical solutions in the area of study allows to draw a conclusion about the absence of signs, similar to the set of features of the proposed solution.

The proposed solution is illustrated by drawings: figure 1 shows the U-shaped heat-conducting sotopaneley blocks (POTS); figure 2 - section of the middle of setpanel; figure 3 is a cut side of the radiator-emitter; figure 4 - instrument container P KA in the collection.

P KA for implementing the method contains the instrument container 1 formed by combining two POTS 2, each of which is designed with built-in outer main layer internal structure 3 (rigidly connected to the parallel metal casings 4), arranged in parallel with l-shaped adjustable heat pipes (GORTT) 5, an evaporator 6 which is made in the middle of setpanel 7 POTS 2 pairs directly connected to each other in the longitudinal direction, and the capacitors 8 each specified pair of heat pipes are made in the outer sotapanna layer 3 side radiators-radiators 9; parallel spaced P-shaped heat pipes (POTT) 10, evaporators 11 and a capacitor 12 which is embedded in the internal additional sotopaneley layers 13 of the respective side radiators-radiators (BRY) 9 and with an orthogonal arrangement is otnositelno their capacitors 8; devices 14 mounted on the inner sides of the BRIE 9 and the average setpanel 7, with the outer side of which is installed insulation 15.

After the withdrawal of the SPACECRAFT into orbit in the instrument container 1 carry out the engagement devices 14. Excess heat generated by the devices 14 mounted on the average setpanel 7, is passed through the aluminum siding 4 (thickness of 0.3-0.5 mm) on the evaporator 6 GORTT 5 and in the process TT excess heat is transferred respectively to the capacitors 8, through plating 4 BRIE 9 excess heat is radiated into free space.

The outer surfaces BRIE 9 performed with the temperature-controlled coating-type solar reflector that provides minimum thermal load on them from direct illumination by the Sun and the maximum heat-radiating ability (respectively AS≤0,43; ε≥0,85).

Depending lit whether the Sun or not BRIE 9, their temperatures will be correspondingly higher or lower with the difference of temperatures between 40°C. as they are in thermal relation with the average setpanel 7 through GORTT 5, despite the fact that GORTT 5 provide an adjustable thermal link, the influence of different temperatures BRIE 9 on average honeycomb 7 will be significant due to the exclusion of the absolute accuracy of regulation, inertia telomere the ACI, degradation of thermal control coatings for extended lifetime AC (10 years and over), unregulated heat transfer design of instrument container 1. But as evaporators neighboring GORTT 5 each POTS 2 pairs directly connected to each other in the longitudinal direction with their capacitors 8, made respectively in BRY 9, it is possible to improve the alignment temperature average setpanel 7 and devices 14 and thereby improve the reliability of their work.

During operation of SPACECRAFT large part of the active lifetime radiator-emitter BRIE 9 each POTS 2 is in a state illuminated by the Sun, and the other in the shade. However, their temperature can make a big difference, for example 10-40°C. But as in the proposed solution, the said radiators-radiators in addition to thermal coupling through GORTT 5 is additionally connected through the POTT 10, the temperature difference between the radiators-radiators each POTS 2 significantly reduced. The ability to reduce this temperature difference in the General case depends on the number of installed POTT 10 on each POTS 2 and their heat transfer capability. The minimum achievable temperature radiators-radiators on each POTS 2 may be provided at the level of 4-6°as the temperature difference between the at the evaporator and the condenser TT during its operation at maximum heat transfer in practice level 3° C. by reducing the temperature difference on the radiators-radiators are improving the control accuracy of temperature as the average setpanel 7 and the inner surfaces BRIE 9, which installed devices 14.

