阅读说明：本技术 一种航天器真空热环境模拟结构及温度调节方法 (Spacecraft vacuum thermal environment simulation structure and temperature adjusting method ) 是由 王晶 李日华 杜春林 秦家勇 刘波 吴东亮 于晨 安万庆 朱琳 张丽娜 于 2021-06-11 设计创作，主要内容包括：本申请公开了一种航天器真空热环境模拟结构及控温方法,其中模拟结构包括真空容器,真空容器内底部设置有支撑架；还包括固定在支撑架中部的梁式框架、若干对称固定在梁式框架上的卫星安装架、安装在卫星安装架上的卫星、竖直固定在梁式框架上且位于相邻的卫星安装架之间的星间主冷板、设置在每颗卫星底部的下冷板；每颗卫星的上表面及下表面设有灯阵及用于测温的温度传感器；灯阵用于加热,提供热环境。本申请通过在真空容器内将多颗卫星对称式布置在一梁式框架上,在卫星正下方与相邻两颗卫星之间设置冷板,避免了卫星热试验背景热流过大而引起温度模拟误差；在一个容器内成功实现了两颗卫星背景热流一致性和卫星间的热流无扰。(The application discloses a spacecraft vacuum thermal environment simulation structure and a temperature control method, wherein the simulation structure comprises a vacuum container, and a support frame is arranged at the bottom in the vacuum container; the system also comprises a beam type frame fixed in the middle of the supporting frame, a plurality of satellite mounting frames symmetrically fixed on the beam type frame, satellites installed on the satellite mounting frames, an inter-satellite main cooling plate vertically fixed on the beam type frame and positioned between the adjacent satellite mounting frames, and a lower cooling plate arranged at the bottom of each satellite; lamp arrays and temperature sensors for measuring temperature are arranged on the upper surface and the lower surface of each satellite; the lamp array is used for heating and providing a thermal environment. According to the satellite thermal test system, a plurality of satellites are symmetrically arranged on a beam type frame in a vacuum container, and a cold plate is arranged between the position under each satellite and two adjacent satellites, so that temperature simulation errors caused by overlarge background heat flow of a satellite thermal test are avoided; the consistency of background heat flows of two satellites and the noninterference of the heat flows among the satellites are successfully realized in one container.)

1. A spacecraft vacuum thermal environment simulation structure comprises a vacuum container, wherein a support frame is arranged at the bottom in the vacuum container; the method is characterized in that: the device also comprises a beam frame fixed in the middle of the support frame, a plurality of satellite mounting frames symmetrically fixed on the beam frame, satellites installed on the satellite mounting frames, an inter-satellite main cooling plate vertically fixed on the beam frame and positioned between the adjacent satellite mounting frames, and a lower cooling plate arranged at the bottom of each satellite; the upper surface of each satellite is provided with an upper lamp array and a temperature sensor for measuring temperature, and the lower surface of each satellite is provided with a lower lamp array and a temperature sensor for measuring temperature; the upper lamp array and the lower lamp array are used for heating and providing a thermal environment.

2. The spacecraft vacuum thermal environment simulation structure of claim 1, further comprising a horizontal tilt sensor mounted on the satellite, and a plurality of horizontal adjusting mechanisms uniformly arranged at the bottom of the beam frame; the horizontal adjusting mechanism is used for adjusting the vertical height of the beam type frame according to the signal of the horizontal inclination angle sensor.

3. A spacecraft vacuum thermal environment simulation structure according to claim 2, wherein the horizontal adjustment mechanism comprises a mounting seat which penetrates up and down and the top end of the mounting seat is fixed at the bottom of the beam type frame, and a push rod which penetrates through the mounting seat; the bottom end of the push rod is fixed on the support frame;

the push rod comprises a threaded cylinder positioned at the bottom and a threaded rod in threaded connection with the threaded cylinder;

a transmission case is fixed at the bottom of the mounting seat; the transmission case is driven by a motor to drive the threaded rod to rotate.

4. A spacecraft vacuum thermal environment simulation structure according to claim 3, wherein the side wall of the transmission case is provided with a heating sheet, and the heating sheet is controlled by a temperature controller to heat.

