Method of control of temperature of spacecraft equipped with solar batteries

FIELD: control of temperature of spacecraft and their components.

SUBSTANCE: proposed method includes measurement of temperatures in spacecraft temperature control zones, comparison of these temperatures with high and low permissible magnitudes and delivery of heat to said zones at low limits. Heat is delivered by conversion of electrical energy into thermal energy. Power requirements are measured at different standard time intervals of spacecraft flight forecasting orientation of its solar batteries to Sun. Magnitude of electric power generated by solar batteries is determined by forecast results. Measured magnitudes of consumed electric power are compared with forecast data. According to results obtained in comparison, flight time is divided into sections at excess of energy generated by solar batteries over consumed power, equality of these magnitudes and shortage of generated energy. High magnitudes of temperature are maintained at excess energy sections by conversion of difference of generated energy and consumed energy into heat. In case of reduction of generated energy in the course of changing the orientation of solar batteries on Sun, temperature in these zones is reduced to low limits at simultaneous equality of energies. In case of further increase of generated energy, temperature in said zones is increased to high limits at equality of energies. Then, in the course of change of generated energy, temperature correction cycles in temperature control zones are repeated.

EFFECT: avoidance of excess of consumed energy above generated energy of solar batteries.

7 dwg

 

The invention relates to space technology and is intended for use on space vehicles (SV) in space, where it is necessary to maintain a preset temperature, just as the KA and its individual elements.

Known methods of passive thermal control radiation surfaces [1], which are implemented through the use of materials with specific thermal characteristics (radiation and thermal insulation), by choosing the appropriate geometric shape of the device and its orientation relative to the Sun and through the use of heat of phase transitions.

The main elements of thermal control systems (P)built using these methods are thermal control coatings and high-efficiency screen-vacuum thermal insulation.

Passive temperature control with the use of surfaces with specific radiation characteristics and highly efficient thermal insulation reduces heat flows inside the AC (or heat loss to space) and reduce the heat load on the operating system. The main criteria in the selection of external thermal control coatings for practical use on the SPACECRAFT are the coefficient of absorption of solar radiation As the degree cher is the notes E, as well as the stability of these characteristics after a long stay in space under the influence of UV radiation from the Sun and components of space radiation.

The most efficient insulation in the space flight environment is multi-layered insulation recruited from radiation screens and insulating spacers. As radiation shields are used, for example, metallic aluminum (less gold) Mylar (for temperatures not exceeding 160° (C) or kartonowy (for temperatures up to 430° (C) films [2].

A typical example of a passive thermal control for SPACECRAFT modular design presented in [3].

During the flight the specified SPACECRAFT orientation relative to the Sun regularly. When one of its units H-shaped configuration with radiant panels "East" and "West" is alternately periodic daily repeated exposure to solar radiation.

By drawing on radiation setpanel coating type optical solar reflector" (CCA) to ensure throughout the specified life time (not less than 10 years) AS/E<0,43 expected to provide the radiation panels temperature not more than the maximum allowable value 45°C. the Specified upper allowable value is emperatur chosen with consideration of the temperature modes of the devices, installed on the panels.

In addition, the radiation sotapannas block H-shaped configuration, set the drop-down and compact folding mechanical drives on the edges or in the middle of the panels heat roller blind. Roller blind made of layer-by-layer combo layered Mat screen-vacuum thermal insulation. With these blinds, you can:

to reduce external heat flow inside the device by closing their panels when flying SPACECRAFT into the lighted area of the orbit;

to reduce heat loss into space, holding the panel closed shutters when flying SPACECRAFT on an unlit stretch of the orbit;

to carry out temperature control of the radiation panels by measuring the temperature in different zones of the panels, comparing the measured temperatures with the upper and lower values to limits and opening-closing of the shutters to maintain the temperature at the radiation surfaces between the upper and lower limit values by periodically supplying external heat flows to sotapannas.

Passive methods of thermal radiation surfaces KA have certain advantages over active. Systems and devices that implement passive ways, more reliable in operation, their design generally has lower mass.

However, active methods t is samoregulirovania as internal compartments KA, and their surfaces can maintain the necessary heat mode when the external and internal heat loads in a wide range. Moreover, the accuracy of temperature is much higher than in systems that implement passive methods of thermal control [4].

Therefore, active thermal control are used to ensure that thermal regime quarters manned SPACECRAFT, as well as for temperature control instrument compartments with complex and sensitive electronic equipment.

