Acta Polytechnica https://doi.org/10.14311/AP.2022.62.0498 Acta Polytechnica 62(5):498–504, 2022 © 2022 The Author(s). Licensed under a CC-BY 4.0 licence Published by the Czech Technical University in Prague LOW EARTH ORBIT NANOSATELLITE: INFLUENCE OF HEAT DISSIPATION ON PASSIVE THERMAL ANALYSIS Amine Akkaa,∗, Farid Benabdelouahaba, Randa Yerroub a Abdelmalek Essaadi University, Department of Physics, Laboratory of physics and condensed matter, BP. 2121 M’Hannech II. 93030 Tetouan, Kingdom of Morocco b Abdelmalek Essaadi University, Department of Physics, ERSN Laboratory, BP. 2121 M’Hannech II. 93030 Tetouan, Kingdom of Morocco ∗ corresponding author: a_akka@hotmail.com Abstract. The use of small satellites in ambitious missions presents challenges related to thermal breakdowns as one of the critical issues contributing to their failure. Heat dissipation and thermal management are still the major challenges in nanosatellite systems design. To meet the thermal stability requirements, it becomes statutory to manage passive and active thermal control to reach this goal while a variety of factors, such as high-powered components, sunlight and shadow on orbit, or a tight spacecraft layout, remain imposed. A spherical nanosatellite thermal analysis was performed to show the effect of energy dissipation in a low earth orbit and the stability of the system with a special attention to batteries, which persist as the weak link among electronics parts. Additionally, a set of different material coatings was used to demonstrate their impact on the nanosatellite’s thermal behaviour, hence highlighting their importance while designing such a spacecraft. Keywords: Nanosatellite, thermal stability, material coatings, heat dissipation, passive thermal control. 1. Introduction Satellites have always been developed to capitalise on the advantages that provide on all levels of weather monitoring, scientific observation, communication, re- mote sensing, and surveillance. The novelty with nanosatellites is that these favours are acquired at minimal costs [1–4]. Table 1 [5] clearly shows the benefits of the nanosatellite approach as compared to the traditional satellite approach when designing each satellite. The need for a better thermal control on nanosatel- lites with temperature-sensitive components on-board requires many adjustments before launching. A simple shape of nanosatellites will narrow the range of tem- peratures experienced by internal components, and also outer irradiance coming from the Sun and the Earth if absorptance α and emissivity ϵ, which are the primary means of passive thermal control, are well-chosen, depending on the material used [8–14]. The objective is to sustain the temperature of all subsystems within their operating range. As each part of the satellite is coupled to the structure by conduction, the internal temperature is fairly uniform. Therefore, the design of the thermal control system depends on the strictest temperature range, namely the batteries [0 ◦C, 40 ◦C] [15] and operating electronic equipment [−15 ◦C, 50 ◦C] [16]. The purpose of this paper is to establish a passive thermal analysis to ensure optimal operating condi- tions for inner components of a spherical nanosatellite, namely by keeping the temperature within the spec- ified limits. The thermal analysis has been carried out with simulation tools based on the finite element approach for various coating materials, with considera- tion of heat dissipation in steady state conditions and, then also in nominal conditions. Obtained results were very promising, as, in outer space, the possi- bilities offered by a passive or even active thermal control should be considered to overcome the difficul- ties encountered, in particular for the sensitive parts of nanosatellites. 