The positive effect is ensured, on the one hand, by equalizing the temperature between the radiators-radiators each POTS 2 by highly efficient thermal coupling through GORTT 5 and POTT 10 that allows GORTT 5 ensure that the average temperature of setpanel in a narrower range. On the other hand, due to the fact that the ranges of the average temperature of GORTT 5 and POTT 10 of about 3-6°already the temperature ranges of the outer surfaces of the radiators-radiators each POTS 2. As a consequence, the temperature of the devices 14 that are installed on the middle sotapannas 7 and the inner surfaces BRIE 9, will be provided in a narrower range compared to the prototype, which ensures their high reliability. Precision thermal control system devices allows you to expand the possibilities of applying the proposed PAGE for KA with more stringent requirements on temperature conditions.

During operation of the AC is alternating maximum and minimum internal and external heat loads on it. Due to the fact that the proposal is hinnon solving highly effective equalization of the temperature fields BRIE 9, average setpanel 7 and devices 14 by the specified application of the TT, the change of the temperature fields of these elements occurs in a narrower range due to the high thermal mass communication of the heat capacities of the individual elements into a single more inertial mass heat capacity, providing change its temperature in a narrow range in a changing thermal loads on the SPACECRAFT. Which ultimately improves the accuracy of temperature control devices, and CA in General and thereby increases the reliability of their work.

Use insulation 15 on the outer surface average setpanel 7 allows you to minimize the unregulated heat exchange AC with the environment and thereby contribute to providing a more accurate temperature control.

The working of the body, which filled GORTT 5 and POTT 10 - ammonia (coolant). Governing body for filling of GORTT 5 - nitrogen. Dose of their dressings are made with regard to design dimensions of the TT set temperature conditions of their work in the operation and requirements provide the temperatures of the evaporators 6 and 11. The work of GORTT 5 is as follows. The maximum heat load on the evaporator TT vapour pressure of ammonia increases, and as the movement of steam is supplied from the evaporator to the condenser, nitrogen (seconden is ruusila gas) replaced vaporous ammonia to the end of the condenser or in a specially made for nitrogen capacity, when this condensation of ammonia vapor occurs throughout the volume of the capacitor and TT runs with maximum heat transfer. Mode minimum heat load on the evaporator TT vapour pressure of ammonia is reduced, while non-condensable gas expands and displaces vapors of ammonia from the condenser and thereby eliminates the possibility of condensation of ammonia vapors in the condenser. This cuts the heat transfer heat from the evaporator to the condenser TT due to evaporation-condensation effect and thus thermal interchange between the average setpanel 7 and BRIE 9.

The POTT 10 provides heat transfer from the evaporator to the condenser (or from the warmer area to a less heated in any part of it) by vapor-liquid circulation of the coolant inside her tight cavity. When exposed to heat liquid in the evaporator evaporates and moves into the zone of lowest pressure in the condenser where it is cooled, condensed and due to the wettability of the capillary structure on the inner surface of the wall of the TT moves to the evaporator, thus providing a closed heat transfer duty cycle TT.

Instrument container 1 is made of modular type, which allows its arrangement with one or more other similarly completed the ies instrument containers with providing thermal communication between them without performing any complex design of their revisions. This allows to extend the applicability of the proposed PAGE for KA of different types and layouts (for example, with the two connected containers in orbit, or to run one rocket-carrier multiple SPACECRAFT and so on).

This solution is currently undergoing a study on the conditions and the optimal range of applications.

1. The method of thermal control of the spacecraft, including the removal of excess heat from each device installed on the conductive setpanel, through built-in honeycomb and pairs directly connected to each other in the longitudinal direction of the first two and second two evaporator l-shaped adjustable heat pipes, and heat from the condensers of these heat pipes respectively on the first side radiators-radiators and located orthogonal them second lateral radiators-radiators, wherein implementing the additional removal of excess heat from the inner surfaces of one of the radiators-radiators of the first lateral radiators-radiators and located orthogonal to them second lateral radiators-radiators - for evaporators of the U-shaped heat pipes located orthogonal to the capacitors l-shaped adjustable heat pipes that transfer heat to the specified one R is dietary-emitters, and the removal of excess heat from the capacitors above the U-shaped heat pipes operate on the inner surface of the other radiators-radiators, respectively, from among the first lateral radiators-radiators and located orthogonal them second lateral radiators-radiators and condensers data of the U-shaped heat pipes are orthogonal capacitors l-shaped adjustable heat pipes that transfer heat to these other radiators-radiators.