Using active methods of thermal control systems are used with the circulation of the refrigerant, changing thermal resistance between the internal volume of the compartments and their shell), heaters and thermostats, bimetallic actuators for blind control, thermostatic and other devices.

In P with forced circulation of the liquid (or gas) in the closed circuits of the heat from the cooling source is transmitted to the fluid, which is then cooled by radiation surfaces, discharging heat radiation into space. Most often for these purposes as radiation surfaces are used radiators capacitors. Moreover, these radiators capacitors can be implemented as a liquid, and on the basis of heat pipes (TT).

System with TT more effectively is effective in thermal relation, more reliable and have less weight in comparison with similar systems without heat pipes because the TT compared to the commonly used elements of P (heat exchangers, pumps, etc. have a number of advantages:

does not require energy for pumping coolant;

pipe more reliable and silent in the absence of moving parts;

no additional regulating devices as may be used for self-regulating pipe;

radiator panel using heat pipes is more reliable (less vulnerability radiator when the meteor hit);

able to provide high thermal conductivity between the heat sources and sinks that enables you to use less of the surface and, therefore, reduce the weight.

Using TT created various devices that determine the radiation surface of the SPACECRAFT, such as radiation panel placed on their devices (see, for example [3]), cooled TT bearings with radiation surfaces in the form of boots on the Shuttle SPACECRAFT [5] and others

Normal TT variable thermal conductivity able to maintain its temperature at a constant level, despite the fact that the applied heat capacity and the surrounding conditions change. If thermal resistance between the TT and the heat source is small, so the temperature value of the source will also be approximately constant.

In practice, this resistance is often quite large, resulting in the source temperature will vary over a wider range than the temperature of the TT.

These temperature fluctuations can be significantly reduced when controlling for feedback. The most commonly used method of control by feedback - electric. Regulatory systems with electric feedback include a thermistor, electronic control unit and the heater.

Thus, there is in space technology with a wide range of designs radiation surfaces using active methods of control. As a rule, for the specified control is used on-Board electrical energy devices. Electricity costs associated with the supply of heat to the radiation surfaces to maintain the temperature of operation of the devices installed on radiation panels, to prevent freezing of the liquid coolant in the contours of radiators condensers and other cases.

As a prototype of the invention it is proposed to elect the method of thermal control of various heated radiation surfaces and structural elements of the SPACECRAFT, which is necessary to make the cost of on-Board electric energy to the AI KA, generated by solar batteries (SB), see [3]. According to the classification in P this method applies to active methods and systems for SPACECRAFT thermal control.

The essence of the method lies in the fact that thermal zones KA measure temperature, then make a comparison of the measured temperatures from the upper and lower values of their permissible limits. When the temperature limit lower values provide a supply of heat to thermal zones KA by conversion of electrical energy into thermal energy. Application of heat is terminated by reaching the measured temperatures of the upper limit values.

The disadvantage of the prototype method is that at times the expenditure of electrical energy for heating KA ignored orientation SAT in the Sun.

When generated on-Board electricity than consumed, including consumable and specified on heating, the situation is not critical. The lifespan of the orbit is determined by the rated lifetime (years).

In the case of on-Board consumption from the secondary power source (batteries) the lifespan of the orbit is determined by the backup time interval from the point in time that defines the current supply of the electric power in the battery the arts (AK) until complete consumption of on-Board energy consumption.

If the lifespan of the orbit is determined by the years, the standby time is determined by the clock.

Given the fact that the P KA consumed about 1/4 of the total on-Board energy consumption, it is important task to increase the backup time KA by eliminating or reducing the load on the onboard power supply system (PSS) in critical situations when the transition SES to work from secondary sources of electricity - by eliminating or reducing the heat radiation surfaces and structural elements of the SPACECRAFT. While the current values of the temperature of the device must be within a permissible nominal limits.

In addition, the connection of the secondary power source to the electrical load will result in the need for additional holding charge-discharge cycles. These cycles are among the main resource settings and as they develop, as a rule, decreases the value of the bit energy AK.

The present invention aims to eliminate the possible connections AK power under load when changes in orientation SAT in the Sun, leading to excess consumption of electricity generated over SB KA.

