The basic purpose of thermal design is to maintain the temperature of all spacecraft components within desired limits. We also wish to minimize the temperature fluctuation (thermal cycling) that the spacecraft components are subjected to. FalconSat-2’s internal components, which are the most thermally sensitive parts of the satellite, are fairly thermally decoupled from the external heat flux the satellite is subjected to. This is due to the design with the inner column and outer structural shell. This allows us to control the temperature with a passive thermal design approach. We will modify the thermo-optical properties (absorptivity and emissivity values) of the external facets of the satellite so that the satellite and all components are maintained within the optimal temperature range.
On FalconSat-2, the operational temperatures are limited by the electronic components within the satellite, and specifically by the battery. The battery is the most thermally sensitive of the satellite subsystems because it cannot be recharged below 0˚C. As a result, the nominal temperature range targeted for the batteries and internal components of FalconSat-2 is +5 to +30 deg C. The other commercial electronics within the satellite have temperature limits of –40 and +85 deg C. The structural components and solar panels have much more relaxed temperature limits. Table 1 lists the temperature limits for FalconSat-2.
Table 1 – Temperature limits for FalconSat-2 subsystems
To design the thermal subsystem and ensure that FalconSat-2 will meet these temperature limits, we had to first simulate the thermal behavior of the satellite. This will allow us to see how the satellite will behave without any thermal control implemented, which will in turn show us what design we must implement to meet the temperature range requirements. In order to simulate the satellite’s thermal behavior, a model had to be created.
We require a detailed thermal model of FalconSat-2 for several reasons. Primarily, we need to simulate expected on-orbit thermal behavior of the satellite and ensure that no spacecraft components exceed their maximum or minimum temperature limits. We also need to ensure that the temperature fluctuation (thermal cycling) of all spacecraft components is minimized. By simulating varying on-orbit scenarios, including varying attitude modes and varying subsystem operation modes, we can also simulate worst-case hot and worst-case cold temperature scenarios. Furthermore, we wish to use the thermal model to simulate testing environments that we will subject the satellite to at various phases throughout the development. Furthermore, we wish to integrate the thermal model into an overall behavioral model of the satellite to assess the interaction of the thermal design with the rest of the satellite.
The inputs to the flux history calculation routine are the satellite’s epoch classical orbital elements, epoch date and Universal Time, the satellite’s attitude control method (Sun-tracking, velocity tracking, tumbling, or quaternions), and the time of flight taken from the simulation clock. The outputs are insolation, Earth infrared, and albedo fluxes for each face with respect to time for an entire orbit.
The flux history calculation model is broken into five modules within MatLab. These modules, along with their inputs and outputs, are discussed here:
COE Update--This module updates the classical orbital elements (COEs) from the epoch time to the current simulation time. Inputs are the epoch COEs, the epoch date and time, and the time of flight, taken from the MatLab simulation clock. This module outputs updated COEs for the satellite and the current Julian date.
Light--This module calculates the sun position vector, the satellite position and velocity vectors, and whether or not the sun currently illuminates the satellite. Inputs are the current COEs and Julian date. Outputs are the satellite position vector (R), satellite velocity vector (V), sun position vector (Rsun), illumination flag (Vis) and satellite/sun Beta angle.
Surface Normals--This module calculates the surface normal vectors of each of the six faces of the satellite. This routine is used if the satellite is sun-tracking, velocity-tracking, or randomly tumbling. There is a switch where the user can choose which tracking mode to use. Alternatively, the surface normal vectors can be calculated using quaternions from an interface with Satellite Tool Kit. There is a switch that allows the user to choose which method of calculating the surface normal vectors they would like to use. Inputs are the satellite position vector (R), satellite velocity vector (V), sun position vector (Rsun), illumination flag (Vis) and satellite/sun Beta angle. Outputs from the module are the surface normal vectors for each face of the satellite, the angle from the +K axis to the satellite R vector (phi), and the angle from the +I axis to the satellite R vector (theta).
Insolation--This module calculates the insolation flux on each of the six faces of the. Its inputs are the surface normal vectors, sun position vector, and illumination flag. It outputs the insolation flux on each face in Wm-2 in both graphical and matrix form.
Earth Effects--This module calculates the Earth Infrared and Albedo flux on each of the six faces of the satellite. This part of the model takes the longest time, as there is a double discrete summation to calculate the Earth IR and Albedo view factors for each face of the satellite. Inputs are the surface normal vectors for each face of the satellite, the satellite position vector (R), the sun position vector (Rsun), the angle from the +K axis to the satellite R vector (phi), and the angle from the +I axis to the satellite R vector (theta). It outputs the Earth infrared and Albedo flux on each face in Wm-2 in both graphical and matrix form.
You can read more about thermal design here.
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