CN204998798U - Close on spacecraft and hot accuse system thereof - Google Patents

Abstract

The utility model discloses an aircraft near space and a thermal control system thereof. The thermal control system of the near-space vehicle of the present invention comprises a shrinkable capsule with expansion and contraction functions arranged on the outer surface of the near-space vehicle shell. There is also a phase-change heat-absorbing part on the surface, and the phase-change heat-absorbing part is used to absorb the radiant heat in the cabin of the adjacent space vehicle. layer; said first radiant coating absorbs radiant heat within said adjacent space vehicle cabin and conducts it to said phase change heat sink. The thermal control system of the utility model can ensure that the temperature in the cabin of the adjacent space vehicle is relatively stable, and has a simple structure and a simple thermal control method. The utility model adopts the passive thermal control cost for the near-space aircraft, which can effectively ensure that the temperature in the near-space aircraft cabin is relatively constant, so that the equipment in the cabin can work normally.

Description

Near space vehicle and its thermal control system

technical field

The utility model belongs to the technical field of near-space aircraft, in particular to a thermal control system of a near-space aircraft and a near-space aircraft containing the thermal control system.

Background technique

Near space is a new space developed by the human economy, and near space vehicles will reside in this area for a long time in the future. Whether the equipment cabin of the aircraft can provide a relatively stable temperature environment determines the basic conditions for the normal operation of the equipment in the cabin.

Because the ambient temperature in the airtight cabin of the aircraft is affected by the ambient temperature in the adjacent space, solar radiation, infrared radiation and solar reflection, and the heating of equipment in the cabin, the temperature may be too high during the day and too cold at night. In the external environment where the heat of electronic equipment is difficult to discharge, it is necessary to take effective thermal control measures to ensure that the temperature of electronic equipment is within a safe temperature. Taking supersonic aircraft flying in the atmosphere for a long time as an example, the total heating capacity of the aircraft is large, the thermal environment is relatively harsh, and the cabin space is small and the thermal sealing performance is good. The ambient temperature in the cabin is also high or low, and the heat is difficult to Evacuate to atmosphere or have difficulty absorbing heat. It is impossible to keep the ambient temperature in the cabin relatively constant, and it is difficult to effectively ensure the normal operation of the equipment in the cabin.

The current thermal control measures for aircraft are usually divided into passive thermal control methods and active thermal control methods. However, due to the impact of the current aircraft structure, the load of the energy system is limited, and it is not suitable to adopt the active temperature control strategy. Therefore, the thermal control of the equipment compartment is proposed. new requirements.

Among them, the passive thermal control method is easy to install, reliable in performance, light in weight, low in cost and good in versatility. The passive scheme mainly suppresses the heating effect of the thermal environment of the airtight cabin on the equipment by changing the installation form of the instrument, adding heat insulation pads, adding phase change materials, installing temperature control covers, and pasting aluminum foil, so as to realize long-term working Temperature control of electronic equipment. Traditional passive thermal control methods generally only take thermal control measures for electronic devices. For high-temperature airtight environments with complex, narrow interiors and multiple electronic devices, taking thermal control measures for individual electronic devices one by one will bring greater thermal control costs. Taking the entire cabin as the thermal control object, a multi-level and step-by-step thermal control method is adopted to suppress the heat introduced into the electronic equipment. Take thermal control measures for the airtight cabin shell to reduce the heat introduced into the cabin from the source, and adopt different thermal control measures or a combination of multiple thermal control measures for the different heating characteristics and structural characteristics of electronic equipment to suppress heat transfer through heat conduction. , radiation and natural convection heat the heat entering the electronic equipment to achieve temperature control for the electronic equipment in the complex cabin. Therefore, the current passive temperature control measures are complicated and still bring a large thermal control cost.

Utility model content

The purpose of the utility model is to overcome the shortcomings of the above-mentioned prior art, and provide a near-space aircraft and its thermal control system, which solves the current problem of complex passive temperature control measures for near-space aircraft and brings relatively large thermal control costs. technical problem.