This technical result is achieved by a method for thermal control of satellites with the B, including the measurement of temperatures in thermal zones KA, comparing the measured temperature with the upper and lower values to limits and application of heat to these areas through the conversion of electrical energy into thermal energy when reaching the measured temperatures lower limit values, and to reach the specified temperature upper limit value during the current flight SPACECRAFT measured the power consumption at different sample intervals of flight time, then before running the AC flight program specified duration smash it on the standard intervals of flight time, with previously measured values predict energy consumption forecast for the same intervals of flight time orientation on SAT The sun, according to the results of the forecast orientations SAT in the Sun and previously measured values of the generated power for these various orientations determine the values of the generated power at typical intervals, compare the predicted values of the power consumption with defined according to the forecast values of the generated electricity on the same sample intervals of flight time, on the comparison of the share of flight time KA specified duration to phase the intervals, for which the generated electricity has exceeded consumption, equal consumed and less consumption of electricity, next to plots generated electricity, less than consumption, determine the area of thermal control of SPACECRAFT, which produce temperature by turning electricity into heat energy, and for these zones during the flight SPACECRAFT in the previous sections, where the generated electric power consumption, maintain the temperature at the upper levels until it reaches the upper limit value by supply of thermal energy by converting the difference of these energies into heat energy, then in the process of flying a SPACECRAFT to the predicted sites, as the amount of generated electric energy when changing the current orientation SAT in the Sun, adjust temperature by reduction in these areas until the lower limit values and achieve this equality generated and consumed electricity, then, as the increase of the generated electricity in the process of changing orientation SAT in the Sun and exceeding its electricity consumption, adjust the temperature in thermal zones, starting with the lowest values, by their increase until it reaches the upper observe the Sabbath. 's values and achievements of this equality generated and consumed power plant fueled and as changes in the value of the generated electricity repeat cycles correction of the temperature in thermal zones KA the above by the way, and after the completion of the current mission program specified duration and before the next flight, make another breakdown on some of the above, the areas used for the prediction of the measured values consumed and generated power on the model performed intervals flight, and then repeat the cycle thermal control of SPACECRAFT in this way.

To explain the essence of the present invention are represented figure 1-7.

Figure 1 is a conditionally shows a diagram of one of the radiation panels of the SPACECRAFT.

Figure 2 shows a graph of daily values of load current KA (In) prior to the accumulation of thermal energy in thermal zones of the SPACECRAFT.

Figure 3 shows the daily schedule of the orientation angles SAT in the Sun.

Figure 4 shows a graph of current (ISAT)generated by SB KA during the day in accordance with the daily schedule of orientation angles SAT in the Sun.

Figure 5 shows a graph of the temperature in the i-x zones panel (Tziin the process of accumulation of thermal energy.

Figure 6 shows a graph of daily values of load current AC in the accumulation of thermal energy in thermal zones of the SPACECRAFT.

7 shows a graph of the temperatures in the zones brackets installation of stationary plasma thrusters (SPT).

Figure 1 before the taulani:

1 - radiation panel, i-mi zones, where i=1...8;

2ithe instruments installed on the panel in the i-x zones;

3i- TT installed on the panel in the i-x zones;

4i- electric heaters (EN) installed on the TT i-x zones;

5i- temperature sensor (TD), installed on the panel in the i-x zones;

6 is a cross TT.

For the description of the invention will take one of the most typical cases of SPACECRAFT thermal control - radiation control panels. These panels are widely used on the SPACECRAFT with solar-terrestrial oriented circular and highly-elliptical orbits.

Radiation surface in the form of sotapanna and are both radiators "South" and "North" of the parties KA (see, for example [3])Rassmotrim temperature control one of the radiation sotapanna geostationary satellite, see figure 1.

As can be seen from figure 1, the temperature of each of the i-th zone is provided by two TTi, 3icontrol of temperature in the area is three TDi, 5ion each TTiinstalled one at a time-N, 4i. The equalization of temperature between the zones is carried out using cross-TT 6. Control of thermal conditions in the i-x zones panel is to compare the actually obtained values testimony DTi, 3iwith a given temperature di is the scoring range for each of them.

To control the operation of the CN from the onboard control circuit SC (see, for example, [7]) provides commands to enable and disable groups of EN on-Board digital computer system (BCS) according to the indications DTi, 5iinstalled in i-x, i=1...8 zones panel. Management of the CN zone is BCS average temperature parameter sensors temperature zones.

Algorithm P automation controls the operation of the CN zone heating in accordance with the following logic:

- in the original arrays temperatures on each step of the survey revealed inoperative and locked the temperature sensors, which are excluded from the analysis to generate control commands EN;

for the algorithm for each zone heating the parameters of the control channel EN - nominal temperatures to enable the CN, as well as the lower and upper limit values for thermal control.

