Orbit / Environment Preset
External Thermal Environment
Solar constant Gs [W/m²]1361
Earth IR flux q_IR [W/m²]230
Albedo coefficient a0.30
Spacecraft Surface Properties
Solar absorptance αs0.15
IR emissivity ε0.85
Projected area As [m²]1.00 m²
Total surface area A [m²]6.00 m²
Internal power Q_int [W]100 W
Eclipse fraction β0.35
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Sunlit T_sun [°C]
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Eclipse T_ecl [°C]
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Absorbed Q_in [W]
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Radiator area [m²]
Heat Budget Breakdown (Sunlit)
▲ Orbital Temperature Variation (Sunlit/Eclipse Cycle)
▲ Required Radiator Area vs Design Temperature (for varying Q_int)
Theory
Steady-state heat balance: $Q_{in} = Q_{out}$
$$Q_{in} = \alpha_s A_s G_s + \alpha_s A_s a G_s F_{alb} + \varepsilon_{IR} A_{IR} q_{IR} + Q_{int}$$ $$Q_{out} = \varepsilon \cdot A_{total} \cdot \sigma \cdot T^4$$Equilibrium temperature:
$$T_{eq} = \left(\frac{Q_{in}}{\varepsilon \cdot A_{total} \cdot \sigma}\right)^{1/4}$$Radiator sizing: $A_{rad} = Q_{reject}/(\varepsilon \sigma T_{rad}^4 - q_{abs})$
CAE Note: Dedicated spacecraft thermal tools include Thermal Desktop/SINDA and ESATAN-TMS. ANSYS Mechanical can model orbital thermal loading via radiation boundary conditions and view factors computed by Monte Carlo ray tracing. MLI effective emittance: ε_eff = 1/(N/ε_sheet + 1). OSR surface finish: αs ≈ 0.05, ε ≈ 0.80.