Thermal Control Design Research Articles

Thermal Modelling of a Remote Sensing Satellite Panel in an Inclined Orbit

Satellites consist of different subsystems; one of them is the thermal control subsystem (TCS), which warranties the specified temperature settings of other subsystems and devices through satellite lifetime. Satellite Thermal modelling is performed by solving thermal budget equation, taking into consideration maximum and minimum external fluxes, and heat rejection from internal devices which fixed on the internal surface of the panel. To reach the optimum design for the thermal control of the panel (radiation surface areas and power of the electric heaters), the thermal analysis results should meet the design requirements, and the temperature ranges of each device or subsystem inside the satellite. Multilayer Insulation (MLI) is one of the most important passive elements of thermal control subsystem covered the satellite as a blanket consists of some layers from thin pressed Mylar or Kapton sheets. MLI is important to minimize the heat exchange between the inside (devices heat dissipation), and outside satellite (external heat fluxes). A parametric investigation is presented for an inclined satellite in Low Earth Orbit (LEO) at 650 km altitude, to study MLI covering area effect on panel thermal control design. Satellite thermal analysis is performed by Thermal desktop/SINDA FLUINT Software by decreasing and increasing MLI on the outer surfaces of the satellite, to ensure that the satellite electrical equipment temperatures maintained in the required ranges for normal operation.

Thermal control technology for the space station adjoint modular satellite based on new thermal control materials

A modular satellite could perform various functions with various modules and the function of the satellite could be expanded easily. Therefore, the modularization design is an important development direction of space station adjoint satellites and others. However, every module of a modular satellite is structurally independent and the heat dissipation is unevenly distributed, while the interface between modules should support repeatable connection-separating. The traditional thermal control design could not satisfy the thermal control demand. In this paper, the thermal control technology for the space station adjoint modular satellite based on the thermal interface of carbon nanotube array on copper substrate, graphene coating and smart thermal control coating is proposed. By using the new technology, the thermal connection of the assembly and reconstruction system is built and the synergistic heat dissipation of the whole satellite is achieved. As to validate the proposed technology, the finite element model of the space station adjoint modular satellite is established and the whole flight process is simulated. The result indicates that the thermal control technology proposed in this paper can satisfy the thermal control demand of the modular satellite.

Design and simulation of hybrid thermal energy storage control for photovoltaic fuel cell

Objective: Through the design and simulation of hybrid thermal energy storage control of photovoltaic fuel cell, the hybrid thermal energy storage system of photovoltaic fuel cell is further optimized. Method: Firstly, the mathematical model of photovoltaic power generation is established. Then voltage feedback, power feedback, disturbance observation method and conductance increment method are used to track the maximum power of the system. After that, the dynamic model of proton exchange membrane fuel cell is established, and the former maximum power point tracking control strategy is used to keep the voltage stable. Finally, simulation experiments are carried out to verify the effectiveness and superiority of the proposed control strategy and battery model. Results: The hydrogen pressure on the anode side of the fuel cell can be maintained at 0.3 MP at a fast speed. In the process of output, the voltage of fuel cell is much smaller than the polarization voltage of fuel cell. Its voltage decreases gradually from 14 seconds to 16 seconds. Once the illumination changes suddenly, the system can also accurately locate and track the maximum power point, and output the electric quantity. Conclusion: Based on the mathematical model of photovoltaic power generation and the dynamic model of proton exchange membrane fuel cell, the hybrid thermal energy storage system of photovoltaic fuel cell has great advantages. It can keep the voltage stable and track the maximum power of the system in time, which is of great significance for the follow-up research in photovoltaic power generation.