For achieving the above object, the technical scheme that the utility model adopts is:

A thermal control system for a near-space vehicle, comprising a shrinkable capsule with a shrinking function arranged on the outer surface of the close-to-space vehicle shell, and on the surface of the shrinkable capsule opposite to the contact surface with the outer surface of the close-to-space vehicle shell There is also a phase-change heat-absorbing part on it, and the phase-change heat-absorbing part is used to absorb the radiant heat in the cabin of the adjacent space vehicle, and a first radiation coating is also provided on the outside of the phase-change heat-absorbing part ; the first radiant coating absorbs the radiant heat in the cabin of the adjacent space vehicle and conducts it to the phase-change heat-absorbing component.

Preferably, the thermal control system of the near-space vehicle further includes a flow storage tank, and the flow storage tank communicates with the shrinkage capsule body through a heat transfer fluid channel, so that the heat transfer fluid flows between the flow storage tank and the shrinkage bladder body. circulation.

Specifically, the accumulator tank is arranged on the surface of the contraction bladder opposite to the contact surface with the outer surface of the shell of the adjacent space vehicle.

Further preferably, the shrinkable balloon includes a first shrinkable balloon and a second shrinkable balloon, the cavity of the first shrinkable balloon communicates with the cavity of the second shrinkable balloon, and the phase change suction The thermal component is arranged on the surface of the first shrinkable bladder opposite to the surface contacting the outer surface of the adjacent space vehicle shell, and the accumulator is arranged on the surface of the second shrinkable bladder opposite to the adjacent outer surface of the spacecraft shell. The outer surface of the spacecraft shell is on the surface of the contact surface, and the accumulator tank communicates with the second contraction bladder through the heat-conducting fluid channel.

Further preferably, on the surface of the first shrinkable capsule body opposite to the surface in contact with the outer surface of the adjacent space vehicle shell, there is a device that prevents the phase change heat absorbing component from radiating heat into the cabin of the near space vehicle. Second radiant coating.

Specifically, the thickness of the second radiation coating is 50-150 microns.

Further preferably, the surface of the first shrinkable capsule opposite to the contact surface with the first radiation coating is further provided with a third radiation coating which is conducive to absorbing radiant heat in the cabin of the near-space vehicle.

Specifically, the thickness of the third radiation coating is 50-150 microns.

Further preferably, the outer surface of the second shrinkable capsule is further coated with a fourth radiation coating that facilitates heat radiation to the heat transfer fluid in the second shrinkable capsule.

Specifically, the fourth radiation coating has a thickness of 50-150 microns.

Preferably, the outer surface of the flow storage tank is further provided with a fifth radiation coating for preventing the heat transfer fluid in the flow storage tank from radiating heat to the outside.

Further preferably, the thickness of the fifth radiation coating is controlled to be 50-150 microns.

Further preferably, the cavity of the first shrinkable balloon and/or the second shrinkable balloon is divided into several compartments, wherein the several compartments in the first shrinkable balloon communicate with each other or are respectively connected to Several compartments in the second shrinkable bladder communicate; several compartments in the second shrinkable bladder communicate with each other or communicate with the flow storage tank through the heat-conducting fluid channel.

Preferably, the thickness of the first radiation coating is controlled to be 50-150 microns.

And, a near-space vehicle, which includes the thermal control system of the near-space vehicle described in the present utility model.

Compared with the prior art, the utility model provides a thermal control system for a near-space vehicle by arranging a shrinkable capsule on the outer surface of the near-space vehicle shell, and setting a phase change absorber on the surface of the shrinkable capsule away from the outer surface of the near-space vehicle shell. The thermal component uses the first radiant coating to absorb the radiant heat in the cabin of the adjacent space vehicle and conducts it to the phase-change heat-absorbing component, so that the phase-change heat-absorbing component generates and transmits sunlight adjacent to the equipment in the cabin of the space vehicle At the same time, the expansion and contraction of the shrinkable capsule can also effectively absorb the radiant heat produced in the cabin of the adjacent space vehicle, thereby keeping the temperature in the cabin of the adjacent space vehicle relatively constant . Therefore, the thermal control system of the near-space vehicle of the present invention adopts a passive thermal control method to make the temperature in the cabin of the near-space vehicle relatively stable, which provides a guarantee for the normal operation of the equipment in the cabin of the near-space vehicle, and its structure is simple, and the thermal control method Simple.