2. Materials and methods 2.1. Space thermal environment The space environment in Figure 1 is very complicated and erratic; therefore, the simulation concerns mod- elling two extreme cases: The hot case, and the cold case. As their name indicate, those describe the most serious situations where thermal loads are relevant as indicated in Table 2 [17]. Since the orbit is Sun- synchronous [18] and circular at a 400-km altitude “98.13° inclination” with spacecraft pointing-earth, the β angle that determines the time during which the spacecraft is exposed to direct sunlight remains almost constant. When considering the Earth and its atmosphere as a whole, the calculation of the rate of absorption of solar energy, and the terrestrial in- frared radiation emission averaged over a certain time 498 https://doi.org/10.14311/AP.2022.62.0498 https://creativecommons.org/licenses/by/4.0/ https://www.cvut.cz/en vol. 62 no. 5/2022 Nanosatellite passive thermal analysis Design approach Mission flexibility System perfor- mance Risk tole- rance Deve- lopment time Cost System focus Traditional/military Low High Low High High Performance Traditional/commercial Low High Low High High Profit Traditional/experimental Low High Medium Medium High Science Nanosatellites Medium Low Medium Medium Medium Cost Table 1. Comparison of traditional and, nanosatellite design approaches. Parameters Cold Case Nominal Case [6] Hot Case Incident Solar Flux [W m-2] 1317 Summer Solstice 1367 1419 Winter Solstice Albedo Factor 0.22 at β = 0° 0.28 0.59 at β = 90° Earth IR [W m-2] 217 242 261 Heat dissipation [W] [7] 6.594 – 22.438 Temperature and Pressure Vacuum at 2.7 K Table 2. Chosen conditions for simulation purposes. Figure 1. Nanosatellite heat exchange. interval may unfold the radiative balance of the Earth with the Sun and outer space [19]. In low Earth orbit (LEO), the altitude is less than the diameter of the Earth and satellites can only see a small part of the Earth at any given time. This means that the conditions will change dramatically as the satellites move through different combinations of environments. These changes must be given priority in the design of the thermal control satellite system. 2.2. Thermal analysis The primary objective of the thermal analysis is to ensure the preservation of the satellite’s internal com- ponents within the specified temperature threshold, especially batteries as mentioned earlier. The steady- state thermal analysis, performed for the nanosatellite in Figure 2, is governed by the equation: Asat × ϵ × σ × T 4 = Qsun + Qalb + Qear + Qint, (1) where on the left side of Eq. (1), Asat [m2] is the satel- lite’s total area emitting radiation, which has the same external area as a Cubesat of 10 cm × 10 cm × 10 cm, ϵ [-] is the emissivity, σ [W m2 K−4] is the Stephan- Boltzmann constant and T [K] is the temperature required, however, on the right side Qsun [W] is the heat input from the solar radiation, Qalb [W] is the heat input from the Albedo radiation, Qear [W] is the heat transferred due to Earth Infrared, and 499 A. Akka, F. Benabdelouahab, R. Yerrou Acta Polytechnica Figure 2. A: Panoramic view, B: Front plane section view, C: Inside view. Qint [W] is the nanosatellite internal heat load, which refers to the dissipated energy from the electronic components and batteries. At this stage of analysis, the absorptivity factor α can’t be remarked in Eq. (1), yet, dissecting Qsun, Qalb, and Qear terms clearly shows the impact of the absorptivity on the equilibrium balance. As a matter of fact, the three quantities are defined as follows, Eqs. (2), (3) and (4): Asun · α · Js = Qsun, (2) where Asun is the projected area receiving solar radi- ation, α is the absorptance factor, and Js is the solar constant. Aalb · α · Jalb = Qalb, (3) where Aalb is the projected area receiving albedo radi- ation, and Jalb is the intensity of the albedo radiation. Aear · ϵ · Jear = Qear, (4) where Aear is the projected area receiving earth radia- tion, and Jear is the intensity of the planetary infrared radiation. As it can be seen in Figure 2, the main component of the nanosatellite, almost totally made of aluminium alloys, is a disc panel, where different electronic cards may be installed, and which is not dissipating power in the upcoming simulation. The disc panel is mounted in the equator enclosed by batteries, the most power- dissipating elements, that are supported by a double tube in the axis of spinning. Coating α ϵ Black Body 1 1 White Paint V200 0.26 