Main EN included in the achievement of the measured temperature value of specified nominal value. Next is heated zone to the upper limit temperature (maximum values), and off the CN. Re-enable main EN are produced by lowering the temperature in the area to the nominal value. In the case when the main EN not obespechivayuschee the upper limit of the temperature values and the temperature is lowered in areas to the lower limit (minimum) values, extras include backup EN. Off the main and backup EN is at achievement in the areas of the upper limit of the temperature values.

For further explanation will choose a typical day of the flight SPACECRAFT geostationary communications (GSS).

The measured values of the load current InGSS during the day are presented in figure 2. The current composition includes components from the consumption of the payload (MO), on-Board equipment (BA) and the CN. Component characterizing the work of MO, almost close to a constant value (for example, through continuous trunks enable the onboard repeater). Variables the values of Incontains due to the inclusion of BA and ES.

On-Board power consumption is compensated by the current generated by SB (ISAT), which, in turn, is a function of the angles between the normal to the working surface of the battery and the direction to the Sun. For two rotary SB (see[3]) SATELLITES that are constantly orbiting the orientation specified current is a function of the rotation angle α1and α2the first and second batteries, respectively, ISAT=f(α1that α2). When the excess generated electricity on Board the SPACECRAFT over consumption, which is always the case for early flight KA, SA can perform other functions on Board. For example, the R, may play a role propulsion for the GSS or the Executive body, creating a discharge points for the system installed on Board the SPACECRAFT power gyroscopes (see [8, 9]). In the first case uses the force of radiation pressure on the surface of the SAT for orbit correction ("solar sail"), and the second, those same forces, supplemented by others (e.g., forces from the interaction of the intrinsic magnetic moment SAT with the Earth's magnetic field), create a control point, CA.

In each of these cases, the effect is achieved by the lapel SB from the direction to the Sun. For example, consider the case of formation on the SPACECRAFT using SAT and control forces and moments that can not only provide the specified speed change characteristic, but also the management of the kinetic moment [8].

In accordance with the logic inherent in [8], is calculated forecast orientation SAT in the Sun for the next day's flight.

The results of this prediction are shown in figure 3, where the corners α1and α2presents a number of rotary zones each of SA (one of the readings is the angle of rotation ˜2,8°). For the specified mode is also necessary relining SAT" on the corners α1and α2(turns from an initial position on the corners ˜90°).

According to the results of the forecast position SB-defined values of the generating system is constituent of their current I SATin the interval the upcoming days (smfg). The determination can be made by calculation (see [10], p.109), and on the basis of the forecast on the measured values of the specified current, obtained, for example, in the previous day of the flight GSS.

Compare current at a constant value of voltage ˜28VDC measured values of energy consumption with its projected values generated by SB.

As can be seen from figure 2 and figure 4 in areas of flight time ˜16:00 to ˜16:30 and ˜18:30 to ˜19:30 current day, the arrival of electric power may be less than consumption, which, in turn, will lead to the violation of the energy balance and will require the inclusion of AK. In addition, when conducting the specified comparative evaluation, it is necessary to bear in mind an additional supply of electricity on a "different kind of uncertainty." Under these "uncertainties" are understood to be more unpredictable enable BA, as well as the maximum possible fluctuations of the load current. For example, activation of BA during emergency situations in the service systems and MO can lead to the increase of In. Evaluation for the specified CA, will require the specified "supply current"equal to not less than 4 A.

Given this stock under the limiting current are lots of flight time with ˜10:45 to ˜11:30 and ˜13:45 what about the ˜ 14:30.

In these days you can also define areas of approximate equality (including stock) consumption BA, MO, EN electricity and its generation using SAT (˜16:30 to ˜18:30), and areas with excess generated electricity consumed over (all unquoted earlier sections in the specified interval of flight time). In all of these areas can be periodic inclusion of EN in i-x zones panel (see figure 1).

For example, the first two plots (with ˜10:45 to ˜11:30 and ˜13:45 to ˜14:30) is the inclusion of N in zones 1-5 (see figure 1). The valid range control limit values is from 0°to 35°C. it is Suggested to exclude the N on the interval of flight time with ˜10:45 to ˜14:30, thereby to reduce the load average ˜4 A.