The near space aircraft of the utility model adopts the above-mentioned thermal control system of the near space aircraft of the utility model, so the cost of passive thermal control for the near space aircraft is small, and it can effectively ensure that the temperature in the cabin of the near space aircraft is relatively constant, so that the temperature in the cabin is relatively constant. The device works normally.

Description of drawings

Fig. 1 is a schematic structural diagram of a thermal control system for a near-space vehicle provided by an embodiment of the present invention;

Fig. 2 is an enlarged schematic diagram of the structure of the shrinkable capsule in the thermal control system of the near-space vehicle provided by the embodiment of the present invention.

detailed description

In order to make the purpose, technical solutions and advantages of the utility model clearer, the utility model will be further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described here are only used to explain the utility model, and are not intended to limit the utility model.

It should be noted that when an element is described as being “fixed” or “disposed on” another element, it may be directly on the other element or there may be an intervening element at the same time. When an element is described as being "connected to" another element, it can be directly connected to the other element or intervening elements may also be present.

It should also be noted that the orientation terms such as left, right, up, down, top, and bottom in this embodiment are only relative concepts or refer to the normal use state of the product, and should not be regarded as having restrictive.

The embodiment of the utility model provides a passive thermal control system with low cost of thermal control, which can effectively keep the temperature in the cabin of an adjacent space vehicle relatively constant. The structure of the thermal control system of the near-space vehicle is shown in Figures 1 and 2, which includes a shrinkable capsule 20 arranged on the outer surface of the close-to-space vehicle shell 4. A phase change heat absorbing component 10 is also provided on the surface of the surface contact surface of the outer shell 4 of the aircraft.

Wherein, the phase-change heat-absorbing component 10 contains a phase-change heat-absorbing material, so it has the functions of absorbing heat, storing heat and releasing heat. In one embodiment, the phase-change heat-absorbing material in the phase-change heat-absorbing component 10 may be an inorganic phase-change material, an organic phase-change material, or a mixture of inorganic and organic phase-change materials. In a specific embodiment, the inorganic phase-change material can be inorganic substances such as crystalline hydrated salt, molten salt, metal alloy; the organic phase-change material can be organic substances such as paraffin, carboxylic acid, ester, and polyol. Therefore, in a specific embodiment, the phase-change heat absorbing component 10 may be a phase-change material layer structure or a capsule structure formed by these phase-change materials.

In order to enable the phase-change heat-absorbing component 1 to effectively absorb and dissipate heat through phase-change, in one embodiment, a first radiation coating 11 is provided on the outer surface of the phase-change heat-absorbing component 10 . The setting of the first radiant coating 11 can absorb the radiant heat in the cabin of the adjacent space vehicle and sunlight and conduct it to the phase-change heat-absorbing component 10, so that the phase-change of the phase-change heat-absorbing material can store heat. The first radiation coating 11 is preferably coated on the entire surface of the phase-change heat absorbing component 10 .

In order to improve the effect of absorbing radiant heat of the first radiation coating 11, in one embodiment, the material of the first radiation coating 11 is a radiation material with a high emissivity coefficient, and the radiation material with a high emissivity coefficient can be selected from conventional materials in the field. The high emissivity material, such as in a specific embodiment, the high emissivity material is selected from SiC, metal oxide, boride, sulfide, selenide and the like.

In another embodiment, the thickness of the first radiation coating 11 is controlled to be 50-150 microns.

