RU2353923C1 - Device for measurement of radiant fluxes intensity in process of heat-vacuum testing of spacecrafts - Google Patents

Abstract

FIELD: physics.

SUBSTANCE: device is arranged from two neighboring assemblies, in every of which sensitive element on insulating substrate is installed inside corresponding assembly body; mentioned bodies are arranged in the form of rectilinear right-angle prism and/or circular straight cylinder; for every assembly one base of body is arranged with maximum high coefficient of heat radiation (ε₁ ^max>0.5 and ε₂ ^max>0.5), and the other base and side surface of body for every assembly are arranged with minimum coefficient of heat radiation (ε₁ ^min<0.5 and ε₂ ^min<0.5); at that base of the first assembly body with ε₁ ^max is installed parallel to surface of controlled section of spacecraft and is directed to the side opposite to it, and base of the second assembly body with ε₂ ^max is installed parallel to surface of controlled section of spacecraft and is directed towards it.

EFFECT: simplification of design and increased reliability of measurements.

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Description

The invention relates to space technology, namely, to control the heat transfer of a spacecraft (SC), with a space environment simulated in ground-based vacuum chambers during thermal vacuum tests (TWI), using infrared thermal receivers.

As a rule, TWI is subjected to spacecraft or its individual elements with a very complex configuration of the outer surface, the temperature of which is entirely determined by the heat balance with the surrounding elements and internal heat sources. In thermal vacuum chambers (TCEs), the position of the spacecraft and its orientation with respect to external energy sources are simulated using heat radiation simulators, in particular infrared radiation simulators (IKI). A feature of the spacecraft, and this is modeled during tests in TCEs, is that the main factor determining the reliability and durability of the spacecraft is the stability of its thermal regime, which is in the temperature range (200-350) K. That is, at temperatures that correspond to the maximum intensity infrared radiation in the region (7-15) microns, characteristic of the effective radiation of the spacecraft. In this case, the intensity of the incident radiation to individual parts of the outer surface of the spacecraft can vary over a wide range. In addition, it must be borne in mind that individual parts of the spacecraft surface made of different materials or having different coatings have different optical characteristics, which are generally a function of the radiation wavelength, direction of the rays, and body temperature. In addition, if other spacecraft surfaces are “visible” from some part of the surface, then the effective degree of blackness (ε) of such a part also depends on the geometry of the body, which makes an exact theoretical calculation of the effective degree of blackness practically impossible [1, p.120] . Hence the need to control both the external thermal fluxes incident on the controlled portion of TCEs (q _TCEs) including background radiation from the elements and from TCEs IZHI and effective radiation controlled area CA (q _KA).

Calorimetric, thermoelectric and photometric radiant energy detectors are used to measure the intensity of radiant fluxes during thermal vacuum tests of spacecraft [2, p.130].

The calorimetric radiant energy receiver is a vessel, the walls of which are lined with turns of a thin-walled metal (copper) tube, through which a coolant flows, heated by absorbed radiant energy incident on the input area of the calorimeter. The inner surface of the calorimeter is blackened, which leads to the absence of selectivity for the wavelengths of the received radiation. The external thermal insulation of the receiver ensures that there is no influence of extraneous heat fluxes on its readings. The absorbed power is determined by the known flow rate of the coolant and its temperature at the inlet and outlet of the calorimeter. The intensity of the radiant flux is determined from the equality of power absorbed by the coolant and incident on the known input area of the calorimeter.

The disadvantage of this radiation receiver is the large inertia and the need to introduce flexible tubes into the TCEs for water supply [1, p. 301].

Thermoelectric radiant energy receivers generate an electrical signal proportional to the temperature difference of two surfaces, one of which perceives the measured radiant flux, and the second (back) is maintained at a constant temperature. The irradiated surface of the receiver is blackened to eliminate frequency selectivity, and the back side is thermostabilized using a circulating coolant or electric heaters.

The disadvantage of such receivers is their low sensitivity [1, p. 301].