With this purpose until the time of the beginning of the specified interval produced by adjusting the operation of specified EN in the first to fifth areas, for example, nominal values 20°and the width of the area to work ±5°C. Thus, the heaters will operate at temperatures in the range of 15°...25°C. Further, taking into account the gradient of increasing temperatures in these zones, choose the time of their inclusion for ˜4 hours prior to entrance into the specified flight interval. Include them in the set time (˜0:00 on the given interval) and produced by the accumulation of heat in these areas. In ˜09:15 produced by the reconfiguration of work of the same ES the nominal value of 5°and the width of the area to work ±5°C. Since the new setup 5°With less than 20°With the previous settings, all EN off and will stay in this position until reaching the nominal temperature lower settings 5° (enable EN is manufactured at par settings).

Experimental graphs of temperatures Tziin these zones and the load current Inshown respectively in figure 5 and 6. As can be seen from figure 5, in the process of heating increased temperatures in zones ˜4°...˜10°s to ˜17°...˜22°C. Increase in temperature with the increase of the load current Insee Fig.6. As can be seen from the graphs in figure 5, the inclusion of EN will happen much later time 19:30, i.e. in the interval energonapryazhennosti" section.

To reduce the current load on the interval from 16:00 to 16:30, 14:30 included two EN installed in areas brackets SPD. The valid range of temperature for the specified design elements from 0°to 35°C. Similarly is the accumulation of heat for follow-up "energonapryazhennosti" section.

Warming up the AC in thermal zones, in these examples was provided with the necessary sufficiency is to solve the task of unlocking "energonaprjazhenie areas and to prevent the inclusion of AK.

In case of different conditions orientation swivel SAT in the Sun, setting P KA must be made differently. For other installations, SAT (increase of their quantity, change the operating conditions, the lack of a systems orientation SAT etc) on the SPACECRAFT, as well as looking at other flight conditions KA (the presence of "shadow" areas of the orbit, different orientations, and so on) must take into account all the factors, influencing the advent of electricity from SB, and N P KA.

In General, you should available a positive balance of electric energy generated by SB, to turn into heat energy with the aim of further reducing the demand for electricity to control the AC. Next, you need to ensure equality between generated and consumed power by the removal of thermal energy by maintaining the temperature in thermal zones on the lower limit values.

Upon receipt of a positive balance between generated and consumed power again to turn electricity into heat form of energy. It is necessary, first, to raise the temperature in the zones with the lowest values of temperature.

This strategy allows for a more evenly to provide temperature control of SPACECRAFT in General, avoiding excessive "Sahelian" googeling areas and structural elements. In turn, this helps to ensure a more even work PAGE as a whole, as well as align the load current of the AC. The total capacity of the heaters PAGE (for this example) more than two times may be higher power SAT. It is therefore necessary to distribute the work (NBC during the day, without creating a local overloads SES from their work. This management strategy allows you to do it.

So, in the example considered, as can be seen from Fig.7, ˜16:30 has been the inclusion of EN on one of the brackets of the state Duma. To carry out the specified action "allowed" current ISAT, who by this time had increased by more than 5A (see figure 4). In principle, the range of temperature control on the specified design element was allowed to lower the temperature to the lowest level (0°). Subsequent flight interval did not require further heat storage to achieve the above purposes, therefore, to the upper limit values of the temperature re-heating to produce in this case was inappropriate. It should be noted that the principle of reasonable sufficiency saving resource of EN - the number of inclusions and duration of their work.

Cycles of inlet-exhaust thermal energy in thermal zones are repeated as described above. Each time proizvoditsa their clarification on the forecast for the angular orientation SAT in the Sun. Gradients of rising temperatures in the zones and structural elements KA with a sufficient degree of accuracy predicted by the results of flight operation of the apparatus. This is used, for example, analysis of SPACECRAFT telemetry data received from the AP.

Rated power AES for prediction of power consumption from their inclusion is also known in advance.

Positive main effect of the proposed method lies primarily in the resource conservation electrochemical AK on the number of possible charge-discharge cycles. It is known that the dynamics of degradation of these batteries is largely determined by the number of specified cycles prescribed in the technical specifications for operation.

As can be seen from this example, using the proposed solution will be able to avoid at least two short charge-discharge cycles per day. In the year this figure may be too high a value that will not allow you to use SB to control the movement of the SPACECRAFT. As a result, for performing dynamic operations will require additional consumption of the working fluid.

