ACOUSTIC DAMPING MECHANISMS OF HALF-WAVE RESONATORS IN A ROCKET ENGINE ENVIRONMENT
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- NGN 3000
ABSTRACT
Combustion instabilities in rocket engines are caused by coupling between the combustion processes and pressure oscillations in a combustor, and they are characterized by the frequencies and shapes of the acoustic modes of the combustion chamber. Acoustic resonators are commonly installed in combustors to provide passive acoustic damping and prevent combustion instabilities. Previously, it has been proposed that the propellant injectors in a combustor can be tuned to act as half-wave resonators and provide acoustic damping. This requires a thorough understanding of the acoustic damping mechanisms of injectors, and producing this understanding forms the basis of the work described in this dissertation. In this work, the acoustic damping of propellant injectors is measured experimentally. A new experimental facility is developed and employed to measure the sound power reflection, transmission, and dissipation under the conditions of mean flow, high temperature, high acoustic amplitude, and higher-order modes. The effects of common design parameters—namely, the typical features of an injector, the ratio between the crosssectional area of the injectors and the combustion chamber, the number of injectors, and the relative position of the injectors within the cross-section of the combustion chamber—on the acoustic absorption coefficient are investigated experimentally using the new facility. Measurements of the fraction of sound power dissipated by the injectors and the velocity flow fields at the ends of the injectors are used to elucidate the physical mechanisms responsible for the acoustic damping. Attempts are made to quantify the separate contributions of viscothermal dissipation and the conversion of acoustic energy into vorticity. Analytical and numerical models incorporating some of these dissipation mechanisms are developed to predict the absorption coefficient of the injectors.
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