0.89 Black Paint H322 0.96 0.86 Brilliant Aluminum Paint 0.70 0.13 Buffed Aluminum 0.16 0.03 Blue Anodised Titanium Foil 0.30 0.31 Table 3. Used coatings for the nanosatellite [23]. It is always accurate to process a perfect thermal, mechanical and electrical design of useful loads at the very beginning of the design process to avoid anoma- lies that may occur due to the details of the payload packaging. Indeed, there are many challenges that engineers face when designing a spacecraft, namely thermal ones that have made the subject of several research papers for the optimisation of such an analyt- ical approach [20] and even experimental testing [21]. The passive thermal analysis, shown in Table 3, con- cerned multiple coating materials, in extreme condi- tions with heat dissipation, but also in nominal condi- tions with no heat dissipation as mentioned in Table 2. The temperature is calculated by finite element code in accordance with all the boundary conditions [22]. 500 vol. 62 no. 5/2022 Nanosatellite passive thermal analysis Figure 3. Coating effect on temperature in nominal conditions. Coating α ϵ Temp [◦C] White Paint V200 0.292 −40.8 Brilliant Aluminum Paint 0.967 17.9 Black Body (No coating) 1 19.9 Black Paint H322 1.116 26.8 Buffed Aluminum 5.333 158.2 Blue Anodised Titanium foil 5.384 164.4 Table 4. Temperatures in nominal conditions without heat dissipation. Coating α ϵ Min Temp [◦C] Max Temp [◦C] White Paint V200 0.292 13.4 15.4 Brilliant Aluminum Paint 0.967 55.1 57.0 Black Body (No coating) 1 56.7 58.7 Black Paint H322 1.116 62.4 64.3 Buffed Aluminum 5.333 186.5 188.3 Blue Anodised Titanium foil 5.384 186.8 188.6 Table 5. Min and Max temperatures for cold case simulation. 3. Results and discussion 3.1. Nominal case results Referring to the temperature of the batteries, which should remain within the predefined ranges to main- tain the proper functioning of the nanosatellite, we hardly notice that, under the nominal working condi- tions (Table 4 and Figure 3), three of the six coatings meet the above condition and allow the batteries’ to run at acceptable temperatures, namely [17.887 ◦C, 19.901 ◦C, 26.848 ◦C]. For all the results obtained, the temperature of the batteries was the highest and the simulation made it possible to clearly identify the problem and try to start on a good basis when proposing solutions. All these coatings have an α ϵ ratio around unity. For the report of the other coatings, it is obvious that other passive thermal controls should be con- sidered, namely heat-conducting elements, adiabatic spacers, modifying the geometry of the spacecraft, or even active thermal control, but one must keep in mind that the latter should only be used when it is impossible to meet the requirements. 3.2. Cold case results For the cold case simulation, Table 5 and Figure 4, it is clear that “White paint V200” perfectly follows the temperature required for the proper functioning of the nanosatellite, the three ratios of coatings that follow need only a complement of passive control to fall within the optimal operating temperature range of the batteries. However, for “Buffed Aluminum and Blue Anodised Titanium” coatings, an active thermal control is required because the temperatures have exceeded the limits. 3.3. Hot case results Finally, for the hot case, Table 6 and Figure 5, the obtained results show that under the conditions de- fined at the outset, a temperature exceeding 200 ◦C is reached, and therefore another type of coating mate- rial must be applied in addition to an active thermal control to decrease the temperatures and ensure opti- mal operating conditions for the spacecraft. 501 A. Akka, F. Benabdelouahab, R. Yerrou Acta Polytechnica Figure 4. Coating effect on temperature in cold case conditions with heat dissipation. Figure 5. Coating effect on temperature in hot case conditions with heat dissipation. Coating α ϵ Min Temp [◦C] Max Temp [◦C] White Paint V200 0.292 80.1 86.6 Brilliant Aluminum Paint 0.967 103.6 109.9 Black Body (No coating) 1 104.6 110.9 Black Paint H322 1.116 108.2 114.5 Buffed Aluminum 5.333 199.9 205.8 Blue Anodized Titanium foil 5.384 201.0 206.9 Table 6. Min and Max temperatures for hot case simulation. 