Moreover, the mode of operation of the SPP, with lapel SB from the direction to the Sun, the whole system more favorable. The control equipment and monitoring SES have smaller amounts to in turn heat vostrebovannuyu generated electricity, that, in turn, leads to lower thermal loads on the instrument composition of the SES and other benefits in the system.

Thus, in the overall strategy of the SPACECRAFT control using the proposed method allows for flexible reallocation of resources available inside the SPACECRAFT, and also to make their rational use, which ultimately increases the life of the AC as a whole.

Literature

1. Spacecraft, edited by Ceptionist. M: Voenizdat, 1983, s.

2. Two-dimensional thermal conductivity of the multilayer thermal insulation. Overview of GONTI-4. 1972.

3. Iaashraf and other spacecraft modular design. RF patent 2092398, class B 64 G 1/10.

4. Asselineau. Technology of space flight. M: mechanical engineering, 1983.

5. Heat Pipe Application for the space Shuttle, AIAA Paper, 1972, No. 272.

6. Khachaturov and other thermal control System. RF patent 2168690, class F 28 D 15/02, dated 25.08.1999.

7. System for thermal regime CA. GC. U 0000 A-0. RSC Energia. Korolev, 2002.

8. Avigate and other methods of forming the control effects on spacecraft power gyroscopes and rotating solar panels. RF patent 2207969, 08.05.2001.

9. Avigate and other Method of generating control torques on the spacecraft with power gyroscopes and tilt with the solar panels and system for its implementation. RF patent 2196710, dated 28.02.2001.

10. Vagaries, Low, Lbook. Solar and space flight. M.: Nauka, 1984.

The method of thermal control of space vehicles (SV) with solar panels, including the measurement of temperatures in thermal zones KA, comparing the measured temperature with the upper and lower values to limits and application of heat to these areas through the conversion of electrical energy into thermal energy when reaching the measured temperatures lower limit values, and to reach the specified temperature upper limit value, wherein in the process the current flight SPACECRAFT measured the power consumption at different sample intervals of flight time before you run the AC flight, given the duration of the break this program on typical intervals of flight time, with previously measured values predict the power consumption and orientation of the solar panels in the Sun, according to the forecast this orientation of solar panels and previously measured values of the generated power for different orientations of solar panels define the values of the generated power at typical intervals specified split, compare the predicted values consumed elec is renergie with defined according to the forecast values of the generated electricity on the same sample intervals, by comparing the results share the time specified flight program KA at sites with intervals, for which the generated electricity has exceeded consumption, equal consumption and less power consumption, then for areas with generated electricity, less than consumption, determine the area of thermal control of SPACECRAFT, which produce temperature by turning electricity into heat energy, and for these zones during the flight SPACECRAFT in the previous sections, where the generated electric power consumption, maintain the temperature at the upper levels until it reaches the upper limit value by supply of thermal energy by converting the difference of these energies into heat energy, then during the flight the SPACECRAFT at the predicted sites as the amount of generated electric energy when changing the current orientation of solar panels on the Sun adjust temperature by reduction in these areas until the lower limit values and achieve this equality generated and consumed electricity, then, as the increase of the generated electricity in the process of changing the orientation of solar panels in the Sun and exceeding its electricity consumption adjust temperatures in the zones of the x control starting with the lowest values, by their increase up to the upper limit values and as values of the generated electricity repeat cycles correction of the temperature in thermal zones KA as above described, and upon completion of the current program flight KA specified duration and before the next again produce the above breakdown for typical intervals of flight time and plots with different levels consumed and generated electricity for the specified forecast use values consumed and generated power measured on the model performed intervals flight, SPACECRAFT, and then repeat the cycle correction temperature in thermal zones KA in this way.



 

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Heat pipe // 2254533

FIELD: heat power engineering.

SUBSTANCE: heat pipe comprises vertical housing with evaporation and condensation zones and partially filled with heat-transfer agent and coaxial hollow insert in the evaporation zone which defines a ring space with the housing and is provided with outer fining. An additional hollow cylindrical insert of variable radius made of a non-heat-conducting material is interposed between the condensation zone and coaxial hollow insert. The outer side of the additional insert and inner side of the housing of the heat pipe define a closed space.

EFFECT: reduced metal consumption.

1 dwg

Microcooling device // 2247912

FIELD: cooling equipment, particularly heat exchange apparatuses.