2. thermal control system of the spacecraft, including the instrument container with exterior insulation formed by combining two U-shaped heat-conducting stopplng blocks, each of which is designed with built-in inner structure with l-shaped parallel-arranged adjustable heat pipes, vaporizers which made pairs directly connected to each other in the longitudinal direction in the middle of setpanel, on the inner side of which is installed devices, and capacitors each specified pair of heat pipes are made respectively in the side sotapannas a U-shaped block, which is the lateral radiators (emitters, wherein each U-shaped heat-conducting sotopaneley the unit is made with parallel-arranged P-abranimations pipes, evaporators and condensers which are embedded in the inner side of the additional sotopaneley layers of the respective lateral radiators-radiators and are orthogonal relative to the capacitors of these l-shaped heat pipes made in these radiators-radiators.



 

Same patents:

FIELD: temperature control of surrounding medium parameters on board manned space object.

SUBSTANCE: proposed device includes absolute pressure gauge, vacuum pump and hermetic reservoir divided into liquid and gas chambers by means of elastic diaphragm. Liquid chamber is filled with working medium and is connected with section of hydraulic line being checked by means of pipe line through first shut-off valve. Gas chamber filled with air at pressure of space object compartment atmosphere is connected with pressure gauge. It is also communicated with vacuum pump inlet and with said compartment via second and third shut-off valves. Vacuum pump outlet is communicated with compartment atmosphere. Pressure and temperature of working medium in hydraulic line are equalized with pressure and temperature of compartment atmosphere. Then, hydraulic line is divided into separate sections which are not connected hydraulically and each section is brought in turn in communication with liquid chamber. Every time after communication with liquid chamber, it is disconnected from section being checked and gas chamber is evacuated till pressure of saturated vapor of working medium corresponds to its temperature. Then, liquid chamber is brought in communication with gas chamber at maintenance of constant temperature and change of air pressure in gas chamber is checked. In case there is no change in pressure, section is considered to be perfectly hermetic and vice versa in case of change in pressure, section is rejected.

EFFECT: enhanced safety for space object crew in finding fault in hydraulic line sections.

3 cl, 1 dwg

FIELD: spacecraft temperature control systems.

SUBSTANCE: proposed method consists in maintenance of temperature in zone where instruments are mounted within preset limits and measurement of rates of change of temperatures. Electric heaters of thermal tubes are switched on or switched off depending on these rates, thus maintaining the tolerable magnitudes of these rates. Prior to switching-on the instruments, total thermal energy released by these instruments and total thermal energy released by electric heaters shall be determined. These magnitudes are compared and electric heaters may be switched off or delivery of thermal energy supplied to area of installation of instruments shall be changed by definite difference. When high limits of temperature are achieved in these zones, electric heaters are switched off till rates in temperature change get below permissible limits. Electric heaters in adjacent zones are switched off similarly but for temperature effect disturbing thermostatting of adjacent zones by temperature low limits. For subsequent interval of joint operation of instruments, temperature is changed with the aid of electric heaters. At change of external heat fluxes acting on panel, modes of switching on/off the electric heaters are changed for new distribution of temperature changes in zones. Operations similar to above-mentioned operations are performed at constant operation of instruments and in case of connection of other instruments. After complete or partial disconnection of instruments, cycle of change of temperatures is performed by the above-mentioned method depending on number of instruments working in zones.

EFFECT: increased service life of spacecraft due to avoidance of adverse cyclic action of temperature.

4 dwg

FIELD: spacecraft temperature control systems.