5. A spacecraft vacuum thermal environment simulation structure according to any of claims 1 to 4, further comprising a vacuum gauge for measuring pressure in the satellite; the vacuum gauge is installed on a cabin plate of the satellite through a bracket; a cabin plate of the satellite is provided with an opening corresponding to the measuring port of the vacuum gauge; the opening is sealed after the vacuum gauge is fixedly installed.

6. A spacecraft vacuum thermal environment simulation structure according to claim 5, wherein a thermal insulating spacer is provided between the support and the deck of the satellite.

7. A spacecraft vacuum thermal environment simulation structure according to claim 5, wherein the vacuum gauge is a hot cathode ionization gauge.

8. A spacecraft vacuum thermal environment simulation structure according to any one of claims 1 to 4, wherein the satellite mounting rack is an L-shaped mounting rack, and the L-shaped support and the satellites are symmetrically arranged on two sides of the main cooling plate between the satellites.

9. A spacecraft vacuum thermal environment temperature regulation method applied to the spacecraft vacuum thermal environment simulation structure of any one of claims 1 to 8, comprising the following steps of:

dividing each upper lamp array and each lower lamp array into a plurality of temperature control areas, wherein each temperature control area controls a current signal of the temperature control area through a PID arithmetic unit, and the current signal of each controller is determined according to the following steps:

starting timing by taking the experiment starting time as a reference;

acquiring real-time temperature signal data y of the current control period_t；

The set target temperature y for the current control cycle is determined by the following function:

y＝r-Δy+Δy*[1-exp(-t/T)]；

wherein r is the final temperature control target, Δ y is the real-time temperature signal data y_tThe difference from r; t is the acquisition of real-time temperature signal data y_tThe time of (d); t is a set time constant;

determining a control increment u (t) of the PID operator according to the following equation:

e(t)＝y-y_t；

wherein K_pIs the set adjustment coefficient.

Technical Field

The disclosure relates to the field of spacecraft ground thermal tests, in particular to a spacecraft vacuum thermal environment simulation structure and a temperature adjusting method.

Background

Before a launching field of a spacecraft or a satellite is launched, equivalent simulation of a space environment is required to be carried out in a vacuum container, namely, simulation of a vacuum thermal environment experienced by the satellite is required to be carried out, so that verification of a satellite thermal design model and reliability of on-board equipment is completed. Typically, vacuum thermal environment simulation is performed on a single satellite in a single container. Aiming at the requirement of batch production of Beidou satellite III, parallel thermal tests of a plurality of satellites are required to be carried out in a single container. Under the circumstance, the problem of consistency between mutual thermal interference among satellites and background heat flows of the satellites when a plurality of satellites are tested in parallel to simulate the environment needs to be solved. Meanwhile, a plurality of satellites are subjected to thermal tests, the exhaust volume is increased during the thermal tests of the satellites, the quantity of onboard equipment and external heat flow simulation equipment is increased in equal proportion, the problem of satellite exhaust needs to be solved, and the problems of low temperature control efficiency, poor control accuracy and the like of the thermal vacuum tests are solved. The invention provides a one-device multi-satellite parallel vacuum thermal test method, which provides a technical scheme of vacuum thermal environment simulation when a plurality of satellites simultaneously perform thermal tests in a single container.

Disclosure of Invention

In view of the above-mentioned drawbacks and deficiencies of the prior art, it is desirable to provide a spacecraft vacuum thermal environment simulation structure, which includes a vacuum container, a supporting frame disposed at a bottom of the vacuum container; the device also comprises a beam frame fixed in the middle of the support frame, a plurality of satellite mounting frames symmetrically fixed on the beam frame, satellites installed on the satellite mounting frames, an inter-satellite main cooling plate vertically fixed on the beam frame and positioned between the adjacent satellite mounting frames, and a lower cooling plate arranged at the bottom of each satellite; the upper surface of each satellite is provided with an upper lamp array and a temperature sensor for measuring temperature, and the lower surface of each satellite is provided with a lower lamp array and a temperature sensor for measuring temperature; the upper lamp array and the lower lamp array are used for heating and providing a thermal environment.