PRELIMINARY THERMAL CONTROL DESIGN ANALYSIS OF LAPAN SAR-MICROSATELLITE DEPLOYABLE SOLAR ARRAY PANEL USING ONE NODAL METHODE

LAPAN SAR-Microsatellite is the first LAPAN micro satellite being developed and planned to carry a Synthetic Aperture Radar (SAR) payload. Different from optical satellites that have been developed, LAPAN SAR-Microsatellite requires a lot more power. Solar array panels are needed to generate solar radiation into electrical energy which is used by all of subsystem satellite as energy to turn on and turn off all of components. More larger the area of solar array panel, more greater to the energy obtanied. Therefore, the needed of deployable solar array panel is a must caused by the dimension of the main body structur (MBS) are not large enough. Solar cells will experience a decrease in efficiency if they experience excessive heat. If there is a decrease in efficiency in the solar cell, it will have an impact on the decrease in power produced. The purpose of this study is to conduct thermal design on solar panels to maintain temperature according to their working temperature. Calculations are carried out on the temperature of solar array panel with various optical properties in the upper and lower panel that best which will be use using one nodal analysis methode According to the calculation result from all of the designs passive thermal control system on this study show that Design-5 is the best thermal control design which can be used for LAPAN SAR-microsatellite deployable solar array panel which α=0.24 at the top and ε=0.9 at the bottom surface of solar array panel.

Analytical Study of Tri-Generation System Integrated with Thermal Management Using HT-PEMFC Stack

Recently, extensive studies on power generation using clean energy have been conducted to reduce air pollution and global warming. In particular, as existing internal combustion engines lose favor to power generation through hydrogen fuel cells, the development of tri-generation technology using efficient and reliable fuel cells is gaining importance. This study proposes a tri-generation thermal management model that enables thermal control and waste heat utilization control of a high-temperature PEMFC stack that simultaneously satisfies combined cooling, heating, and power (CCHP) load. As the high-temperature PEMFC stack operates at 150 °C or more, a tri-generative system using such a stack requires a thermal management system that can maintain the operating temperature of the stack and utilize the stack waste heat. Thus, to apply the waste heat produced through the stack to heating (hot water) and absorption cooling, proper distribution control of the thermal management fluid (cooling fluid) of the stack is essential. For the thermal management fluid control design, system analysis modeling was performed to selectively design the heat exchange amount of each part utilizing the stack waste heat. In addition, a thermal management system based on thermal storage was constructed for complementary waste heat utilization and active stack cooling control. Through a coupled analysis of the stack thermal management model and the absorption cooling system model, this study compared changes in system performance by cooling cycle operation conditions. This study investigated into the appropriate operating conditions for cooling operation in a tri-generative system using a high-temperature PEMFC stack.

Investigation of self‐adaptive thermal control design in passive direct methanol fuel cell

AbstractA novel anode thermal optimization approach is presented and developed in this paper, based on the detailed investigations of operating temperature on the performance of passive micro direct methanol fuel cell (μDMFC). First, a two‐dimensional model of two‐phase cell is established to observe methanol solution transport and pressure distribution under different operating temperatures. Moreover, passive stainless steel‐based μDMFC preparation and fabrication use micro laser‐cut technology, and the influences of cell temperature on corresponding cell performances are under experimental investigation. Eventually, according to the consequences of simulation and experimental analysis, a practical anode self‐adaptive thermal control strategy equips with the self‐made silicone heating sheet is designed, and the stable work of μDMFC at the optimal temperature is realized through this method. Compared with the conventional fuel cell, the optimized fuel cell system can significantly improve the net output power density and efficiency (Power density increased from 20.8 mW cm−2 to 52.16 mW cm−2).

Study on Thermal Distortion and Thermal Design Optimization of Satellite Antenna Reflector during Orbit Period

Reflector thermal distortion status is essential to performance stability of communication satellite antenna in geosynchronous orbit, and should be carefully validated in design period. Thermal distortion simulation of current satellite antenna reflector is usually conducted by numerical calculation of extreme temperature field condition from thermal analysis results. However, telemeasuring result shows that, the worst performance time of antenna is not usually under these extreme temperature conditions, and it is necessary to analyze the thermal distortion features of satellite antenna during the whole orbit period to provide efficient instructions for design of the antenna thermal control.

