No abstract

Down-scaled solid propellant motors are a valuable tool in the study of thrust oscillations and the underlying, vortex-shedding-induced, pressure instabilities. These fluctuations, observed in large segmented solid rocket motors such as the Ariane 5 P230, impose a serious constraint on both structure and payload. This paper contains a survey of the numerous configurations tested at ONERA over the last 20 years. Presented are the phenomena searched to reproduce and the successes and failures of the different approaches tried. The results of over 130 experiments have contributed to numerous studies aimed at understanding the complicated physics behind this thorny problem, in order to pave the way to pressure instability reduction measures. Slowly but surely our understanding of what makes large segmented solid boosters exhibit this type of instabilities will lead to realistic modifications that will allow for a reduction of pressure oscillations. A ''quieter'' launcher will be an important advantage in an ever more competitive market, giving a easier ride to payload and designers alike.

Numerical simulation is performed and compared with a firing test of an H 2 O 2 ∕polyethylene hybrid rocket engine equipped with a catalyzer. The simulation is performed assuming an axisymmetric flow, unsteady Reynolds-averaged Navier-Stokes equations are solved based on a single-phase flow. Stable pressure oscillations are observed during the firing test and could be explained by the simulation with the formation of large-scale vortices at the end of the fuel grain and in the postcombustion chamber. These vortices transport unburnt fuel to the nozzle and play a major role in the mixing process in the postchamber. Two semi-empirical approaches are applied and provide a good order of magnitude for the frequency and frequency shift during the firing test. Another simulation has been performed with an added oxidizer postchamber injection. It revealed that small-scale vortices along the fuel grain in the baseline case are probably generated by the oscillatory motion of the flow due to pressure fluctuations in the engine. Nomenclature C μ = turbulence model constant c = speed of sound, m · s −1 c th = theoretical characteristic velocity, m · s −1 D = fuel port diameter, m D t = nozzle throat diameter, m f = frequency, Hz k = turbulence kinetic energy, m 2 · s −2 k Ross = ratio of vortex transport velocity to freestream velocity L = total combustion chamber length, m L f = fuel grain length, m L pre ch = length of the prechamber, m L Ross = distance between shear layer initiation and impingement point, m l = exhaust nozzle length, m M = Mach number m = stage number (number of coherent structures) _ m = mass flow rate, kg · s −1 O∕F= oxidizer-to-fuel mixture ratio p = pressure, Pa Q = Q criterion, rad 2 · s −2 R = specific gas constant, J · kg −1 · K −1 R f = fuel port radius, m r = distance from wall, m S = symmetric part of velocity gradient tensor S D = Strouhal number based on end-port diameter T = temperature, K T Ross = period, s U = freestream velocity, m · s −1 U vortex = vortex transport velocity, m · s −1 V = combustion port total volume, m 3 V x = axial velocity, m · s −1 x = axial position, m Y = mass fraction y = radial position, where y equal to zero is the axis, m α = empirical constant Δt = correction factor, s η c = combustion efficiency κ = von Kármán constant ρ = density, kg · m −3 τ bl = boundary-layer delay time, s ψ = function of isentropic coefficient Ω = skew-symmetric part of velocity gradient tensor ω = turbulence specific dissipation rate, s −1 Subscripts 1L = first longitudinal acoustic mode end = end of firing test H = Helmholtz mode hy = hybrid intrinsic instability ini = beginning of firing test ox = oxidizer

No abstract