According to the technical scheme provided by the embodiment of the application, the system further comprises a horizontal tilt angle sensor arranged on the satellite and a plurality of horizontal adjusting mechanisms uniformly arranged at the bottom of the beam type frame; the horizontal adjusting mechanism is used for adjusting the vertical height of the beam type frame according to the signal of the horizontal inclination angle sensor.

According to the technical scheme provided by the embodiment of the application, the horizontal adjusting mechanism comprises an installation seat which penetrates through the beam type frame from top to bottom, and a push rod which penetrates through the installation seat; the bottom end of the push rod is fixed on the support frame;

the push rod comprises a threaded cylinder positioned at the bottom and a threaded rod in threaded connection with the threaded cylinder;

a transmission case is fixed at the bottom of the mounting seat; the transmission case is driven by a motor to drive the threaded rod to rotate.

The lateral wall of transmission machine case is equipped with the heating plate, the heating plate is controlled by the temperature control appearance and is heated.

According to the technical scheme provided by the embodiment of the application, the device further comprises a vacuum gauge for measuring the pressure in the satellite; the vacuum gauge is installed on a cabin plate of the satellite through a bracket; a cabin plate of the satellite is provided with an opening corresponding to the measuring port of the vacuum gauge; the opening is sealed after the vacuum gauge is fixedly installed.

According to the technical scheme provided by the embodiment of the application, a heat insulation gasket is arranged between the support and the cabin plate of the satellite.

According to the technical scheme provided by the embodiment of the application, the vacuum gauge is a hot cathode ionization gauge.

According to the technical scheme provided by the embodiment of the application, the satellite mounting frame is an L-shaped mounting frame, and the L-shaped support and the satellites are symmetrically arranged on two sides of the inter-satellite main cooling plate.

In a second aspect, the present application provides a method for adjusting a temperature of a spacecraft vacuum thermal environment, which is applied to the above spacecraft vacuum thermal environment simulation structure, and includes the following steps:

dividing each upper lamp array and each lower lamp array into a plurality of temperature control areas, wherein each temperature control area controls a current signal of the temperature control area through a PID arithmetic unit, and the current signal of each controller is determined according to the following steps:

starting timing by taking the experiment starting time as a reference;

acquiring real-time temperature signal data y of the current control period_t；

The set target temperature y for the current control cycle is determined by the following function:

y＝r-Δy+Δy*[1-exp(-t/T)]；

wherein r is the final temperature control target, Δ y is the real-time temperature signal data y_tThe difference from r; t is the acquisition of real-time temperature signal data y_tThe time of (d); t is a set time constant;

determining a control increment u (t) of the PID operator according to the following equation:

e(t)＝y-y_t；

wherein K_pIs the set adjustment coefficient.

According to the method, a plurality of satellites are symmetrically arranged on a beam type frame in a vacuum container, a cold plate is arranged between the position right below the satellites and two adjacent satellites, the satellites are guaranteed to be positioned under the background heat flow of a liquid nitrogen temperature region in all directions, the low background heat flow separation is guaranteed to be realized by the simulation accuracy of a heat test, and the temperature simulation error caused by overlarge background heat flow of the satellite heat test is avoided by arranging the lower cold plate below the satellites; the consistency of background heat flows of two satellites and the noninterference of the heat flows among the satellites are successfully realized in one container.

According to the technical scheme provided by the embodiment of the application, the horizontal adjusting mechanism and the horizontal inclination angle sensor arranged on the satellite are combined for use, the deformation amount of the beam type frame in the vertical direction is calculated and converted through the horizontal inclination angle measured by the sensor, then the push rod moves in the reverse direction to compensate the thermal structural deformation, and the requirement of high-precision horizontal degree adjustment of the satellite is met.

According to the technical scheme provided by the application, the temperature control target of each stage is gradually approached through the exponential transition process function curve, the temperature rising and falling rate is accurately controlled and adjusted, and the system overshoot is guaranteed to the maximum extent.