SUBSTANCE: proposed method consists in equalizing the rates of change of temperatures in instrument panel zones at the moment of action of external heat fluxes with the instruments switched off. Temperature is maintained at constant level within permissible range by switching on/off the electric heaters. Prior to simultaneous switching-on of instruments at intervals of their joint operation, total thermal energy released by these instruments and electric heaters located in zones of these instruments is determined. Depending on difference in energy, electric heaters are either switched-off or heat delivered by them is reduced. When instruments are disconnected and temperature high limit is achieved, electric heaters of adjacent zones are switched-off, thus ensuring thermostatting of adjacent zones by low limit temperatures. Similar actions are carried out for subsequent intervals of simultaneous operation of instruments and at constant operation of instruments. In case of alternate switching-on of instruments simultaneously with constantly operating instruments, magnitude of thermal energy released by first ones is determined and heat from electric heaters is delivered to them at intervals of their switching-off. When temperature high limit is achieved, electric heaters are switched-off at intermediate intervals of instrument operation; electric heaters of instruments found in adjacent zones may be switched-off if necessary. Thermostatting of these zones is ensured.

EFFECT: increased service life of spacecraft due to avoidance of adverse cyclic action of temperatures.

4 dwg

FIELD: spacecraft temperature control systems; spacecraft ground servicing.

SUBSTANCE: proposed device has drainage and filling tanks, filling line and spillage line which connect drainage tank with system being filled. Tanks are connected with vacuum unit and with compressed gas source by means of lines which are provided with valves and pressure recorders. Filling line is also provided with pump and valves disconnecting the inlet branch pipe of pump from filling and drainage tanks and outlet branch pipe of pump from filling pipe union of system being filled. Vacuum unit is connected with filling and spillage lines by means of additional lines which are provided with valves. Area of filling tank above heat-transfer agent surface may be connected with filling line via valve. Device may be provided with reservoir whose liquid cavity is connected with lower part of filling tank through valve and temperature and pressure recorders; gas cavity is connected with compressed gas source and vacuum unit through pressure recorder and valves. Liquid cavity may be also connected with filling line by means of valve behind the valve disconnecting the pump from filling pipe union. Pump inlet may be connected via valve with definite point of spillage line.

EFFECT: improved quality of filling procedure due to enhanced monitoring of gas content in filling line and high degree of evacuation of system at heat-transfer agent de-aeration.

5 cl, 1 dwg

FIELD: space engineering; satellite temperature control systems.

SUBSTANCE: proposed system has hermetic instrument container filled with heat-transfer agent and made in form of envelope of body of revolution, onboard equipment units located inside container, fan placed in casing and two radiators mounted on different sides of container. Radiator inlets are connected with container interior and outlets are connected with fan casing inlet by means of gas ducts. Container is made in form of cylinder; its axis is perpendicular to satellite axis. Radiators of temperature control system are mounted in parallel relative to cylinder butts. Gas ducts are laid inside instrument bay. Satellite temperature control system is provided with two temperature sensors, control unit and multi-position gate valves. Two gate valves are mounted at butt of radiator and gas duct outlets. Two other gate valves are mounted in gas duct walls. Temperature sensor outputs are connected with control unit inputs and control unit output are connected with gate valve inputs. Temperature sensors are mounted on two heat shields which are thermally insulated from radiators. Thermal physical characteristics of heat shields are identical to those of radiators; their planes are matched with radiator planes.

EFFECT: enhanced accuracy of temperature control inside instrument bay.

2 cl, 1 dwg

FIELD: thermal control systems, primarily for telecommunication satellites.