SUBSTANCE: device to remove heat from heat-generation component includes coolant stored in liquid coolant storage part, heat absorbing part including at least one the first microchannel and installed near heat-generation component. Heat absorbing part communicates with storage part. Liquid coolant partly fills microchannel due to surface tension force and evaporates into above microchannel with gaseous coolant generation during absorbing heat from heat generation component. Device has coolant condensing part including at least one the second microchannel connected to above coolant storage part separately from the first microchannel, gaseous coolant movement part located near heat-absorbing part and condensing part and used for gaseous coolant movement from the first microchannel to the second one. Device has case in which at least heat-absorbing part is placed and heat-insulation part adjoining heat absorbing part to prevent heat absorbed by above part from migration to another device parts.

EFFECT: reduced size, increased refrigeration capacity, prevention of gravity and equipment position influence on device operation.

22 cl, 4 dwg

The invention relates to heat engineering, namely, devices for heat transfer

FIELD: space engineering; temperature control systems of communication satellites.

SUBSTANCE: proposed method includes cooling the reservoir for drainage of heat-transfer agent from liquid loop of system to obtain required pressure of saturated vapor of heat-transfer agent above its surface in this reservoir. This pressure is selected from definite ratio so that it is lesser than pressure of saturated vapor for measured temperature of liquid loop at its escape during drainage. Drainage is continued till pressure of saturated vapor of heat-transfer agent at liquid loop outlet gets equal to pressure of saturated vapor in this reservoir.

EFFECT: reduction of losses of heat-transfer agent; enhanced ecological safety.

2 dwg

FIELD: space engineering; filling liquid loops of spacecraft temperature control systems with de-aerated low-boiling heat-transfer agent.

SUBSTANCE: prior to filling the loop with heat-transfer agent, required amount of heat-transfer agent shall be drained to metering reservoir. Increased pressure satisfying definite hydrostatic conditions shall be maintained in lines. Filling device is additionally provided with unit whose inlet is connected to thermostatted reservoir filled with de-aerated heat-transfer agent and is connected with outlet hydraulic joint of unit by means of pipe line through valve and filter; outlet hydraulic joint is connected to inlet hydraulic joint of liquid loop of temperature control system. Connected between filter and valve are compound pressure and vacuum gauge, metering reservoir and evacuation line.

EFFECT: improved quality of filling the loop.

3 cl, 4 dwg

FIELD: space engineering; temperature control systems of communication satellites; refrigerating plants.

SUBSTANCE: deaeration is performed in thermostatted reservoir. Closed cavity above liquid surface in reservoir is connected with vacuum pump by means of valve and adjustable throttle valve. Fitted between valve and adjustable throttle valve is pressure meter for measuring pressure in said cavity. Evacuation of this cavity is performed in stepwise manner at simultaneous measurement of temperature and pressure. Difference between measured steadied pressure and pressure of saturated vapor of liquid shall be minimum at measured temperature.

EFFECT: simplified construction of deaeration facilities; reduced losses of liquid in the course of deaeration.

3 cl, 3 dwg

FIELD: space engineering; manufacture of communication satellites.

SUBSTANCE: proposed method includes testing the payload for serviceability. To this end, technological compensating device is connected to module before and after tests. Gas cavity of this device is preliminarily filled with gas under pressure and definite amount of heat-transfer agent is drained from liquid cavity. Pressure and amount of liquid being drained are selected in definite ratio.

EFFECT: reduction of rejects, hidden rejects inclusive.

2 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: 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: spacecraft temperature control systems.

SUBSTANCE: proposed method includes measurement of temperature of spacecraft structural members and onboard equipment and components of rocket propellant, heating them by celestial body heat and conversion of electrical energy into thermal energy as measured temperatures reach low limits of thermostatting range. In flight, intervals of thermal energy accumulation in propellant components (at excess of thermal energy and electric power on board) and intervals of its free liberation are determined. In case expected magnitude of accumulated energy during predetermined interval exceeds upper level for preset volume of propellant, heat of celestial bodies is accumulated till the end of this interval. Otherwise, excess of electric power generated on board is converted into heat which is delivered to propellant components. In predicting release of thermal energy from propellant components, its residual amount required for maintaining the propellant component temperature within required ranges is determined; temperature of structural members and onboard equipment is also measured. In case this temperature exceeds permissible levels, delivery of heat is discontinued. When temperature of propellant component gets beyond threshold magnitudes, removal of heat from propellant components is discontinued. Otherwise, delivery of heat to thermostattable elements and onboard equipment and/or to points of accumulation of heat for subsequent useful conversion is continued till beginning of next interval of accumulation of thermal energy. Then, thermal energy accumulation cycle is repeated.