In summary, the high-strength bearing platform, the rotational symmetric satellite layout scheme and the low background heat flow separation technology are adopted, and the problems of consistency of background heat flows of two satellites and no interference of the heat flows among the satellites are successfully solved in one container; by adopting a high-precision horizontal regulation and control system, the levelness of the two satellites is superior to 1mm/1m in the test process, and the normal work of the satellite heat pipes is ensured; the in-satellite vacuum degree monitoring of the test full working condition is realized by adopting a high-reliability measuring method, and becomes a key criterion of a newly added degassing working condition of a thermal test; aiming at the problems of low temperature control efficiency, poor control precision and the like of the Beidou satellite III batch production thermal vacuum test, an intelligent control parameter self-setting technology is adopted, the temperature control targets of all stages are gradually approximated through exponential transition process function curves, the temperature rising and falling rate is accurately controlled and adjusted, the system overshoot is guaranteed to the maximum effect, and the efficiency is improved by more than 90% through the automatic calculation algorithm.

Drawings

Other features, objects and advantages of the present application will become more apparent upon reading of the following detailed description of non-limiting embodiments thereof, made with reference to the accompanying drawings in which:

FIG. 1 is a schematic structural diagram of example 1 of the present application;

fig. 2 is a schematic structural view of the horizontal adjustment mechanism in embodiment 1 of the present application.

Reference numbers:

10. a vacuum vessel; 11. a support frame; 20. a beam frame; 30. a satellite mounting rack; 40. a satellite; 50, an inter-satellite main cooling plate; 60. a lower cooling plate; 70. arranging a lamp array; 80. a lamp array is arranged; 90. a horizontal adjustment mechanism; 91, mounting a base; 92. a push rod; 93. a transmission case; 94. a motor 95, a heating plate; 96. and (5) fixing blocks.

Detailed Description

The present application will be described in further detail with reference to the following drawings and examples. It is to be understood that the specific embodiments described herein are merely illustrative of the relevant invention and not restrictive of the invention. It should be noted that, for convenience of description, only the portions related to the present invention are shown in the drawings.

It should be noted that the embodiments and features of the embodiments in the present application may be combined with each other without conflict. The present application will be described in detail below with reference to the embodiments with reference to the attached drawings.

Example 1

Referring to fig. 1, the present embodiment provides a spacecraft vacuum thermal environment simulation structure, which includes a vacuum container 10, wherein a support frame 11 is disposed at a bottom portion in the vacuum container 10; the cooling system further comprises a beam frame 20 fixed in the middle of the support frame 11, a plurality of satellite mounting frames 30 symmetrically fixed on the beam frame 20, satellites 40 installed on the satellite mounting frames 30, an inter-satellite main cooling plate 50 vertically fixed on the beam frame 20 and located between adjacent satellite mounting frames 30, and a lower cooling plate 60 arranged at the bottom of each satellite 40; the upper surface of each satellite 40 is provided with an upper lamp array 70 and a temperature sensor for measuring temperature, and the lower surface of each satellite 40 is provided with a lower lamp array 80 and a temperature sensor for measuring temperature; the upper array of lamps 70 and the lower array of lamps 80 are used for heating to provide a thermal environment.

The beam type frame adopts a double-layer symmetrical structure and has high-strength bearing capacity.

Wherein, the satellite mounting bracket is L-shaped mounting bracket. The L-shaped support and the satellites are symmetrically arranged on two sides of the inter-satellite main cooling plate 50, and the satellites can obtain relatively consistent thermal background and levelness change in a thermal test by adopting the arrangement mode.

The inter-satellite main cold plate 50 is a horizontal liquid nitrogen cold plate, the lower cold plate is a vertical liquid nitrogen cold plate, and the two cold plates can be used for simulating low background radiation heat flow in the design of a thermal test tool. Because the test tool used in the thermal test of the navigation satellite greatly shields the bottom heat sink of the container, the simulation of the low background radiation heat flow of the two cold plates compensates the heat sink, and the cooling effect of the south plate of the satellite during the test is guaranteed. Wherein the horizontal liquid nitrogen cold plate is positioned between the bottom of the L-shaped mounting frame and the lower lamp array.

