Hybrid rockets can provide a similar degree of flexibility in the thrust profile as a liquid engine but keeping a much simpler architecture. In addition to deal with only a single liquid propellant, hybrids are often designed with some sort of ablative protection system borrowed from the less flexible solid rockets instead of the more complex regenerative-type typically found in liquid rockets. Even if it is not a strict rule as regenerative hybrids or ablative liquids exist, this is the most common approach. In order to properly design the hybrid combustion chamber, it is important to determine the behavior of the thermal protection system during the burn and after the burn, particularly when the multiple fire capability of hybrids has to be exploited. In this paper, the thermal response of thermal protections is investigated trough numerical modelling. A one-dimensional code has been developed that solves the transient heat equation with or without regression of the surface. The code considers the heat transfer normal to the surface from the combustion trough the thermal protection up to the external surface of the hybrid casing, where the radiative heat transfer toward space is applied. The results highlight the importance of the heat soak back after burn, which force the use of thicker thermal protections, higher temperature resistant materials and more careful design if the hybrid is fired multiple times or when the motor case is foreseen to be reusable. However, it is also shown that, when possible, properly using the thrust termination and re-ignition capability of hybrids can help limiting the amount of thermal protections to a level even lower than that of the single burn expendable case. Nevertheless, on the opposite side, other critical situations like an upper stage performing an Hohmann transfer are also highlighted. The methodology and the analyses performed in this paper can also be applied/extended to non-regenerative cooled liquid engines.
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