Thermal design and analysis of the high resolution MWIR/LWIR aerial camera

The thermal control technology of high resolution MWIR/LWIR aerial camera is studied. Thermal design and thermal analysis of the camera are completed according to thermal control indexes of infrared aerial camera. Firstly the architecture of the aerial camera and the external working environment are introduced, the camera heat exchange model is established, and the boundary conditions of radiation, conduction, convection and aerodynamic heat are analyzed, and then the thermal design of the camera is described in detail. Finally the thermal analysis of aerial camera is done, and a heat balance test has been done for the camera. The results show that: In the specified environment, after two hours work the temperature range of optical system is 18.5 ℃ to 22.2 ℃, the temperature range of lens set is 19.1 ℃ to 20.3 ℃, the temperature range of detector module is 19.7 ℃ to 31.9 ℃, meet the thermal control indexes, at the same time, under extreme operating conditions, the thermal control system can meet the start-up requirements, the thermal design is feasible and reasonable. This research method and technical route have the guidance and reference significance for the thermal control design of other infrared aerial cameras.

Preliminary design and comparative study of thermal control in a nanosatellite through smart variable emissivity surfaces

The thermal radiation that is rejected or absorbed into deep space is highly variable. Ultralight smart surfaces with arrays of unit cells can be designed to change their effective emissivity and absorptivity without energy consumption, actuators, and controllers, and can be used for the temperature control of satellites. The smart surfaces work in a similar manner to thermal louvers but they are hingeless, lighter, and their activation depends on their anisotropic mechanical properties and multilayer structure. The generated thermal stresses between layers that have a high mismatch in the coefficient of thermal expansion cause large deformations and rotations within small temperature changes. The arrays of the surface open or close, and transform their geometry as a function of temperature; therefore, coatings of different thermo-optical properties are revealed or concealed, thus creating variable emissivity surfaces. The emissivity and absorptivity curves of the smart surfaces can be entirely designed as a function of temperature. Theoretically, an emissivity change equal to Δε = 0.8 can be achieved. The small thermal capacitance renders nanosatellites very susceptible to temperature fluctuations. In this study, different emissivity curves were generated to re-calculate the worst cold and hot cases, and to redesign the thermal control system of a certain nanosatellite. We studied a plethora of design cases based on the energy balance equation in steady state while considering the nanosatellite as one-node geometry. In two ideal designs, the temperature deviation of the nanosatellite in the worst cold and hot cases is limited to Δ Τ = 37 ℃ or 43 ℃ without the use of heaters. Moreover, with a power equal to 0.7 W the temperature deviation is limited to Δ Τ = 20 ℃. Consequently, the thermal fatigue is minimized and the energy consumption during the eclipse phase is reduced.

Thermal control design and experimental verification of light off-axis space optical remote sensor in the sun-synchronous orbit

Thermal control design and experimental verification of light off-axis space optical remote sensor in the sun-synchronous orbit

A Numerical Study of Fluid Flow and Heat Transfer in Carbon Dioxide Enclosures on Mars

In order to support the future thermal control and energy conservation design for the Mars rover, numerical studies on natural convection in CO2 enclosures on Mars’ surface were conducted for both horizontal and vertical enclosures. The parameters are as follows: the atmospheric pressure was 1000 Pa, the gravitational acceleration was 3.62 m/s2, and the Prandtl number was 0.77. The heat flux, temperature, and velocity fields of the CO2 enclosures were obtained with the aspect ratio ranging from 5.56 to 200 and the Grashof number ranging from 430 to 2.6 × 104. It was found that natural convection formed more easily in the horizontal enclosures than that in the vertical enclosures when the enclosures had same thickness. With the increasing thickness of the enclosures, Rayleigh–Bénard convections formed in the horizontal enclosures, while only single-cell convections formed in the vertical enclosures. The heat flux through the horizontal enclosures was greater than that through the vertical enclosures with the same thickness when natural convection formed. The maximum difference between them reached 35.26%, which was illustrated by the field synergy principle. A hysteresis phenomenon of the natural convection dominating the heat transfer was found in the vertical enclosure on Mars’ surface. New values for the critical Grashof number and correlations for the average Nusselt number for both the horizontal and vertical CO2 enclosures on Mars’ surface were also developed.