In this embodiment, the number of the satellite supports is 2, and the number of the corresponding satellites is 2, namely, the satellite a and the satellite B; during installation, firstly, the satellite A is overturned in place, then a lower cold plate is installed under the satellite A through a tool, and the satellite A and a corresponding satellite support are hoisted into a container; the inter-satellite cold plate is fixedly arranged between the two satellites; b, after the satellite is turned in place, installing a corresponding lower cold plate through a tool; b, the satellite is directly below the satellite, and the satellite support are hoisted into a container finally by analogy; and finally, connecting an internal liquid nitrogen tank circuit.

In other embodiments, the number of satellites may also be set to 4, 6, and other even numbers according to requirements, and when the satellites are arranged, the satellites are arranged in a symmetrical arrangement manner.

Wherein, the upper lamp array 70 and the lower lamp array 80 are infrared lamp arrays and are used as heaters which are controlled by a program-controlled direct-current power supply; the temperature sensor is a thermocouple and is used for sensing the experiment temperature; the upper lamp array 70 is provided with a plurality of control areas, and each controller is provided with a thermocouple for measuring temperature.

Preferably, the present embodiment further comprises a vacuum gauge for measuring pressure within the satellite; the vacuum gauge is installed on a cabin plate of the satellite through a bracket; a cabin plate of the satellite is provided with an opening corresponding to the measuring port of the vacuum gauge; the hole is sealed after the vacuum gauge is fixedly installed, and the hole is sealed by using multiple layers of heat-insulating materials and 3M adhesive tapes after the vacuum gauge is installed, so that heat leakage is prevented.

And a heat insulation gasket is arranged between the support and the cabin plate of the satellite, and the heat insulation gasket is a polyimide gasket.

According to the technical scheme provided by the embodiment of the application, the vacuum gauge is a hot cathode ionization gauge, and the magnetic field of the hot cathode ionization gauge is weaker, so that the harmful influence on equipment on a satellite can not be generated. Wherein, the data of the vacuum gauge are read by devices such as a transmission cable, a vacuum gauge and the like.

When the device is installed, a vacuum gauge is fixedly connected with a support, the support is fixed on a cabin plate of a satellite through a heat insulation gasket and a titanium screw, and a measuring port of the vacuum gauge is opposite to a hole of the cabin plate; then, sealing the opening of the satellite cabin plate by using a plurality of layers of heat insulation materials and a 3M adhesive tape to prevent heat leakage; and finally, connecting the signal control cable with the vacuum gauge joint to test the conduction and insulation performance and ensure that the connection is error-free.

The measuring method of the vacuum gauge has high reliability, realizes the intra-satellite vacuum degree monitoring of the test under all working conditions, and becomes a key criterion for adding degassing working conditions in the thermal test.

Preferably, the present embodiment further comprises a horizontal tilt sensor installed on the satellite, and a plurality of horizontal adjusting mechanisms 90 uniformly arranged at the bottom of the beam frame 20; the horizontal adjustment mechanism 90 is used for adjusting the vertical height of the beam frame according to the signal of the horizontal tilt angle sensor.

As shown in fig. 2, the horizontal adjustment mechanism 90 includes an installation seat 91 that penetrates through the top and bottom and has a top end fixed to the bottom of the beam frame, and a push rod 92 that penetrates through the installation seat 91; the bottom end of the push rod 92 is fixed on the support frame 11 through a fixing block 96;

the push rod 92 comprises a threaded cylinder positioned at the bottom and a threaded rod in threaded connection with the threaded cylinder;

a transmission case 93 is fixed at the bottom of the mounting seat 91; the transmission case 93 is driven by a motor 94 to drive the threaded rod to rotate. The motor drives the threaded rod to rotate through the worm gear structure.

The side wall of the transmission case 93 is provided with a heating sheet 95, and the heating sheet is controlled by a temperature controller to heat. The heating plate 95 heats the transmission case, and the transmission assembly of the transmission case 93 adopts vacuum lubrication, so that the whole mechanism can adapt to a vacuum low-temperature environment.

