Overview
This bachelor thesis developed a COMSOL Multiphysics simulation framework to estimate heat losses of a compact high-temperature lunar ISRU payload. The payload was intended as a small demonstration system for lunar regolith processing, where oxygen and metals can be extracted at high temperatures. Because the payload had to remain within CubeSat-scale dimensions and operate with less than 50 W of heating power, the thermal design was the central feasibility driver.
The work was connected to ESPLORE: Electrochemical Sustainable Production of indigenous Lunar Oxygen and metals from REgolith. The goal was not to design a complete flight-ready payload, but to identify which parameters dominate heat loss and which insulation and geometry choices are most favorable for reaching high oven temperatures with limited input power.
Motivation
Lunar ISRU can reduce the amount of oxygen, metals, and other resources that must be transported from Earth. However, many regolith-processing methods require temperatures above 850 degrees Celsius. Reaching these temperatures in a small lunar lander payload is difficult because heat is lost through insulation, radiation, and interaction with the lunar thermal environment. This thesis studied which parameters dominate these losses and how they affect the achievable oven temperature.
Simulation Framework
The framework combined a simplified lunar thermal environment with a compact payload model. The lunar model included regolith layers, solar flux, radiation to space, subsurface heat flow, surface absorptivity, and solar elevation. The payload model included a high-temperature oven cavity, crucible, insulation layer, and external 1U or 2U CubeSat-scale geometry. Four payload variants were compared: 1U cuboid, 1U cylindrical, 2U cuboid, and 2U cylindrical.
The lunar thermal framework was compared against Apollo 17 lunar surface temperature data, DSNE steady-state estimates, and Lunar Reconnaissance Orbiter temperature data. The validation showed that the framework was suitable for the intended sensitivity and heat-loss study, even though accuracy decreased for high-latitude regions where terrain effects become important.
Payload Design
The study focused on a small 1U/2U CubeSat-scale high-temperature oven payload for lunar lander accommodation. Four payload variants were modeled: 1U and 2U outer sizes, each with either a cuboid or cylindrical oven cavity. The payload model consisted of a high-temperature oven volume inside a crucible, surrounded by high-temperature insulation. After integration, the payload was positioned above the lunar surface and heat exchange between the payload and surface was modeled radiatively. Legs, supports, power electronics, wiring, mechanisms, and the rest of the ISRU process hardware were outside the scope.
Parameter Studies
The study used parameter sweeps, sensitivity studies, and thermal conductivity studies. The parameter sweep compared different payload sizes, geometries, insulation conductivities, surface emissivities, solar absorptivities, solar flux values, and lunar surface conditions. The sensitivity study isolated the most important parameters. The thermal conductivity study calculated the theoretical insulation performance required to maintain oven temperatures between 1000 K and 1500 K using only 10 W to 50 W of input power.
Key Results
The insulation thermal conductivity had the strongest effect on heat loss at high temperatures. At lower temperatures, the payload infrared emissivity became more important. Reducing infrared emissivity from 0.2 to 0.01 improved the oven temperature by about 30%. The required insulation performance was demanding: approximately 0.042 W/(m*K) was required to reach 1000 K with 50 W, while approximately 7.50 x 10^-4 W/(m*K) was required to reach 1500 K with 10 W. The cuboid oven geometry was found to be more favorable than the cylindrical geometry under the assumptions of the model.
High-temperature multi-layer insulation was selected as the most relevant insulation concept for the study. The thesis used literature-based effective thermal conductivity values that depend on the temperature gradient across the insulation, because the insulation cannot be represented accurately by one constant conductivity value across all thermal cases.
- Approximately 0.0024 W/(m*K) effective conductivity at Delta T = 75 K
- Approximately 0.0044 W/(m*K) effective conductivity at Delta T = 400 K
- Approximately 0.011 W/(m*K) effective conductivity at Delta T = 730 K
- Approximately 0.020 W/(m*K) effective conductivity at Delta T = 950 K
Analytical Model
A simplified MATLAB model was derived from the COMSOL results to allow fast early-stage thermal estimates. Depending on the case, the model reproduced the COMSOL oven temperature with an error of roughly 1% to 6%. This made it useful for rapid design screening before running more expensive COMSOL simulations.
My Contribution
This was an individual bachelor thesis. My contribution included:
- Building the lunar thermal environment model in COMSOL.
- Implementing solar flux, radiation to space, subsurface heat flow, lunar regolith properties, and time-dependent solar elevation.
- Validating the lunar framework against Apollo 17, DSNE, and LRO temperature references.
- Developing the 1U and 2U payload models with cuboid and cylindrical oven geometries.
- Implementing high-temperature insulation and emissivity assumptions.
- Integrating the payload and lunar surface frameworks.
- Performing mesh, radius, parameter sweep, sensitivity, and thermal conductivity studies.
- Comparing payload sizes and oven geometries.
- Deriving a simplified MATLAB analytical model from the COMSOL results.
- Writing the final bachelor thesis documentation in LaTeX.
Limitations
The project should be presented as a thermal simulation and feasibility framework, not as a flight-ready payload design. The model did not include a complete heater system, electrical architecture, structural supports, deployment hardware, ISRU chemistry, or detailed lander integration. The lunar environment was simplified, and the analytical model is only valid within the parameter range investigated in the thesis.
- No detailed ISRU chemistry was modeled.
- No real heater design, power electronics, wiring, or control system was included.
- Payload legs and structural supports were neglected.
- Heat transfer between payload and lunar surface was modeled as radiation only.
- The payload geometry was simplified to idealized 1U and 2U shapes.
- The insulation behavior was based on literature-derived effective conductivity values.
- Terrain effects at high lunar latitudes were not fully resolved.
- The analytical MATLAB model is only valid within the parameter space used to derive it.
Tools and Methods
- COMSOL Multiphysics
- MATLAB
- LaTeX
- Thermal simulation
- Radiative heat transfer
- Parameter sweeps
- Sensitivity analysis
- Mesh study
- Model validation
- Analytical approximation