Photometric receivers of radiant energy, as a rule, are elements of solar cells that convert the radiation incident on them directly into electric current. However, these elements have significant selectivity in the spectrum of absorbed radiation, and their signal depends on the temperature of the element [1, p.302].

The closest to the principle of action and the use of thermometric properties embedded in the design of the device for measuring the intensity of radiant fluxes is a sprayed metal bolometer [3, p. 241], [4, p. 50], [6, p. 56], which is chosen for prototype.

The device consists of a sensitive element representing a conductive layer of metal, which is deposited on a dielectric that serves as an electrically insulating substrate. The sensing element prepared in this way consists of a glass container in which a certain pressure of air or some inert gas is maintained. The cylinder has a window made of a material that is transparent to radiation from the spectral region for which the bolometer is intended. Wire bends from the ends of the conductive layer are brought out from the cylinder. Using sensitive equipment, the resistance of the sensitive element of the bolometer is measured, and the temperature acquired by the conductive layer (metal tape) due to the thermal radiation absorbed is determined by the value of this resistance. Thus, the intensity of the radiant flux is judged.

The disadvantages of the device, in addition to the sufficient complexity of the design and the cost of the experiments, include the fact that when measuring heat fluxes in TCEs, at low temperatures (for example, liquid nitrogen temperature), moisture condensation is possible on the sensitive element and on the cylinder window [4, p. .88], which will lead to a distortion of the measurement results and a decrease in the reliability of measurements.

The objective of the invention is to provide a device for measuring the intensity of radiant flux during thermal vacuum tests of a spacecraft, which would have sufficient simplicity of design with its reliability of measurements.

The problem is solved by a device for measuring the intensity of radiant flux during thermal vacuum tests of spacecraft, including a metal conductive sensitive element placed on an electrically insulating substrate, while the device is made of two adjacent assemblies, in each of which a sensitive element on an electrically insulating substrate is installed inside the housing of the corresponding assembly; the housing of each of the assemblies is made of a material with high thermal conductivity and is electrically insulated from the sensing element on the substrate; said bodies are made in the form of a regular straight prism and / or circular straight cylinder; the base areas of the housings of each of the two assemblies (s _o1 and s _o2 ) and the area of the side surfaces of the housings of these assemblies (s _b1 and s _b2 ) correspond to the relations s _b1 <s _o1 and s _b2 <s _o2 , and the heights of the shells of these assemblies (h ₁ and h ₂ ) and the characteristic sizes of the bases of these buildings (L ₁ and L ₂ ) correspond to the relations h ₁ << L ₁ and h ₂ << L ₂ ; for each of the assemblies, one housing base is made with the highest possible coefficient of thermal radiation

(ε ₁ ^max > 0.5 and ε ₂ ^max > 0.5), and the other base and side surface of the housing for each assembly are made with a minimum coefficient of thermal radiation

(ε ₁ ^min <0.5 and ε ₂ ^min <0.5); moreover, the base of the body of the first assembly with ε ₁ ^{max is} installed parallel to the surface of the controlled portion of the spacecraft and directed in the opposite direction from it, and the base of the body of the second assembly with ε ₂ ^{max is} installed parallel to the surface of the controlled portion of the spacecraft and directed towards it.

In order to reduce the influence of leaks (or influx) of thermal radiation through the side surface of the housing of each assembly and obtain close to the isothermal surface of the body, it is necessary that the area of the side surfaces (s _b1 and s _b2 ) and the values of the coefficient of thermal radiation (ε ₁ ^min and ε ₂ ^min ) were minimal, therefore it was suggested:

1) execute the bodies of each assembly, in which the heights (h ₁ and h ₂ ) and the characteristic dimensions of the bases (L _i and L ₂ ) correspond to the relations h ₁ << L ₁ and h ₂ << L ₂ ;

2) by special processing of the side surface (for example, polishing or applying a thin layer of a special coating) to obtain the minimum coefficient of thermal radiation of the side surface of the bodies of each assembly.

