Method and results of an experimental study of heat losses in a combustion chamber of a solid-propellant propulsion system

A semi-empirical method is proposed for determining one of the key performance characteristics of a solid-propellant rocket motor, i.e., the heat loss coefficient in a combustion chamber. Computational and theoretical determination of the heat loss coefficient typically requires three-dimensional modeling of the gas dynamics of combustion products within the engine flow passage with a charge of complex configuration, as well as the use of several empirical constants, which does not always ensure sufficient accuracy of calculations. The presented method is based on measuring the pressure-time dependence during the adiabatic gas outflow from the chamber after the propellant charge has burned out. The advantage of the method lies in the determination of the heat loss coefficient without thermocouple measurements, which are limited both in the range of measurable temperatures and in their operating time in a high-temperature environment. The relations are obtained for selecting the relaxation time of the free volume of the combustion chamber and the diameter of the nozzle throat that ensure the adiabaticity of the outflow. The method was tested on a model solid-propellant rocket motor with a non-insulated cylindrical steel combustion chamber and an inserted tubular-channel charge of the N-type ballis-tite propellant. The obtained values of the discharge coefficient of a cylindrical nozzle with a sudden contraction (φ = 0.85 ± 0.03) and the heat loss coefficient in the combustion chamber (χ = 0.61 ± 0.02) are consistent with known empirical formulas reported by other authors.

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