One surface of the shrinkable capsule 20 is in contact with the first radiation coating 11 on the phase-change heat-absorbing component 10 , and the other surface of the shrinkable capsule 20 is in contact with the surface of the shrinkable capsule 20 relative to the contact surface with the first radiation coating 11 . Adjacent to the outer surface of the spacecraft shell. In this way, after the first radiant coating 11 absorbs radiant heat, such as the heat in the near-space environment and the radiant heat generated by the equipment 5 in the near-space vehicle cabin, especially the radiant heat of sunlight during the day, the heat is directly transferred to the corresponding The heat-changing heat-absorbing component 10 performs phase-change heat storage. At the same time, in the process of the first radiant coating 11 absorbing radiant heat, the shrinkable capsule body 20 will also exert its expansion characteristics due to the absorbed radiant heat, realizing the phase transition from the radiant heat generated by the equipment 5 in the adjacent space vehicle cabin to The heat-absorbing component 10 transfers and prevents external heat from being transferred to the cabin of the near-space vehicle, avoids excessive temperature rise in the cabin of the near-space vehicle, and keeps the temperature in the cabin of the near-space vehicle relatively stable.

When the ambient temperature of the near space is too low, especially at night, the temperature in the cabin of the near space vehicle will drop suddenly. During this process, the volume of the shrinkable capsule 20 will shrink in the low-temperature environment, so that the phase-change heat-absorbing component 10 will also approach the adjacent space vehicle shell 4 due to the shrinkage of the shrinkable capsule 20, and pass The thermal radiation effect of the first radiation coating 11 on the surface of the phase-change heat-absorbing component 10 radiates heat into the cabin of the adjacent space vehicle, so that the temperature in the cabin of the adjacent space vehicle remains relatively stable.

In one embodiment, the expansion and contraction functions of the deflationary capsule 20 are realized by the heat-conducting fluid filled in the deflationary capsule. In some specific embodiments, the heat transfer fluid may be a heat transfer fluid commonly used in the field, for example, in a specific embodiment, the heat transfer fluid may be alcohol or the like. The selection of these heat-conducting fluids can play a very good heat-conducting effect and/or cold-heat shrinkage effect.

In a specific embodiment, the cavity of the shrinkable bladder 20 is divided into several compartments 23, and the several compartments 23 communicate with each other or the several compartments 23 are respectively connected to the flow tank 30 through the heat conducting The fluid channels 24 communicate, as shown in FIG. 2 . The cavity of the shrinkable bladder 20 is divided into several compartments 23, so that the heat-conducting fluid can be evenly filled in the cavity of the shrinkable bladder 20, and it is also possible to avoid local deformation of the shrinkable bladder 20 due to excessive expansion or reduce the use of the shrinkable bladder 20 life.

In a further embodiment, the thermal control system of the near-space vehicle further includes a flow storage tank 30 , and the flow storage tank 30 communicates with the shrinkable bladder 20 through the heat-conducting fluid channel 24 . The accumulator tank 30 is used to store the heat transfer fluid. When the heat transfer fluid is radiated with heat, the heat transfer fluid expands and fills the cavity of the shrinkable bladder 20 through the heat transfer fluid channel 24, so that the shrinkable bladder 20 expands and expands. Or buffer the expansion of the heat transfer fluid in the cavity of the deflation balloon 20 . In a specific embodiment, the accumulator tank 30 is arranged on the surface of the shrinkable bladder 20 opposite to the surface in contact with the outer surface of the aircraft shell 4 in the vicinity of space. Of course, the accumulator tank 30 can also be provided according to installation requirements. On other positions adjacent to the shrinkable balloon body 20 .

In order to avoid or reduce the heat transfer fluid stored in the accumulator tank 30 to radiate heat outward, in one embodiment, a fifth radiant coating 31 is provided on the outer surface of the accumulator tank 30, which can effectively prevent the The heat transfer fluid in the accumulator tank 30 radiates heat to the outside.