It was proposed to make the case of each assembly of Al, Cu, Ag or alloys based on them as metals having the highest thermal conductivity values [5, p. 340-343], which makes it possible to obtain an almost insulated surface of the case for each assembly in the equilibrium state. This circumstance (isothermality of the casing of each assembly) allows us to assume that radiation emission over the entire surface of the casing takes place under equal temperature conditions and will be used below in deriving relations for determining the intensity of radiant fluxes.

To identify, by fixing the temperature of the hull of each assembly, the effect on the thermal state of each assembly of separately incident radiation to the controlled spacecraft section from the bodies surrounding the spacecraft installed in the TCEs (q _TVK ) and the effective radiation of the surface of the controlled spacecraft section (q _spacecraft ), one the base of the case for each assembly should be performed with the highest possible coefficient of thermal radiation (ε ₁ ^max > 0.5 and ε ₂ ^max > 0.5), and the other base of the case for each assembly should be performed with the minimum coefficient of thermal radiation (ε ₁ ^min <0, 5 and ε ₂ ^min <0.5).

To create the surface of the base with the highest possible coefficient of thermal radiation, it is proposed to apply a coating on this surface with a high absorption capacity. It is proposed to make a coating of friable "black", representing a metallic substance in a finely divided state, with a layer thickness reaching 30-40 microns or more, in accordance with the maximum wavelengths of the absorbed radiation. It is proposed that this coating layer of loose “mobile” be made of Au, as having a maximum absorption coefficient in a wide range of wavelengths [6, p. 56]. It has also been proposed to use coatings based on pigments Al ₂ O ₃ , CaO, ZrO ₂ , ZnO, CuO to obtain a surface with the highest possible coefficient of thermal radiation, which have the stable and highest values of the coefficient of thermal radiation in the temperature range ~ 100-300 K, corresponding to working the conditions of the proposed device with TWI [5, p.779].

To obtain the base and side surface of the casing of each assembly with a minimum coefficient of thermal radiation (ε ₁ ^min <0.5 and ε ₂ ^min <0.5), it is proposed to make the casing of Al, Cu, Ag or alloys based on them. Polished housing surfaces made of these metals, in addition to high thermal conductivity, have low coefficients of thermal radiation [5, p. 780-783].

In the case of using the device in aggressive environments, leading to an increase in the coefficient of thermal radiation on the surfaces of the case where a low coefficient of thermal radiation is required, one of the assemblies, it is proposed to protect these surfaces with a coating of a thin layer (of the order of or less than a micron) of precious metals or alloys based on them, as chemically inert [7, p.183], and then also polish these surfaces [5, p.783].

The essence of the invention is illustrated in figures 1-4. Figure 1-4 shows an embodiment of the device, where the first assembly is made in the form of a regular straight prism (rectangular parallelepiped with a square base), and the second assembly is in the form of a circular straight cylinder.

Figure 1-4 shows a device for measuring the intensity of radiant flux, consisting of two assemblies:

1 - the first assembly, which includes: 2 - housing; 3 - cavity; 4 - a sensitive element; 5 - substrate; 6 - layer of electrical insulation; 7 - base; 8 - base; 9 - side surface; 10 - current leads;

11 - the second assembly, which includes: 12 - housing; 13 - cavity; 14 - a sensitive element; 15 - substrate; 16 - layer of electrical insulation; 17 - base; 18 - base; 19 - side surface; 20 - current leads.

