Reconstructed Ballistic Data Versus Wax Regression-Rate Intrusive Measurement in a Hybrid Rocket

Abstract

Burning tests of a laboratory-scale hybrid rocket engine were carried out with gaseous oxygen and a microcrystallinewax based fuel to look into the feasibility of using an intrusive resistor-based sensor for measuring the fuel regression rate.This initial screening was driven by the need for real-time control of the oxidizer-to-fuel ratio in altering-intensity swirlingflow-type hybrid rocket engines aiming at performance optimization. To build up a framework for comparison with the measured data, a traditional ballistic reconstruction technique was critically revised; with the measured aft-chamber pressure and oxygen mass flow rate time histories, the fuel regression-rate and port diameter were reconstructed over the firing, by estimating the combustion efficiency with the constraint that calculated and measured fuel mass consumed are equal. This technique invariably suffers from the issue of presenting multiple solutions for the fuel mass flow rate in the proximity of the optimum mixture ratio, for which a novel variable-efficiency approach is proposed. Reconstructed data show that regression rate is nearly constant in each firing, yielding dependence upon the port diameter other than the mass flux.Resistor-sensor raw data displayed large deviation from the ballistic results for the slower burning rate of the sensor support.A detailed analysis is presented.

Bridging the Technology Gap: Strategies for Hybrid Rocket Engines

“…To obtain the temporal evolution of the regression rate, other methods have been developed. These include, for example, the use of X-ray tomography during the burn [111,112], the application of ultrasonic transducers to measure the web thickness during the burn [114], placing resistors in the fuel grain [87,115], and the so-called ballistic reconstruction technique, which dates back to 1966 [116].…”

“…Ballistic reconstruction is based on the mass balance inside the combustion chamber; the total propellant mass (oxidizer and fuel) equals the mass expelled through the nozzle, which is calculated using the characteristic velocity c * , nozzle throat area A t , and chamber pressure P c ) [115]:…”