In order to reduce the effect of the fifth radiation coating 31 radiating heat outward, in one embodiment, the material of the fifth radiation coating 31 is a radiation material with a low emissivity coefficient, and the radiation material with a low emissivity coefficient can be conventional in the art Low emissivity material, such as in a specific embodiment, low emissivity paint or low emissivity metal powder, etc. In another embodiment, the thickness of the fifth radiation coating 31 is controlled to be 50-150 microns.

On the basis of the embodiment of the thermal control system of the adjacent space vehicle including the flow storage tank 30, in one embodiment, the above-mentioned shrinkable bladder 20 includes a first shrinkable bladder 21 and a second shrinkable bladder 22, and the first shrinkable bladder The cavity of the body 21 communicates with the cavity of the second shrinkable balloon body 22 . At this time, in one embodiment, the phase-change heat-absorbing component 10 is arranged on the surface of the first shrinkable bladder 21 opposite to the surface in contact with the outer surface of the adjacent space vehicle shell 4, and the accumulator tank 30 is arranged on the surface of the second shrinkage bladder 22 opposite to the surface in contact with the outer surface of the adjacent space vehicle shell 4 , and the accumulator tank 30 is connected to the second shrinkage bladder 22 through the heat transfer fluid channel 24 connected. By arranging the shrinkable capsule body 20 as the first shrinkable capsule body 21 and the second shrinkable capsule body 22 whose cavities communicate, it is convenient to carry out regional treatment on the surface of the shrinkable capsule body 20, and it is possible to optimize the radiation heat in the adjacent space vehicle cabin to the opposite direction. The radiation transmission of the variable heat absorbing component 10 improves the relative stability of the temperature in the adjacent spacecraft cabin.

As in one embodiment, a second radiation coating 211 is also coated on the surface of the first shrinkage bladder 21 opposite to the surface in contact with the outer surface of the spacecraft shell 4, that is, the first shrinkage A second radiation coating 211 is also coated on the surface of the capsule 21 in contact with the phase-change heat absorbing component 10 . The setting of the second radiant coating 211, when it is necessary to export the radiant heat in the cabin of the adjacent space vehicle, such as in a high-temperature environment during the day, the phase-change heat-absorbing component 10 will continue to be affected by the external heat source, and the temperature will continue to rise. There may be a tendency for the heat absorbed by it to radiate into the cabin of an adjacent space vehicle, and the existence of the second radiation coating 211 can effectively prevent the heat of the external or phase-change heat-absorbing component 10 from radiating into the cabin of an adjacent space vehicle and cause Adverse effects of increased cabin temperature.

Therefore, in a specific embodiment, the material of the second radiation coating 211 is a material with a low emissivity coefficient, and the radiation material with a low emissivity coefficient may be a conventional low emissivity material in the art, such as in a specific embodiment, Low emissivity paint or low emissivity metal powder, etc. In another specific embodiment, the thickness of the second radiation coating 211 is controlled to be 50-150 microns. By adjusting the material and thickness of the second radiation coating 211, the effect of preventing the phase-change heat-absorbing component 10 or the external high-temperature environment of the adjacent space from radiating heat to the cabin of the adjacent space vehicle is improved.

Or in another embodiment, a third radiation coating 212 is also coated on the surface of the first deflationary capsule 21 opposite to the surface contacting the surface of the first radiation coating 11, that is, the first deflationary capsule. A third radiation coating 212 is also coated on the surface of 21 that is in contact with the outer surface of the spacecraft shell 4 . During the day, the heat transfer fluid in the first shrinkable bladder 21 and the second shrinkable bladder 22 expands after being radiated heat, and after the first shrinkable bladder 21 and the second shrinkable bladder 22 are filled, the surface of the first shrinkable bladder 21 The coated third radiant coating 212 can continue to absorb the radiant heat generated in the cabin of the adjacent spacecraft, and transmit it to the phase-change heat-absorbing component 10 .