Research on Hot-Wire Low Pressure Wind Speed Measurement System

In order to verify the correctness of the thermal control design, the Mars rover needs to conduct a thermal test on the ground during the construction of the Mars exploration mission. In the course of the thermal test, real-time measurement of the wind speed at low pressure is required. Therefore, it is needed to establish a low-pressure wind speed measurement system. This article describes the composition of the test system, the choice of measurement sensors, in which the impact of wire resistance, electromagnetic interference and other factors are taken into account, and the design of the test process is completed.

The effect of temperature characteristic variation due to thermal control design on the antenna of satellite laser communication terminals

Cassegrain telescopes are widely used as optical antenna in satellite laser communication terminals, which are assembled outside the satellite when operating in orbit. The thermal environment in space will result in temperature variation and thermal deformation of the telescopes. Then tracking-pointing performance and communication performance will be affected consequently. So, thermal control design of telescope must be taken to make the temperature distribution vary in a certain range, which is permitted to assure the performance and stability of optics. For laser communication terminals in satellite of geosychronous Earth orbit, the effects on figure error of telescope due to three kinds of heating type of thermal control system are analyzed in this paper. It is shown that radial heating type brings the smallest figure error to telescope, followed by circumferentially-equal-area heating type. circumferentially-equal-radius heating type brings the biggest error of all. The result of this paper will bring certain reference to the thermal control design of telescopes in laser communication terminals on satellites in orbit.

Thermal conductance of one-dimensional materials calculated with typical lattice models.

We show through calculations on typical lattice models that thermal conductance σ can well describe the near-equilibrium thermal transport property of one-dimensional materials of finite length, which presents a situation often met in the application of nanoscale devices. The σ generally contains contributions from the material itself and those from the thermal reservoirs. The intrinsic σ of the material, i.e., the one with the fewest external influences, can be efficiently calculated with the help of the "blackbody"-like nonreflective thermal reservoir, either through the nonequilibrium method or through the Green-Kubo-type formula. σ thus calculated would be helpful to guide the design of thermal management and heat control in nanoscale devices.

In-Flight Performance of the OCO-2 Cryocooler

The Orbiting Carbon Observatory-2 (OCO-2) will have completed its first year in space on July 2, 2015. The OCO-2 instrument incorporates three bore-sighted, high-resolution grating spectrometers, designed to measure the near-infrared absorption of reflected sunlight by carbon dioxide and molecular oxygen. The cryocooler system design is coupled with the instrument's thermal control design to maximize the instrument's performance. A single-stage NGAS pulse tube cryocooler provides refrigeration to three focal plane arrays to ∼120 K via a high conductance flexible thermal strap. A variable conductance heat pipe (VCHP) based heat rejection system (HRS) transports waste heat from the instrument located inside the spacecraft to the space-viewing radiators. The HRS provides tight temperature control of the optics to 267 K and maintains the cryocooler at 300 K. Soon after entering the A-Train on August 3, 2014, the optics and focal planes were cooled to their operating temperatures. This paper provides a general overview of the cryogenic system design and reviews the in-flight cryogenic performance during the Observatory's first year.

Thermal Conduction Measurement and Analysis for a Spaceborn Sensor

The present study is aimed at investigating the thermal conduction characteristics of a spaceborn sensor experimentally and computationally. The experimental measurement has been carried out in a high vacuum chamber with twenty selected thermal couples attached to the sensor module. The detailed internal thermal control design, material for heat release, temperature limit, and control logic of the sensor assembly have been described in the study. The total thermal resistance for the sensor has been evaluated from the experimental results of the steady state temperature distributions under a heat power of 5.5 W and the sensor heater power modulation has been obtained to keep the PT1000 within the accepted temperature range. Under the assumed temperature boundary conditions, the measured thermal resistances and the modulated sensor heater power, the maximum PT1000 temperature distribution is found to be 25.26 °C ± 3.44 °C with the finite element analysis. Heat loss has been evaluated for the invar mount of the sensor module. In addition, the overall orbit peak and average heater powers needed are 4.61 W and 1.61 W respectively.