3.4. Batteries issues As it can be seen in Figure 6 and Figure 7 taken from a cold and a hot analysis, the main cause of the high gradient of temperature is the battery “In red”, It is noteworthy that for an operational nanosatellite in the harsh conditions of the outer space, one should think about using all possibilities offered by passive thermal control, or even active thermal one, to overcome the encountered difficulties. In addition to this, we can undoubtedly notice that the temperature distribution at the level of the differ- ent elements of the nanosatellite is almost the same in both extreme cases [24]. It is important to remind that only the dissipation of heat caused by the batteries has been involved, yet if other major parameters are involved in the system with the complexity of their heat dissipation, results will vary greatly. 502 vol. 62 no. 5/2022 Nanosatellite passive thermal analysis Figure 6. Cold case internal temperature distribution α ϵ = 0.29. Figure 7. Hot case internal temperature distribution α ϵ = 5.38. 4. Conclusions Consistent design criteria for the development and comparison of several material coatings associated with heat dissipation parameters were presented. This paper went through a passive thermal analysis of a simply designed spherical nanosatellite. Indeed, it was assumed that the design provided has shown the ef- fect of batteries’ heat dissipation which imposed some changes in temperature plots. Even though, there is an extensive range of material coatings that can offer various α ϵ ratios, practical selection of the type of coating is often limited by the ageing characteris- tics. That’s why one should think about using all the offered possibilities by finite element codes even those of active thermal control, which can be the subject of an extensive research study and may serve as an additional resolution tool for the current space revolu- tion. An extended study must be performed to gauge the viability of that kind of thermal control, which may improve the quality and then the stability of that kind of nanosatellite. List of symbols α Absorptivity [–] β Beta Angle [°] ϵ Emissivity [–] σ Stephan-Boltzmann constant [W m2 K−4] T Temperature [K] Qsun Heat input from the solar radiation [W] Qalb Heat input from the Albedo radiation [W] Qear Heat transferred due to Earth Infrared [W] Qint Nanosatellite dissipated energy [W] Asat Satellite total area emitting radiation [m2] Asun Projected area receiving solar radiation [m2] Aalb Projected area receiving albedo radiation [m2] Aear Projected area receiving earth radiation [m2] Js Solar constant [W m−2] Jalb Intensity of the Albedo radiation [W m−2] Jear Intensity of the planetary infrared radiation [W m−2] 503 A. Akka, F. Benabdelouahab, R. Yerrou Acta Polytechnica References [1] K. Woellert, P. Ehrenfreund, A. J. Ricco, H. Hertzfeld. 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In press. 504 https://doi.org/10.1016/j.asr.2010.10.009 https://doi.org/10.2514/6.2015-1835 https://digitalcommons.usu.edu/smallsat/2000/All2000/32 https://digitalcommons.usu.edu/smallsat/2000/All2000/32 https://doi.org/10.1109/AERO.2013.6497393 https://doi.org/10.2514/1.44468 https://ntrs.nasa.gov/citations/20020004360 https://doi.org/10.1016/S0094-5765(97)00083-0 https://doi.org/10.17654/HM021010133 https://doi.org/10.17654/0973576321011 https://doi.org/10.1016/j.actaastro.2004.09.003 https://doi.org/10.1016/j.actaastro.2006.07.001 https://doi.org/10.1016/j.actaastro.2015.05.012 https://doi.org/10.3390/en13164097 http://hdl.handle.net/2014/37900 https://doi.org/10.4028/www.scientific.net/AEF.35.29 https://doi.org/10.4028/www.scientific.net/AEF.35.29 https://doi.org/10.1109/AERO.2019.8741754 https://doi.org/10.1051/matecconf/20165409001 https://doi.org/10.1051/e3sconf/202233600057 Acta Polytechnica 62(5):498–504, 2022 1 Introduction 2 Materials and methods 2.1 Space thermal environment 2.2 Thermal analysis 3 Results and discussion 3.1 Nominal case results 3.2 Cold case results 3.3 Hot case results 3.4 Batteries issues 4 Conclusions List of symbols References