In a specific embodiment, the material of the third radiation coating 212 is a material with a high emissivity coefficient, and the radiation material with a high emissivity coefficient can be a conventional high emissivity material in the field. For example, in a specific embodiment, the high emissivity Coefficient materials are selected from SiC, metal oxides, borides, sulfides, selenides, etc. In another specific embodiment, the thickness of the third radiation coating 212 is controlled to be 50-150 microns. By adjusting the material and thickness of the third radiation coating 212, the effect of the third radiation coating 212 on absorbing the radiant heat generated by the equipment 5 in the cabin of the adjacent space vehicle during operation is improved.

In a preferred embodiment, the first deflated capsule 21 is provided with the third radiation coating 211 on the surface of the first deflated capsule 21 opposite to the surface in contact with the first radiation coating 11. Coating 212. In this way, the first shrinkable capsule body 21 can effectively prevent the phase change heat absorbing component 10 or the external high temperature environment of the adjacent space from radiating heat to the cabin of the adjacent space vehicle, and at the same time conduct the radiant heat in the cabin of the adjacent space vehicle to the phase change heat absorber. Part 10, improves the stability of the cabin temperature of the adjacent space vehicle.

In one embodiment, the outer surface of the second shrinkable bladder 22 is further coated with a fourth radiation coating 221 that facilitates heat radiation to the heat-conducting fluid inside the second shrinkable bladder 22 . The fourth radiation coating 221 is preferably coated on the entire surface of the second deflationary balloon 22 .

In a specific embodiment, the material of the fourth radiation coating 221 is a material with a high emissivity coefficient, and the radiation material with a high emissivity coefficient can be a conventional high emissivity material in the field. For example, in a specific embodiment, the high emissivity Coefficient materials are selected from SiC, metal oxides, borides, sulfides, selenides, etc. In another specific embodiment, the thickness of the fourth radiation coating 221 is controlled to be 50-150 microns.

On the basis of the above-mentioned embodiments, in a specific embodiment, the cavity of the above-mentioned first shrinkable balloon body 21 or the second shrinkable balloon body 22 is divided into several compartments 23, or the first shrinkable balloon body 21 and the cavities of the second shrinkage balloon body 22 are separated by several compartments 23, wherein the several compartments 23 in the first shrinkage balloon body 21 communicate with each other or are respectively connected to several compartments in the second shrinkage balloon body 22. The chamber 23 communicates; the several compartments 23 in the second contraction capsule 22 communicate with each other or the several compartments 23 communicate with the reservoir tank 30 through the heat-conducting fluid channel 24 respectively, and the second contraction capsule 22 has a structure such as Figure 2 shows. The cavity of the first shrinkable bladder 21 and the cavity of the second shrinkable bladder 22 are divided into several compartments 23, so that the heat transfer fluid can be evenly filled in the cavity of the shrinkable bladder 20, and the first shrinkable bladder 21 can also be avoided. The cavity of the cavity and the second shrinkable balloon body 22 are locally over-expanded to deform or reduce the service life of the shrinkable balloon body 20 .

Therefore, the thermal control system of the near-space vehicle in the above-mentioned embodiments makes the thermal control system of the near-space vehicle according to the embodiment of the present utility model through the phase-change heat-absorbing component 10 and the shrinkage capsule 20 or the flow storage tank 30 further provided. It can ensure that the temperature in the cabin of an adjacent space vehicle is relatively stable, and the structure is simple and the thermal control method is simple.

Correspondingly, on the basis of the above-mentioned thermal control system for a near-space vehicle, the embodiment of the present invention also provides a near-space vehicle, which includes the above-mentioned thermal control system for a near-space vehicle according to the above-mentioned embodiment of the utility model. Therefore, the cost of adopting passive thermal control for the near-space vehicle is low, which can effectively ensure that the temperature in the cabin of the near-space vehicle is relatively constant, so that the equipment 5 in the cabin can work normally.