The device for measuring the intensity of radiant flux consists of two assemblies - the first assembly 1 and the second assembly 11, each of which consists of housings 2 and 12 made of a material with high thermal conductivity, inside of which in the cavities 3 and 13 the sensitive elements 4 and 14 are placed on electrically insulating substrates 5 and 15. In each assembly, the sensitive elements 4 and 14 on the substrates 5 and 15 are separated from the housings 2 and 12 by an electrical insulation layer 6 and 16. The housings 2 and 12 each consist of two bases 7, 8 and 17, 18, respectively, and side surfaces 9 and 19. Reasons 7 and 17 are made with the highest possible coefficient of thermal radiation ε ₁ ^max for the first assembly 1 and ε ₂ ^max for the second assembly 11. The bases 8, 18 and side surfaces 9, 19 are made with the minimum coefficient of thermal radiation ε ₁ ^min and ε ₂ ^min respectively for the first assemblies 1 and second assembly 11. The sensitive elements 4 and 14 on the substrates 5 and 15 have pre-made calibration characteristics, expressed as a function of resistance to temperature. Through the electrically insulated current leads 10, 20, the sensitive elements 4, 14 of the first assembly 1 and the second assembly 11 can be connected either to a bridge circuit or to a circuit with load resistance (not shown), which allow detecting a change in the resistance of the sensitive elements 4 and 14 depending on temperature cases 2 and 12. In the first assembly 1 in figure 3 shows an embodiment of the production of the sensing element 4 by rectangular winding (zigzag) of a conductive metal wire or tape made, for example, of Cu or Pt on a substrate 5. In the second assembly 11 in Figure 4 represents a variant of manufacturing the sensor element 14 in the form of spirally winding a conductive metal wire or tape made, for example, of Cu or Pt on the substrate 15.

A device for measuring the intensity of radiant flux during thermal vacuum tests of spacecraft works as follows.

In accordance with the requirements for testing the spacecraft, suppose that it is necessary to conduct a TWI of a spacecraft placed in a thermal vacuum chamber (not shown). A spacecraft of complex shape, the temperature of which is determined by the heat balance with the surrounding bodies that simulate the external effect on the spacecraft's spacecraft, including those placed in the TCEs with infrared radiation simulators and cryogenic screens. In addition, the heat balance of the spacecraft takes into account sources of internal heat from equipment placed on the spacecraft. In accordance with preliminary theoretical calculations of the thermal regime of the spacecraft as a whole or of individual elements of its design, each controlled section of the spacecraft, which, according to the experimenter, is of greatest interest, is located next to the first assembly 1 and the second assembly 11.

In each of the respective assemblies — the first assembly 1 and the second assembly 11 — the sensitive elements 4 and 14 on the insulating substrates 5 and 15 are installed inside the housings 2 and 12. The housings 2 and 12 of each of the assemblies are made of material with high thermal conductivity and are electrically insulated from the sensor 4 and 14 on the substrate 5 and 15 with a layer of electrical insulation 6 and 16. Housings 2 and 12 of each of the assemblies are made in the form of a regular straight prism and / or circular straight cylinder. Accordingly, for each assembly (first assembly 1 and second assembly 11), the areas (s _o1 and s _o2 ) of the bases 7, 8 and 17, 18 of buildings 2 and 12, as well as the areas (s _b1 and s _b2 ) of the side surfaces 9 and 19 correspond to the relations s _b1 <s _o1 and s _b2 <s _o2 . The heights (h ₁ and h ₂ ) of buildings 2 and 12 and the characteristic dimensions (L ₁ and L ₂ ) of the bases 7, 8 and 17, 18 correspond to the relations h ₁ << L ₁ and h ₂ << L ₂ . For each assembly (first assembly 1 and second assembly 11), one base of the housing 7 and 17 is made with the highest possible coefficient of thermal radiation (ε ₁ ^max > 0.5 and ε ₂ ^max > 0.5), and the other base 8 and 18 and the lateral surface 9 and 19 of the housing 2 and 12 are made with a minimum coefficient of thermal radiation (ε ₁ ^min <0.5 and ε ₂ ^min <0.5).

The first assembly 1 and the second assembly 11 are installed so that for the first assembly 1, the base 7 with ε ₁ ^{max of the} hull 2 is parallel to the surface of the controlled portion of the spacecraft and is directed to the opposite side, and for the second assembly 11 the base 17 with ε ₂ ^{max of the} hull 12 is installed parallel to the surface of the controlled portion of the spacecraft and is directed towards it.

