Theory of acoustic instability in solid-propellant rocket combustion

Hart, R. W.; McClure, F. T.

doi:10.1016/s0082-0784(65)80246-6

Cited by 67 publications

(22 citation statements)

References 27 publications

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“…7 For an end burner (the mass addition situation most closely corresponding to the present experiment), …”

Section: Comparison Of Results With Theorymentioning

confidence: 68%

“…This loss is associated with the presence of a mean flow across the cavity boundaries. As noted by Hart and McClure,7 this convective loss at the nozzle is normally balanced for axial mode oscillations by a convective gain at the gas inflow surface when the in and outflow regions of the burner are at the ends. Neither of the present experiments appear to have the capability to measure these convective effects.…”

Section: Discussion Of the Datamentioning

confidence: 95%

See 1 more Smart Citation

Acoustic attenuation experiments on subscale, cold-flow rocket motors.

Buffum¹,

Dehority²,

Price³

et al. 1967

AIAA Journal

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Acoustic dissipation studies were undertaken using a subscale cold-flow rocket motor model to determine the effect of purely geometric variables on the acoustic performance leading to axial mode combustion instability. Cold air served as the working fluid to simulate the internal flow and to obtain a critical nozzle throat. By using interchangeable parts, it was possible to vary the nozzle throat diameter and to simulate the geometry of a cylindrical-bore grain at various motor burn times. The model was acoustically driven by three separate methods: pulse decay, steady-state decay, and steady-state resonance. The methods are compared, and the applicability of each is discussed. Test results show that the transmission of acoustic energy through the nozzle was the largest source of loss for the axial mode and that the losses increase linearly with / (ratio of nozzle throat area to grain port area). Acoustic bulk and wall losses were negligible compared with the nozzle losses. The results also showed that the nozzle losses are independent of the way Jis changed and thus can be scaled. The test equipment has been established as a semi-automated facility to evaluate models with more complex geometries, to test chamber damping devices, and to gather design and research data. Nomenclature a* = critical nozzle velocity, m/sec c = velocity of sound at stagnation conditions, m/sec c p = specific heat at constant pressure, joules/kg °C D = diameter of the grainport channel, m /o = first mode resonant frequency, cps f lt f 2 = half-power frequencies, cps J = ratio of nozzle throat area to grain port area L = length of motor chamber, m M = Mach number p = acoustic or a. c. pressure (p-to-p), newton/m 2 po = initial acoustic or a.c. pressure (p-to-p), newton/m 2 p r = Prandtl number Q = quality factor t = time, sec V = d.c. or mean chamber gas velocity, m/sec a = temporal damping constant, sec" 1 7 = ratio of specific heats K = thermal conductivity of gas, joule/m sec°C I JL = viscosity, newton sec/m 2 TT '= 3.1416 p = d.c. or mean gas density, kg/m 3 PO = stagnation gas density at nozzle, kg/m 3 p* = d.c. gas density at sonic nozzle throat, kg/m 3 co = angular frequency, r ad/sec UQ = angular resonant frequency, rad/sec

show abstract

“…7 For an end burner (the mass addition situation most closely corresponding to the present experiment), …”

Section: Comparison Of Results With Theorymentioning

confidence: 68%

Section: Discussion Of the Datamentioning

confidence: 95%

Acoustic attenuation experiments on subscale, cold-flow rocket motors.

Buffum¹,

Dehority²,

Price³

et al. 1967

AIAA Journal

View full text Add to dashboard Cite

show abstract

“…Nearly a decade later, the work of McClure and coworkers would prove instrumental in the understanding of rocket motor stability, especially in the development of the energy balance framework. However, principal efforts in this direction have focused on the thin region near the injecting surface (Hart & McClure, 1965;Hart et al, 1960;Hart & McClure, 1959;McClure et al, 1960). In fact, McClure et al (1963) may have been the first to employ a mean flow approximation in their analysis of the aeroacoustic field in SRMs.…”

Section: Relevance To Propulsion Systemsmentioning

confidence: 99%

Internal Flows Driven by Wall-Normal Injection

Majdalani¹,

Saad²

2012

Advanced Fluid Dynamics

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“…1,2 Owing to this fact, analytical methodologies put forward to describe flow oscillations lean heavily on the assumption of small acoustic disturbances. [3][4][5][6][7][8][9][10][11][12][13][14][15][16][17] Contrary to this assumption, however, a vast body of experimental evidence conveys a dissimilar picture, specifically, one involving large amplitude oscillations with steep gradients in flow variables. For example, in the extensive experimental studies of Clayton, Sotter, and co-workers, [18][19][20][21] a heavily instrumented, laboratory scale, 20 000 lbf thrust engine was used to investigate high amplitude tangential oscillations.…”

Section: Introductionmentioning

confidence: 99%

Acoustic streaming in simplified liquid rocket engines with transverse mode oscillations

2010

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This study considers a simplified model of a liquid rocket engine in which uniform injection is imposed at the faceplate. The corresponding cylindrical chamber has a small length-to-diameter ratio with respect to solid and hybrid rockets. Given their low chamber aspect ratios, liquid thrust engines are known to experience severe tangential and radial oscillation modes more often than longitudinal ones. In order to model this behavior, tangential and radial waves are superimposed onto a basic mean-flow model that consists of a steady, uniform axial velocity throughout the chamber. Using perturbation tools, both potential and viscous flow equations are then linearized in the pressure wave amplitude and solved to the second order. The effects of the headwall Mach number are leveraged as well. While the potential flow analysis does not predict any acoustic streaming effects, the viscous solution carried out to the second order gives rise to steady secondary flow patterns near the headwall. These axisymmetric, steady contributions to the tangential and radial traveling waves are induced by the convective flow motion through interactions with inertial and viscous forces. We find that suppressing either the convective terms or viscosity at the headwall leads to spurious solutions that are free from streaming. In our problem, streaming is initiated at the headwall, within the boundary layer, and then extends throughout the chamber. We find that nonlinear streaming effects of tangential and radial waves act to alter the outer solution inside a cylinder with headwall injection. As a result of streaming, the radial wave velocities are intensified in one-half of the domain and reduced in the opposite half at any instant of time. Similarly, the tangential waves are either enhanced or weakened in two opposing sectors that are at 90°angle to the radial velocity counterparts. The second-order viscous solution that we obtain clearly displays both an oscillating and a steady flow component. The steady part can be an important contributor to wave steepening, a mechanism that is often observed during the onset of acoustic instability.

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Theory of acoustic instability in solid-propellant rocket combustion

Cited by 67 publications

References 27 publications

Acoustic attenuation experiments on subscale, cold-flow rocket motors.

Acoustic attenuation experiments on subscale, cold-flow rocket motors.

Internal Flows Driven by Wall-Normal Injection

Acoustic streaming in simplified liquid rocket engines with transverse mode oscillations

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