The above descriptions are only preferred embodiments of the present utility model, and are not intended to limit the present utility model. Any modification, equivalent replacement or improvement made within the spirit and principles of the present utility model shall be included in this utility model. within the scope of protection of utility models.

Claims (15)

1. a near space vehicle heat control system, it is characterized in that: comprise the contraction utricule with dilation function be arranged near space vehicle housing outer surface, described contraction utricule relative to the surface of described near space vehicle housing outer surface contact surface on be also provided with decalescence parts, described decalescence parts, for absorbing the radiations heat energy in described near space vehicle cabin, the outside of described decalescence parts are also provided with the first radiation coating; Described first radiation coating absorbs the radiations heat energy in described near space vehicle cabin and conducts to described decalescence parts.

2. near space vehicle heat control system as claimed in claim 1, it is characterized in that: also comprise Chu Liuxiang, described storage stream case is communicated with described contraction utricule by heat-conducting fluid passage, realizes the circulation of heat-conducting fluid between described storage stream case and described contraction utricule.

3. near space vehicle heat control system as claimed in claim 2, is characterized in that: described storage stream case be arranged on described contraction utricule relative to on the surface of described near space vehicle housing outer surface contact surface.

4. near space vehicle heat control system as claimed in claim 3, it is characterized in that: described contraction utricule comprises the first contraction utricule and second and shrinks utricule, described first cavity shrinking utricule communicates with the described second cavity shrinking utricule, described decalescence parts be arranged on described first shrink utricule relative to on the surface of described near space vehicle housing outer surface contact surface, described storage stream case be arranged on described second shrink utricule relative to on the surface of described near space vehicle housing outer surface contact surface, and described storage stream case shrinks utricule by described heat-conducting fluid passage and described second is communicated with.

5. near space vehicle heat control system as claimed in claim 4, is characterized in that: described first shrink utricule relative to the surface of described near space vehicle housing outer surface contact surface on be also provided with the second radiation coating stoping described decalescence parts to radiations heat energy in described near space vehicle cabin.

6. near space vehicle heat control system as claimed in claim 5, is characterized in that: the thickness of described second radiation coating is 50-150 micron.

7. near space vehicle heat control system as claimed in claim 4, is characterized in that: described first shrink utricule relative to described first radiation coating contact surface on be also provided with the 3rd radiation coating being beneficial to and absorbing radiations heat energy in described near space vehicle cabin.

8. near space vehicle heat control system as claimed in claim 7, is characterized in that: the thickness of described 3rd radiation coating is 50-150 micron.

9. near space vehicle heat control system as claimed in claim 4, is characterized in that: the outside face of described second contraction utricule is also coated with the 4th radiation coating being beneficial to and shrinking the heat-conducting fluid radiations heat energy in utricule to described second.

10. near space vehicle heat control system as claimed in claim 9, is characterized in that: the thickness of described 4th radiation coating is 50-150 micron.

11. as arbitrary in claim 2-10 as described near space vehicle heat control system, it is characterized in that: the outside face of described storage stream case is also provided with the 5th radiation coating of the outside radiations heat energy of described heat-conducting fluid for preventing in described storage stream case.

12. near space vehicle heat control systems as claimed in claim 11, is characterized in that: the THICKNESS CONTROL of described 5th radiation coating is 50-150 micron.

13. as arbitrary in claim 4-10 as described near space vehicle heat control system, it is characterized in that: the described first cavity shrinking utricule and/or the second contraction utricule is divided to be separated with several compartments, wherein, several compartments in described first contraction utricule communicate with each other or several compartments shunk in utricule communicate with second respectively; Described second several compartments shunk in utricule communicate with each other or several compartments flow case by described heat-conducting fluid channel connection with described storage respectively.

14. as arbitrary in claim 1-10 as described near space vehicle heat control system, it is characterized in that: the THICKNESS CONTROL of described first radiation coating is 50-150 micron.

15. 1 kinds of near space vehicles, it comprise as arbitrary in claim 1-14 as described near space vehicle heat control system.