Correlation of thermal mathematical models for thermal control of space vehicles by means of genetic algorithms

The design of the thermal control system of space vehicles, needed to maintain the equipment components into their admissible range of temperatures, is usually developed by means of thermal mathematical models.These thermal mathematical models need to be correlated with the equipment real behavior registered during the thermal test campaign, in order to adapt them to the real state of the vehicle “as built”. The correlation of this type of mathematical models is a very complex task, usually based on manual procedures, which requires a big effort in time and cost. For this reason, the development of methodologies able to perform this correlation automatically, would be a key aspect in the improvement of the space vehicles thermal control design and validation.The implementation, study and validation of a genetic algorithm able to perform this type of correlation in an automatized way are presented in this paper. The study and validation of the algorithm have been performed based on a simplified model of a real space instrument. The algorithm is able to correlate thermal mathematical models in steady state and transient analyses, and it is also able to perform the simultaneous correlation of several cases, as for example hot and cold cases.

Thermal Control Design for the Combination of Star Sensor Using C-C Graphite Material

To solve the problem that thermal control design of the combination of star sensor is high requirement and difficult to implement, the thermal control design of the combination of star sensor using C-C graphite material was proposed in this paper. By taking some combination of star sensor requiring difference in temperature on bracket of star sensor less than 4.5°C and temperature of flange of star sensor in 8.5±2°C as an example. Thermal analysis was achieved to validate the thermal control design scheme proposed in the paper. Analysis results show that the thermal control design scheme can make difference in temperature of bracket of star sensor in orbit less than 2.73°C, and flange of star sensor in orbit between 7.47°C and 8.93°C,which meet the temperature requirement very well. The design scheme is reasonable.

Thermal Control Design Conception of the Amazonia-1 Satellite

Amazonia-1 is a Brazilian remote sensing satellite providing mainly images, in order to observe and monitor deforestation, especially in the Amazon region. This paper describes the thermal control design, which uses passive and active concepts. The active thermal control is based on heaters regulated by software via thermistors.The passive thermal control consists of multi-layer insulationblankets and radiators, paints, surface finishes to maintaintemperature level of the overall carrier components within an acceptable value. The thermal control design is supported by thermal analysis using thermal mathematical model. The temperatures and heater power are predicted for critical cases.

皮卫星及其星箭分离机构的热-结构耦合与间隙设计

The thermal-structure coupling for a pico-satellite and its separation mechanism was analyzed to guide a rational design of the fit clearance value between the satellite and guide rail and to ensure safe and reliable satellite separation.On the basis of the geometry relations of the satellite,the earth,and the sun in the process of launching,the internal and external heat flows absorbed by the pico-satellite and its separation mechanism were analyzed.The temperature field and thermal deformation in satellite launching were obtained by finite element analysis.The clearance fit value between the satellite and guide rail was designed on the basis of above analysis.The analysis results show that the maximum deformation occurs on the shady side because of the cold dark environment and conduction heat,and the pico-satellite shell and its guide rail surface are not distorted and deformed seriously.The amount of maximum deformation is 62.3μm when the preload force is 600N and that is 63μm when the preload force is 300N.Based on above analysis results,the fit clearance was selected.These results demonstrate that the pico-satellite and its separation mechanism can separate normally under the high and low temperature tests.The on-orbit separation attitude is in agreement with simulation results well.All of these verify that the fit clearance between the satellite and its guide rail is reasonable and it can be used in the structural design and thermal control design of pico-satellite separation mechanisms.