After checking the operability of the spacecraft equipment and all systems that provide TWI in TCEs, they begin testing. Reduce the pressure in the TCEs to 10 ^-3 ... 10 ^-4 Pa, cool the cryogenic screens to the temperature of liquid nitrogen, turn on the IR simulators and the equipment installed on the spacecraft, and begin the experiment.

On the bases 7 and 18 and on the side surfaces 9 and 19, thermal radiation from the bodies surrounding the spacecraft installed in the TCEs (q _TVK ) is incident, and on the bases 8 and 17, the effective radiation of the surface of the controlled portion of the spacecraft (q _KA ) is incident.

The radiant fluxes q _TVK and q _{KA are} partially absorbed by the housings 2 and 12, heating them evenly, due to the high thermal conductivity of the material from which these housings 2 and 12 are made. In addition, the geometric conditions are met

(s _b1 <s _o1 and s _b2 <s _o2 , for h ₁ << L ₁ and h ₂ << L ₂ ) and optical (for side surfaces 9 and 19 with ε ₁ ^min and ε ₂ ^min ) also allows to obtain a close to the isothermal surface of buildings 2 and 12.

Accordingly, the conductive metal tape or wire of the sensing elements 4 and 14 is heated to an average equilibrium temperature of the buildings 2 and 12.

A conductive metal strip or wire of the sensing elements 4 and 14 through current leads 10 and 20 is included in a small current circuit. In this case, the voltage at the ends of the tape or wire, which varies depending on the temperature of the medium (temperature of the housings 2 and 12), is fed to a fixing device (not shown). After reaching the equilibrium thermal state of the system under consideration, the electrical resistance of the metal tape or wire of the sensing elements 4 and 14 is simultaneously measured and the temperature T1 of the housing 2 of the first assembly 1 and the temperature T _{2 of the} housing 12 are determined by the calibration characteristics previously performed for these sensitive elements 4 and 14 second build 11.

The intensity of the radiant flux q _TVK and q _KA is determined from the solution of the system of equations

where ε ₁ = ε ₁ ^max / ε ₁ ^min and ε ₂ = ε ₂ ^max / ε ₂ ^min are the relative optical parameters for the first assembly 1 and for the second assembly 11, respectively;

s ₁ = s _b1 / s _o1 and s ₂ = s _b2 / s _o2 are relative geometric parameters for the first assembly 1 and for the second assembly 11, respectively;

σ is the Stefan-Boltzmann constant; σ = 5.67 · 10 ^-8 W / (m ² · K ⁴ ).

The system of equations (1) and (2) meets the following assumptions:

- the energy of radiant fluxes 2 and 12 perceived by the bodies is removed through their external surfaces to the environment only by radiation;

- heat losses through current leads 10 and 20 of the sensitive elements 4 and 14 are negligible and are not taken into account in the heat balance;

- Cases 2 and 12 are considered isothermal under steady state thermal condition of the system in question.

Equations (1) and (2) are obtained from the heat balance equations for the first assembly 1 and second assembly 11, respectively, and have the following form:

Equations (3) and (4) on the left side show the absorbed radiation from each base and the side surface of the hulls 2 and 12 from the spacecraft and the TCE elements, and on the right side the own radiation of each base and the side surface of the hulls 2 and 12.

и

) и относительные геометрические параметры (

и

) отвечают соотношениям

и

можно преобразовать уравнения (1) и (2) к видуIn the particular case when the relative optical parameters (for the first assembly 1 and for the second assembly 11

and

) and relative geometric parameters (

and

) correspond to the relations

and

it is possible to transform equations (1) and (2) to the form

Solving the system of equations (5) and (6), we obtain expressions for determining the intensity of the radiant fluxes q _TVK and q _KA in the form

- безразмерный коэффициент.Where

- dimensionless coefficient.

This particular case is easily realized in the manufacture of two assemblies that are identical in design and characteristics.

We give a calculated example of the use of a device for measuring the intensity of radiant fluxes during thermal vacuum tests of spacecraft.