EFFECT: enhanced efficiency of accumulation and release of thermal energy; reduced mass and overall dimensions; enhanced heat removal.

5 dwg

FIELD: spacecraft temperature control systems.

SUBSTANCE: proposed method includes measurement of temperature in areas of radiation surfaces of temperature control system, comparison of these temperatures with upper and low limiting magnitudes and delivery of heat to radiation surface when temperatures are below low magnitudes. Flight intervals at power requirement exceeding power generated by primary onboard power sources are determined. Amount of electric power consumed for temperature control of radiation surfaces is determined at the same intervals. Flight intervals for maximum possible accumulation of thermal energy on radiation surface in said zones within permissible temperatures are also determined. Expenses for radiation surface temperature control is taken into account. Before beginning of flight intervals at consumed electric power exceeding electric power generated by onboard power sources, heat is delivered to radiation surface zones which require consumption of power for their temperature control at these intervals. Delivery of heat is performed with upper limiting magnitudes of temperatures taken into account.

EFFECT: reduced loading of spacecraft power supply system due to reduced power requirement for radiation surface temperature control at retained preset temperature ranges on these surfaces.

3 dwg

FIELD: manufacture of heat control systems of communication, TV broadcasting and retransmission systems.

SUBSTANCE: proposed method includes manufacture of at least three similar thermal load simulators 2 at width of contact surface equal to width of web of heat-transfer agent collector. Heat transfer factors (K1,K2, K3) are determined at similar forces of pressing the contact plate of simulators 2 to skin surface of panel 1. Heat-transfer temperature at collector inlet is maintained equal to surrounding temperature. Said heat transfer factors are determined as follows: for simulator (K1) separately mounted on panel; for simulator (K2) mounted between two adjacent simulators in way of motion of heat-transfer agent; for simulator under conditions when other simulators (K3) are mounted opposite web of adjacent turns symmetrically relative to it. Quality of construction and technology of manufacture is judged from the following relationship: K2+K3-K1-ΔKj ≥ [K], where ΔKj is factor of influence of thermal resistance of joint between simulator contact surfaces and skin on heat-transfer factor; [K] is permissible magnitude of heat-transfer factor.

EFFECT: simplified procedure of check of honeycomb panel quality; low cost of manufacture of panels.

3 dwg

FIELD: communication, TV broadcasting and information retransmission satellites and their heat control systems.

SUBSTANCE: on-board device with concentrated heat source is placed inside inner cavity of heat-insulated closed liquid-radiation heat exchanger and excessive heat from this device is removed to circulating water supply line. Inlet and outlet of liquid cavity of said heat exchanger are connected with delivery and discharge lines of circulating water supply system. Connecting pipe lines are provided with drainage and cutoff valves mounted before inlet and outlet of said liquid cavity below heat exchanger level. Parameters of pipe line running from delivery main of circulating water supply system to inlet of said liquid cavity are selected according to special condition. Box-shaped liquid-radiation heat exchanger consists of two sections with double wall: base and hood; their liquid cavities are communicated with atmosphere through drainage holes provided with shut-off members.

EFFECT: facilitated procedure; reduction of expenses.

3 cl, 5 dwg

FIELD: space technology.

SUBSTANCE: unconfined space of gas chamber of hydro-pneumatic compensator is subject to periodical change at the same average-mass temperature of heat-transfer agent. The ratio of Vi≤(Vi+l+nϕ) 1) is used to judge if leak-proofness corresponds to standard value, where Vi is volume of gas chamber of hydro-pneumatic compensator for i-th measurement, Vi+l is volume of gas chamber of hydropneumatic compensator for subsequent measurement, n is time interval between i-th and i+1 measurement, ϕ is standard value of volumetric loss of heat-transfer agent during specific time interval. Difference in unconfined spaces achieved between (i+1)-th and i-th measurement is used to determine real leakage of heat-transfer agent from system during specific time interval. Current value of unconfined space of system hydro-pneumatic compensator gas chamber is measured instead of measuring working pressure of the system for the same average-mass temperature of heat-transfer agent. Difference between measured spaces related to time interval between measurements has to be value of real leakage of heat-transfer agent observed during specific time interval.

EFFECT: simplified and reliable method of inspection.

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