📋 Case Study

Thermal Design of Satellite Payload Radiator for Lunar Orbit Mission

Extreme radiative environment: solar flux up to 1360 W/m², albedo up to 0.12, IR emission from hot regolith (~390 K)

🏗️ Project Overview

NASA CLPS payload requiring stable 20±2°C operation during 14-day lunar day/night cycle

🎯 Challenge

Extreme radiative environment: solar flux up to 1360 W/m², albedo up to 0.12, IR emission from hot regolith (~390 K)

🔧 Design Approach

Deployable high-emissivity (ε = 0.92) radiator with variable-emittance coatings (VEC) and thermal louvers controlled by bi-metallic actuators

📐 Design Diagram

AI-generated project design illustration

📐 Key Calculations

Net Radiative Balance

Q_net = εσ(T_rad⁴ − T_space⁴) − αG_solar − αG_albedo

Result: −128 W/m² (day), +18 W/m² (night)

Drives radiator area sizing

Louvers Open Fraction

f_open = (T_rad − T_set)/ΔT_range

Result: 0.35 (day), 1.0 (night)

Passive thermal regulation without power

📊 Results

Payload temperature stability achieved across full orbit: 19.8–20.3°C; no heater use required during night; mass penalty <1.2 kg

💡 Lessons Learned

• •Second-surface optics modeling essential for albedo reflection paths
• •VEC hysteresis must exceed orbital thermal lag to prevent oscillation

✅ Key Takeaways

• 1Second-surface optics modeling essential for albedo reflection paths
• 2VEC hysteresis must exceed orbital thermal lag to prevent oscillation