The dimensions of all parameters in the calculation example are given in the International system of units.

We put the spacecraft in the TCEs. Near the controlled spacecraft section, we install the first assembly 1 and the second assembly 11. The first assembly 1 is installed so that the base 7 with ε ₁ ^{max is} parallel to the surface of the controlled spacecraft section and is directed in the opposite direction, and the second assembly 2 so that the base 17 with ε ₂ ^max was parallel to the surface of the controlled portion of the spacecraft and directed in its direction. Suppose that the housing 2 of the first assembly 1 is made of Ag in the form of a rectangular parallelepiped with a square base, and the housing 12 of the second assembly 11 is made of Ag in the form of a circular straight cylinder, as shown in figures 1 - 4. For each assembly in the cavities 3, 13, sensitive elements 4 and 14 (for example, a thin copper wire wound in the form of a flat tablet) are placed on substrates 5, 15 through an electrical insulation layer 6, 16.

Suppose that the housing 2 and 12 are made with the following geometric characteristics: h ₁ = 3 · 10 ^-3 m; h ₂ = 3 · 10 ^-3 m; L ₁ = 2 · 10 ^-2 m; L ₂ = 2 · 10 ^-2 m. The area of the bases and the side surfaces of the buildings 2 and 12 will respectively be:

s _o1 = 4 · 10 ^-4 m ² ; s _b1 = 2.4 · 10 ^-4 m ² ; s _o2 = π · 10 ^-4 ~ 3.14 · 10 ^-4 m ² ; s _b2 = π · 6 · 10 ^-5 ~ 1.9 · 10 ^-4 m ² . In this case, the relative geometric parameters for buildings 2 and 12 have values s ₁ = 0.6 and s ₂ = 0.6.

Suppose that the bases 7, 17 are made with a maximum coefficient of thermal radiation: ε ₁ ^max = 0.95 and ε ₂ ^max = 0.95. The surfaces of the bases 8, 18 of the buildings 2, 12 and the side surfaces 9, 19 are polished to create surfaces with a minimum coefficient of thermal radiation. Assume that for these surfaces ε ₁ ^min = 0.05 and ε ₂ ^min = 0.05. Moreover, the relative optical parameters of the first assembly 1 and the second assembly 11 have values ε ₁ = 19 and ε ₂ = 19, respectively.

и

, поэтому q_твк и q_KA определяем из (7) и (8).The intensity of radiant fluxes q _TVK and · q _KA can be determined from the system of equations (1) and (2), however, for our case, the relations

and

, therefore, q _TVK and q _{KA are} determined from (7) and (8).

и

,We first determine the dimensionless coefficient k, knowing

and

,

Upon reaching a pressure of 10 ^-3 -10 ^-4 Pa in the TCEs, the cryogenic screens are cooled to the temperature of liquid nitrogen, they include the IRI installed in the TCEs and the equipment installed on the SC, and then the experiment is started.

The radiant fluxes q _TVK and q _{KA are} partially absorbed by the housing 2 of the first assembly 1 and the housing 12 of the second assembly 11, uniformly heating them and, accordingly, the conductive metal wires of the sensing elements 4, 14 to the average equilibrium temperature of these buildings 2, 12. Using the previously performed calibration characteristics sensitive elements 4, 14 at the same time determine the temperature T _{1 of the} housing 2 for the first assembly 1 (put fixed T ₁ = 150 K) and the temperature T _{2 of the} housing 12 for the second assembly 11 (put fixed T ₂ = 120 K).

k, а также зафиксированные T₁, Т₂ и определяем интенсивность лучистых потоков q_твк и q_KA:We substitute the known values into relations (7) and (8)

k, as well as the recorded T ₁ , T ₂ and determine the intensity of the radiant flux q _TVK and q _KA :

q _TWC = k · σ · ( ε · T ₁ ⁴ -T ₂ ⁴ ) = 1/18 · 5.67 · 10 ^-8 · [19 · 150 ⁴ -120 ⁴ ] ≅29.6 W / m ² ;

q _KA = k · σ · [( ε + s ) · T ₂ ⁴ - (1+ s ) · T ₁ ⁴ ] = 1/18 · 5.67 · 10 ^-8 · [(19 + 0.6) · 120 ⁴ - (1 + 0.6) · 150 ⁴ ] ≅10.3 W / m ² .