SUBSTANCE: claimed device contains closed two-phased contour, filled with low boiling point heat carrier (ammonia). The contour includes capacitor with input and output, evaporators, having common input and common output, and hydro-accumulator, installed closely to common input of evaporators, all interconnected by pipelines. Hydro-accumulator has body, which has mutually opposite input and output, connected respectively to output of capacitor and to common input of evaporators, mesh fuse, positioned inside the body near its walls, and electric heater. Capacitor is attached to covers of radiator. Evaporators, in body of which capillary pumps are installed, are connected to thermostatted panel and installed hydraulically in parallel. After common output of evaporators, direct action pressure regulator is additionally installed, connected by its input to common output of evaporators, and by its output - to input of capacitor. Inside the body of hydro-accumulator to its input a branch pipe is connected with output aperture close to fuse zone, positioned near the output. Aforementioned branch pipe has additional contact apertures with fuse in its zone, positioned near the input. Electric heater is installed on external surface of hydro-accumulator near its input. In the zone positioned near the output the fuse is made with smaller cells compared to its remaining part.

EFFECT: ensured reliability of device in all modes of operation (heat emission) of devices under conditions of satellite operation.

3 dwg

FIELD: thermal control systems, primarily for telecommunication satellites.

SUBSTANCE: claimed device contains closed two-phased contour, filled with low boiling point heat carrier (ammonia). Contour includes evaporator, hydro-accumulator and capacitor, attached to case of radiator, all interconnected by connecting pipelines. Evaporator has thermal connection to two oppositely positioned thermostatted surfaces. In the body of evaporator, capillary pump is installed, to external surface of which contact surfaces of evaporator body shelves are closely adjacent. Evaporator and capacitor are built into cell panel of satellite with two thermostatted cases and made in form of profiled pipeline of aluminum alloy, each with two shelves. Shelves are attached (in case of evaporator - glued) to cases of panel. Internal surface of profiled pipeline in zones, positioned in front of shelves, has shelves with longitudinal and transverse grooves. In other zones recesses are made. Capillary pump has bimetallic flange, which is soldered to external surface of pump and welded to ends of aforementioned profile and connecting pipelines.

EFFECT: reduced mass of device (by 7-15%) in comparison to analogical technical solutions, reduced hazard of penetration of mains of evaporator and radiator by meteorites under conditions of operation.

8 dwg

FIELD: space engineering; manufacture of thermostatted honeycomb panels with built-in liquid manifold.

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EFFECT: enhanced reliability of manufacture of liquid manifolds.

3 cl, 2 dwg

FIELD: space engineering; maintenance of temperature conditions of entire spacecraft and its components.

SUBSTANCE: proposed method includes measurement of temperature in areas of panel with the aid of temperature sensors and maintenance of required temperature within tolerable range. Temperature is changed at time intervals determined by orientation of spacecraft relative to Sun and planets. Every time after change of temperature, measurements shall be made in this area and in adjacent areas and difference in magnitudes shall be noted. Temperature dependences between zones shall be determined taking into account heat delivered from adjacent components. Further temperature control is carried out with dependences taken into account. In case present temperature is beyond permissible limits, it shall be increased or decreased by control of heat delivered to zones from adjacent components. In case of failure of temperature control system components, heat is delivered from adjacent zones with serviceable components adhering to the conditions under which temperature dependences were determined beforehand.

EFFECT: enhanced reliability of system due to experimental determination of temperature dependence.

5 dwg

FIELD: spacecraft temperature control systems; ground testing and servicing of spacecraft hydraulic systems.

SUBSTANCE: volume of filled line is localized before draining compensating dose of heat-transfer agent at isolating the hydraulic line from filling means. Then, pressure is measured in hydraulic line and atmospheric pressure is set twice in compensator gas cavity of temperature control system; different initial basic pressures are set in reference reservoir in turn. Said cavity and reservoir are brought in communication twice and steady-state averaged pressures are measured. Volume of undissolved gas is determined by respective formula on basis of measurements. Initial basic pressures in reference reservoir satisfy definite limitations in form of inequalities.

EFFECT: improved quality of filling due to enhanced accuracy of determination of undissolved gas volume and more accurate determination of volume of drained compensating dose of heat-transfer agent.

3 cl, 1 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.

5 dwg

Nanosputnik // 2308401

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.

5 dwg

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.

3 dwg

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.

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: 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

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