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Thermal Analysis of a Small Satellite in Post- Mission Phase Journal of Multidisciplinary Engineering Science and Technology (JMEST) ISSN: 2458-9403 Vol. 6 Issue 2, February - 2019 Thermal Analysis of a Small Satellite in Post- Mission Phase Ahmed Elhefnawy Ahmed Elweteedy Space Technology Center Military Technical College Cairo, Egypt Cairo, Egypt [email protected] [email protected] Abstract— The objective of satellite thermal perigee altitude of 354 km, an apogee altitude of 1447 control is maintaining all satellite components km and an inclination of 71˚. A passive thermal control within their operational temperature range for system was used, with a small electric heater for the different missions. Most small satellites are battery. Thermal models were developed within both designed to have passive thermal control system. MATLAB/Simulink and ESATAN/ESARAD In this paper the thermal analysis of a small environments. satellite in post-mission phase is introduced. The Dinh [4] illustrated the modeling of a nanosatellite European Student Earth Orbiter (ESEO) satellite using Thermal Desktop software. The satellite has a was chosen in this study. This phase starts at the passive thermal control system. It was designed to end of the operational phase and should last at operate in LEO with altitude of 400 km. least 2 years for the chosen satellite. A finite Bulut et al [5] described the thermal control system difference thermal model using a commercial of Turksat-3U Nanosatellite [6]. The satellite orbit is software “Thermal Desktop” was described for the sun-synchronous orbit with altitude 600 km and selected satellite. Model results verification were inclination 98˚. The satellite has passive thermal performed by varying design variables and seeing control system. The thermal model was built in Therm- whether component temperatures increased or XL. decreased as expected. A sample of one component temperature variation on each panel The passive thermal control system of OSIRIS-3U was chosen and introduced to be compared with satellite was discussed by Garzon [7]. The satellite upper and lower bounds on temperature orbit has altitude 400 km and inclination 51.6°. A time- predictions. The results concluded that all satellite dependent finite-element model of the temperature variations of OSIRIS-3U was created in COMSOL components operate properly within their Multiphysics using the heat transfer module. specified temperature limits. Poucet [8] presented the thermal control system of Keywords— Thermal control system; Thermal ESEO satellite. The mission was based on circular analysis modeling; Modeling tool. sun-synchronous LEO of 520 km in altitude and 97.48° for inclination. Finite difference modeling was I. INTRODUCTION completed using ESATAN software package. The The thermal control system on a satellite generally satellite has active thermal control system. uses two basic approaches for temperature Silva et al [9] introduced the thermal control system management: passive and active thermal control. of Amazonia-1 satellite. The thermal control design Many satellite thermal control systems use a was supported by thermal analysis using Thermal combination of passive and active control, though the desktop software package. The target orbit was sun- passive control methods make up the majority of the synchronous orbit with altitude 752.4 km and an system with supplemental active control methods for inclination 98.405°. The satellite active thermal control equipment with small temperature tolerances [1]. Small system was based on heaters regulated by software satellites most commonly employ passive control via thermistors. methods. Passive thermal control techniques include material property selection, controlling the path of heat Appendix A summarizes the university-class transfer, and using insulation systems to ensure that missions for small satellites from 70’s up to 2017[10], temperatures remain within acceptable limits [1]. [11] Studies in small satellite thermal control analysis This paper describes the analysis of the thermal have been conducted by several researchers. Czernik control system for the “ESEO” satellite [8] in one of its [2] discussed the thermal control methods for mission phases. It has five main phases: launch and Compass-1 satellite. It is designed for a circular, sun- early operations phase, operational phase, extended synchronous Low Earth Orbit (LEO) with an altitude of phase, post-mission phase, and satellite disposal. The 600 km and an inclination of 98°. Its mission duration analysis is focused on post-mission phase. This phase is six months. passive thermal control was used and starts at the end of extended phase (if performed) or at ANSYS software has been chosen for modelling this the end of the operational phase. A