Transient heat transfer of liquid methane under forced convection in a 1.8 mm × 1.8 mm asymmetrically heated square channel was investigated. This study is aimed at understanding the heat transfer behavior of cryogenic propellant in cooling channels of a regeneratively cooled rocket engine at the start-up condition. To simulate high heat load conditions representative of regeneratively cooled rocket engines, a high heat flux test facility with cryogenic liquid handing capabilities was developed at the Center for Space Exploration Technology Research. The time history of inlet and outlet fluid temperatures and test section channel wall temperatures were measured at high heat flux conditions (from 1.19 to 3.80 MW/m²) and a Reynolds number (Re) range of 1.88 × 10⁵ to 3.45 × 10⁵. The measured wall temperature data point toward possible film boiling within the test section during certain tests, particularly with higher heat fluxes and lower Reynolds number conditions that resulted in higher wall temperatures. The transient average Nusselt numbers (Nu_L) of the channel obtained from the experimental measurements are lower than those calculated from the Sieder–Tate correlation (Nu_O); however, the ratio (Nu_L/Nu_O) increases with the increase in Reynolds number. The ratio is around 0.25 at the lower end of Re and then increases to 0.7 at the maximum Re studied in the present investigation.

The flame extinction limits of syngas $(H2‐CO)$ flames were measured using a twin-flame counterflow burner. Plots of extinction limits ($%f$: volumetric percent of fuel in air) versus global stretch rates were generated at different fuel blend compositions and were extrapolated to determine the flame extinction limit corresponding to an experimentally unattainable zero-stretch condition. The zero-stretch extinction limit of $H2‐CO$ mixtures decreases with the increase in $H2$ concentration in the mixture. The average difference between the measured flame extinction limit and the Le Chatelier’s calculation is around 7% of the mean value. The measured OH chemiluminescence data indicates that regardless of blend composition the OH radical concentration reduces to a critical value prior to the flame extinction. The measured laminar flame velocity close to the extinction indicates that regardless of fuel composition, the premixed flame of hydrogen fuel blends extinguishes when the mixture laminar flame velocity falls below a critical value.

The need to refuel future missions to Mars and the Moon via in situ resource utilization (ISRU) requires the development of LOX/LCH4 engines, which are complex and expensive to develop and improve. This paper discusses how the use of digital engineering—specifically physics-based modeling (PBM)—can aid in developing, testing, and validating a LOX/LCH4 engine. The model, which focuses on propulsion performance and heat transfer through the engine walls, was created using Siemens’ STAR-CCM+ CFD tool. Key features of the model include Eulerian multiphase physics (EMP), complex chemistry (CC) using the eddy dissipation concept (EDC), and segregated solid energy (SSE) for heat transfer. A comparison between the complete GRI 3.0 and Lu’s reduced combustion mechanisms was performed, with Lu’s mechanism being chosen for its cost-effectiveness and similar output to the GRI mechanism. The model’s geometry represents 1/8th of the engine’s volume, with a symmetric rotational boundary. The performance of this engine was investigated using NASA’s chemical equilibrium analysis (CEA) and STAR-CCM+ simulations, focusing on thrust levels of 125 lbf and 500 lbf. Discrepancies between theoretical predictions and simulations ranged from 1.4% to 28.5%, largely due to differences in modeling assumptions. While NASA CEA has a zero-dimensional, steady-state approach based on idealized conditions, STAR-CCM+ accounts for real-world factors such as multiphase flow, turbulence, and heat loss. For the 125 lbf case, a 9.2% deviation in combustion chamber temperature and a 15.0% difference in thrust were noted, with simulations yielding 113.48 lbf compared to the CEA’s 133.52 lbf. In the 500 lbf case, thrust reached 488 lbf, showing a 2.4% deviation from the design target and an 8.6% increase over CEA predictions. Temperature and pressure deviations were also observed, with the highest engine wall temperature at the nozzle throat. Monte Carlo simulations revealed that substituting LNG for LCH4 affects combustion dynamics. The findings emphasize the need for advanced modeling approaches to enhance the prediction accuracy of rocket engine performance, aiding in the development of digital twins for the CROME.

Flame synthesis of carbon nanostructures including nanotubes on galvanized steel was investigated utilizing laminar diffusion flames of different types of fuel. Methane (CH₄), propane (C₃H₈) and acetylene (C₂H₂) were used as fuels. Distinctive carbon nanostructures were produced depending on fuel types and fuel flow rates. The qualitative and quantitative analysis of many transmission electron microscope (TEM) and scanning electron microscope (SEM) images were performed. Methane produced thin multi wall carbon nanotubes as well as nanorods and nanofibers within the fuel flow rate range of 7.18E‐07 m³/s to 9.57E‐07 m³/s. Propane yielded nanotubes only at the fuel flow rate of 4.20E‐07 m³/s. The nanotubes synthesized by acetylene flames were of different types that included helically coiled and twisted nanotubes.