Speaker
Description
The rapid growth of small-satellite and CubeSat missions has increased the demand for compact propulsion systems capable of providing safe, flexible, and reliable in-space maneuvering capability. Mission scenarios involving orbit insertion, formation flying, constellation reconfiguration, collision avoidance and de-orbiting require propulsion technologies that can be integrated within highly constrained platforms while preserving operational robustness. In this context, hybrid rocket engines represent a promising solution for small spacecraft because they combine intrinsic safety, throttling potential, restart capability, and relatively simple system architecture. However, their application to CubeSat-class propulsion requires accurate predictive models of internal ballistics and fuel regression, since the regression rate directly controls the mixture ratio, chamber pressure, thrust level, and overall engine performance. Among the mechanisms governing fuel regression, the wall heat balance plays a central role and results from the coupled contribution of convective and radiative heat transfer. While thermal radiation has been investigated in medium- and large-scale hybrid rocket engines, its relevance in small-scale configurations remains less clearly established, especially for hydrogen-peroxide-based systems intended for CubeSat propulsion. In this work, gas-phase thermal radiation is investigated in a 10 N-class hybrid rocket engine designed for small-satellite applications. The analysis is carried out through a coupled Computational Fluid Dynamics-gas/surface interaction framework, which allows the reacting flow field, wall heat transfer, and solid fuel regression to be predicted in a consistent manner. The numerical model solves the steady-state RANS equations for compressible, multicomponent reacting flows, using SST turbulence closure and a non-premixed combustion model based on the mixture fraction approach, in which a PDF formulation is coupled with chemical equilibrium calculations to account for turbulence-chemistry interaction. The fuel surface is treated through an iterative gas/surface interaction procedure, in which the local regression rate, wall temperature, and fuel mass injection are obtained from coupled surface mass and energy balances with Arrhenius-type HDPE pyrolysis. Radiative heat transfer in the participating medium is modeled through a WSGGM/DTRM approach, in which the spectral behavior of the gas is represented by a weighted sum of gray gases and the radiative transfer equation is solved by tracing discrete rays through the computational domain. The gas phase is treated as a non-scattering participating medium, and the radiation calculation accounts for the main radiatively active combustion products, namely H2O and CO2. After validation against experimental firing tests of a small-scale hydrogen-peroxide/HDPE hybrid rocket engine, the model is used to examine how chamber pressure, propellant combination, and geometric scale affect the radiative contribution to the wall heat balance. To further assess the robustness of the scale comparison, a dedicated geometric-scaling analysis is also performed. In this case, the pre- and post-chamber lengths of the small-scale GOX/HDPE engine are reduced to match the normalized geometric parameters of a reference large-scale configuration, while maintaining comparable oxidizer mass flux and port-to-injector diameter ratio.
The experimental reference consists of a pressure-regulated firing campaign designed to reproduce the operating envelope of a low-thrust hybrid rocket engine for small-satellite applications. The engine integrates a catalytic decomposition chamber, where hydrogen peroxide is converted into a high-temperature H2O/O2 mixture, with an axial gas-injection system feeding a combustion chamber composed of a pre-chamber, a cylindrical single-port HDPE grain, and a post-chamber. Since the target thrust level is of the order of a few newtons, the oxidizer mass flow rates are necessarily small, whereas the oxidizer mass flux is governed by the compact port geometry and falls in the range of about 19-33 kg/m2s. Under these experimentally validated operating conditions the results show that, for HTP-based operation in the 10 N thrust class, thermal radiation provides only a limited contribution to the fuel surface energy balance. Across the validated firing tests, the radiative fraction remains below 5% of the total wall heat flux, indicating that convective heat transfer dominates under practical CubeSat-scale operating conditions. Increasing the chamber pressure leads to a clear increase in the radiative wall heat flux, mainly because the higher gas density and larger partial pressures of the radiatively participating species increase the absorption coefficient and, consequently, the optical thickness of the medium. However, this enhancement does not translate into a comparable increase in the relative contribution of radiation to the overall wall heat balance. Within the investigated operating envelope, convective heat transfer remains the dominant mechanism, and radiation continues to represent only a secondary contribution to the total heat flux. A different behavior is obtained when the same small-scale engine is operated with gaseous oxygen. In this case, the higher combustion temperature leads to a marked increase in radiative heating, with the radiative contribution rising from 4.5% in the H2O2/HDPE case to 17.5% in the GOX/HDPE configuration. The comparison with a dynamically comparable 1 kN-class GOX/HDPE engine further shows that the increase in characteristic dimension leads to longer radiative path lengths and higher optical thickness, raising the radiative contribution up to 28.1%. The additional geometrically normalized small-scale analysis confirms that pre- and post-chamber geometry affects the detailed thermal field and wall heat balance: when the small-scale geometry is modified to match the normalized large-scale configuration, the radiative contribution decreases from 17.5% to 8.9%. Nevertheless, the overall physical trend remains unchanged.
Overall, the study shows that gas-phase radiation can be reasonably neglected as a first-order contribution in hydrogen-peroxide-based CubeSat-scale hybrid rocket engines operating under practical conditions, whereas it becomes increasingly relevant for higher-temperature propellant combinations and larger characteristic dimensions. The geometric-scaling analysis further indicates that differences in pre- and post-chamber layout influence the quantitative value of the radiative heat-flux fraction but do not modify the main conclusion regarding the governing role of propellant thermochemistry and optical thickness. These findings provide a validated numerical reference for the design and modeling of small-scale hybrid rocket engines for small-satellite missions and help define the conditions under which radiative heat transfer must be explicitly included in predictive internal-ballistics tools.
Summary
Numerical investigation of radiative heat transfer in a CubeSat-scale hybrid rocket engine. Results show that radiation is negligible for HTP/HDPE operation but becomes increasingly important for GOX configurations and larger scales.