Search

An investigation of the instability mechanism present in a laboratory rocket combustor is performed using computational fluid dynamics (CFD) simulations. Three cases are considered which show different levels of instability experimentally. Computations reveal three main aspects to the instability mechanism, the timing of the pressure pulses, increased mixing due to the baroclinic torque, and the presence of unsteady tribrachial flame. The stable configuration shows that fuel is able to flow into the combustor continuously allowing continuous heat release. The unstable configuration shows that a disruption in the fuel flow into the combustor allows the heat release to move downstream and new fuel to accumulate in the combustor without immediately burning. Once the large amounts of fuel in the combustor burn there is rapid rise in pressure which coincides with the timing of the acoustic wave in the combustor. The two unstable cases show different levels of instability and different reignition mechanism.

Combustion instability in rocket chambers is strongly influenced by acoustic interactions at the boundaries of the configuration. Many CFD simulations employ approximate boundary conditions in order to simplify the geometry but the impact that they have on the solution is not well understood. The present study focuses on the use of detailed (exact) boundary representations and an approximate boundary condition in a given longitudinal mode test chamber. The actual inlet boundary of the injector is comprised of a series of choked slots while the approximate boundary condition is a uniform constant mass flow. Both two and three-dimensional simulations are carried out. Differences in the flowfield are evident in the combustion region away from the inlet, including the size of the recirculation region and location of the peak heat release. The amplitudes of the acoustic modes are well predicted for the first two modes especially in three-dimensional simulation, while higher modes are poorly predicted. These results suggest that such boundary condition approximations must be judiciously used and having access to more detailed treatments is important to verify accuracy.

Concurrent experiments and computations are used to analyze combustion instabilities in a transverse mode combustion chamber. The experiments employ a shear-coaxial injector element, positioned within a rectangular chamber. The reacting flow portion of the study element is optically accessible and the chamber is extensively instrumented with high-frequency pressure transducers. High amplitude transverse acoustics modes are driven by unstable injector elements located near the chamber end-walls. Different levels of instability are obtained by varying the operation of these driving elements. High-fidelity computational fluid dynamics simulations are used to model this set-up, although only the study element is fully represented and the transverse acoustics modes are generated by vibrating the side walls at the appropriate frequencies. The computational results are compared quantitatively with the high frequency pressure measurements, and qualitatively by using the CH* chemiluminescence signal from the experiment. The combustion response of the first, second and third transverse modes obtained using a dynamic modal decomposition procedure show excellent agreement between the experiments and simulations. The overall approach shows significant promise for screening the combustion response of candidate injector configurations for rocket applications.