The horizontal adjusting mechanism is installed and fixed on a beam type frame of the bearing platform after being pre-tightened, and the stroke of the push rod can be dynamically adjusted according to the deformation condition of the beam type frame during the test, so that the levelness of the satellite during the thermal test is finely adjusted.

The beam frame deforms under the action of gravity of the self gravity, the satellite and the thermal test tool, and meanwhile, due to the fact that the masses of the satellites and the test equipment are different, the vertical deformation is inconsistent, the vertical deformation is expressed as the difference of the inclination angles of the horizontal inclination angle sensor, and the inclination angles of 2 orthogonal directions in the horizontal plane of the satellite are commonly used. Therefore, two horizontal tilt sensors are installed on each satellite.

During the thermal test, the heating sheet is controlled by the temperature controller to keep the case at the normal temperature. Meanwhile, the deformation of the beam type frame in the vertical direction can be calculated through conversion of the inclination angle of the horizontal sensor. And controlling the stepping motor to move in real time, and adjusting the push rod to move in the opposite direction of deformation so as to compensate the deformation of the beam type frame and enable the levelness of each satellite to meet the requirement.

For example, in the present embodiment, the number of satellites is 2, each satellite has two tilt sensors, and the number of the leveling mechanisms 90 is 4, and the leveling mechanisms are respectively disposed at four corners below the beam frame; the measurement values of the four horizontal tilt angle sensors are x1, y1, x2 and y2 respectively, and the calculated corresponding tiny deformation of each supporting point of the beam type frame is z1, z2, z3 and z4 respectively; then the stepping motor is controlled to extend or shorten the push rod by corresponding micro displacement so as to achieve the aim of compensating deformation and controlling levelness. For example, a pushrod corresponding to an amount of deformation of z1, elongation-z 1; a pushrod corresponding to a deflection z2, elongation-z 2; a pushrod corresponding to a deflection z3, elongation-z 3; the pushrod, corresponding to the amount of deformation z4, extends to-z 4.

By adopting a high-precision horizontal regulation and control system, the levelness of the two satellites in the test process is superior to 1mm/1m, and the normal work of the satellite heat pipes is ensured.

Example 2

The embodiment provides a method for regulating the temperature of a spacecraft vacuum thermal environment, which is applied to the spacecraft vacuum thermal environment simulation structure in embodiment 1, and comprises the following steps:

dividing each upper lamp array and each lower lamp array into a plurality of temperature control areas, wherein each temperature control area controls a current signal of the temperature control area through a PID arithmetic unit, and the current signal of each controller is determined according to the following steps:

starting timing by taking the experiment starting time as a reference;

acquiring real-time temperature signal data y of the current control period_t(ii) a If the control period is 5s, for example, temperature data is acquired every 5 s;

the set target temperature y for the current control cycle is determined by the following function:

y＝r-Δy+Δy*[1-exp(-t/T)]；

wherein r is the final temperature control target, Δ y is the real-time temperature signal data y_tThe difference from r; t is the acquisition of real-time temperature signal data y_tThe time of day; t is a set time constant;

determining a control increment u (t) of the PID operator according to the following equation:

e(t)＝y-y_t；

wherein K_pIs the set adjustment coefficient.

According to the method, the final temperature control target is gradually approached through the exponential transition process function curve, the temperature rising and falling speed is accurately controlled and adjusted, and the system overshoot is guaranteed to the maximum effect.

According to the technical scheme, the intelligent control parameter self-tuning technology is adopted, the efficiency is improved by more than 90%, and the problems of low temperature control efficiency, poor control precision and the like of the Beidou third satellite batch production thermal vacuum test are solved. The method can be applied to one-device and two-satellite parallel thermal tests of a plurality of groups of MEO satellites.

The above description is only a preferred embodiment of the application and is illustrative of the principles of the technology employed. It will be appreciated by a person skilled in the art that the scope of the invention as referred to in the present application is not limited to the embodiments with a specific combination of the above-mentioned features, but also covers other embodiments with any combination of the above-mentioned features or their equivalents without departing from the inventive concept. For example, the above features may be replaced with (but not limited to) features having similar functions disclosed in the present application.