Application of the proposed design of a device for measuring the intensity of radiant fluxes during thermal vacuum tests of spacecraft allows:

1) to control the magnitude of the heat flux incident on the controlled portion of the spacecraft, and the effective radiation of the controlled portion of the spacecraft;

2) use the proposed design of the device and the method of its implementation in the functional control and diagnostics of the system for ensuring the thermal regime of the spacecraft during thermal vacuum tests;

3) reduce the cost of ground-based experimental testing of spacecraft in the conditions of modeling the space environment in TCEs due to the simplicity of the design of the device and the method of its implementation;

4) determine the distribution of the density of heat fluxes on the outer surface of the spacecraft, using a simple analytical apparatus and experimental techniques;

5) measure the magnitude of the intensity of radiant fluxes according to the proposed functional dependencies, including the minimum number of process-controlling parameters that affect the accuracy of measurements;

6) to automate the process of experimental determination of radiant fluxes in a heat-vacuum chamber, using information from the corresponding heat-sensitive elements.

LITERATURE

1. Modeling of the thermal conditions of the spacecraft and its environment. Ed. Acad. G.I. Petrova. M .: Mechanical Engineering, 1971.

2. O. B. Andreychuk, N. N. Malakhov. Thermal tests of spacecraft. M .: Engineering, 1982.

3. Kriksunov L.Z. Guide to the basics of infrared technology. M .: Soviet radio, 1978.

4. M.N. Markov. Infrared receivers. M .: Nauka, 1968.

5. Physical quantities. Handbook Ed. I.S. Grigorieva, E.Z. Meilikhova. M .: Energoatomizdat, 1991.

6. Physical encyclopedic dictionary. M .: Soviet Encyclopedia, 1983.

7. Large Encyclopedic Dictionary of the Polytechnic. Ch. ed. A.Yu. Ishlinsky. M .: Big Russian Encyclopedia, 2000.

8. I.N. Bronstein, K.A.Semendyaev. A reference book in mathematics for engineers and students of technical colleges. M .: Nauka, 1986.

Claims (1)

A device for measuring the intensity of radiant flux during thermal vacuum tests of spacecraft, including a metal conductive sensitive element placed on an electrically insulating substrate, characterized in that the device is made of two adjacent assemblies, in each of which a sensing element on an electrically insulating substrate is installed inside the housing of the corresponding assembly; the housing of each of the assemblies is made of a material with high thermal conductivity and is electrically insulated from the sensing element on the substrate; said bodies are made in the form of a regular straight prism and / or circular straight cylinder; the base areas of the housings of each of the two assemblies (s _o1 and s _o2 ) and the area of the side surfaces of the housings of these assemblies (s _b1 and s _b2 ) correspond to the relations s _b1 <s _o1 and s _b2 <s _o2 , and the heights of the shells of these assemblies (h ₁ and h ₂ ) and the characteristic sizes of the bases of these buildings (L ₁ and L ₂ ) correspond to the relations h ₁ << L ₁ and h ₂ << L ₂ ; for each of the assemblies, one housing base is made with the highest possible coefficient of thermal radiation (ε ₁ ^max <0.5 and ε ₁ ^max <0.5), and the other base and side surface of the housing for each assembly is made with a minimum coefficient of thermal radiation (ε ₁ ^min <0.5 and ε ₂ ^min <0.5); moreover, the base of the body of the first assembly with ε ₁ ^{max is} installed parallel to the surface of the controlled portion of the spacecraft and directed in the opposite direction from it, and the base of the body of the second assembly with ε ₂ ^{max is} installed parallel to the surface of the controlled portion of the spacecraft and directed towards it.

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