finite difference spacecraft. thermal model using a commercial software “Thermal Desktop” (TD) was described for the selected satellite. The thermal control system of OUFTI-1 satellite Model results verification were performed by varying was presented by Jacques [3]. The satellite orbit has a www.jmest.org JMESTN42352826 9495 Journal of Multidisciplinary Engineering Science and Technology (JMEST) ISSN: 2458-9403 Vol. 6 Issue 2, February - 2019 design variables and seeing whether component Once the above-mentioned data are inserted in the temperatures increased or decreased as expected. model, the simulation is utilized to predict the temperatures that the satellite components will II. SATELLITE THERMAL ANALYSIS experience. The satellite component temperatures ESEO has a cuboid shape with six structural panels results are then compared to the specified limits. For a and three solar panels of which two are deployable component whose temperatures do not comply with its and one fixed. The main dimensions are within a limits, alternatives to rectifying the problem may be cuboid of 967 × 750 × 680 mm and the total mass of examined, such as applying different surface coatings satellite is less than 100 kg. Satellite systems / or adding insulation to the satellite model. In the components are given in Table 1. following section the model results are presented. To accurately determine the satellite components III. RESULTS AND DISCUSSION temperature distribution, a finite difference model of Fig. 1 through Fig. 8 show the temperature the spacecraft was created. The model was developed variation for post-mission phase as predicted by TD. by using TD software package. First, the configuration The results show that all equipment normally operate of the spacecraft as geometry, materials, and within specified limits. In these figures, the temperature placement of equipment must be defined. Then, the responses are repeated every orbit as a periodic. The boundary conditions are completely described. The maximum temperature occurs when the satellite boundary conditions are comprised of the spacecraft’s exposed to all three external heat inputs, solar, albedo orbit, attitude, thermal environment, and internal heat and earth emission. After that, the satellite enters the dissipation. The temperature requirements of the eclipse and the temperatures of the entire satellite spacecraft components should be specified. begins to decrease due to the loss of heat input from ESEO orbit leads to thermal environment with solar the solar and albedo sources. When the satellite exits flux of 1371 W/m2, albedo 0.3, and earth infra-red of the eclipse, the temperature starts to raise again. 2 237 W/m [8]. The total heat dissipation in post-mission Fig. 1 shows the transient temperature response for phase for all equipment is 155.28 W. Operating the three solar panel in four orbits. The temperature temperature limits are usually determined by variation (between maximum and minimum manufacturer. The structural panels and the internal temperature levels) of solar panel +X is 75 °C that is equipment have range (-20 to +40) and solar panels smaller than the variation in temperature experienced range (-100 to +100). by solar panels +Z and -Z, 100 °C. This variation for solar panel +X because of its fixation on structural The thermal control system includes material panel 1 (no incident albedo and earth infra-red on its property selection, controlling the path of heat transfer, bottom side). All solar panels operate within their using insulation systems, and radiators. The margins (-100 to +100 °C) conductive paths are controlled by using thermal fillers between the panels with thermal conductivity of 1.3 The temperatures variation for the six structural W/m.K [1]. Most of satellite panels were covered with panels are depicted in Fig. 2. All panels temperature multi- layer insulation. Panels 2 and 5 are used as levels at 25 °C as maximum and 16 °C as minimum radiators which have emissivity of 0.88, 0.6 and temperature. As seen, all panels normally operate absorptivity 0.26, 0.14 respectively. within their margins (-40 to +85 °C). TABLE 1 ESEO SYSTEMS / COMPONENTS No Equipment System No Equipment System 1 AMSAT Payload 5 Langmuir plasma Payload diagnostic probe (LMP) 2 Tridimensional Payload 6 Electric Power System Electric power Telescope dosimeter control unit (EPS PEB) system (Tri-Tel S) Battery 3 Telemetry and Communication 7 Reaction Wheel Attitude Telecommand system system determination and (TMTC) Magneto Torquer control system Telemetry and Gyro Telecommand Antenna (TMTC Magnetometer Antenna) Star Tracker 4 Micro camera (uCAM) Payload 8 On Board data On Board data Handling (OBDH) Handling system www.jmest.org JMESTN42352826 9496 Journal of Multidisciplinary Engineering Science
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