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Nonlinear combustion stability problems in liquid propellant rocket engines, describing unsteady combustion response by Crocco time lag hypothesis
Application of fluidic devices to reduce generation of noxious exhausts from internal combustion engines is discussed. Operation of fluidic system to provide bypass of fuel air mixture into heated loop to provide more complete combustion is explanined. Advantage lies in no moving parts required for fluidic by-pass action.
Device has been designed and developed for igniting nonrecoverable combustible gases and sustaining combustion of gases evolving from various gas vent stacks. Igniter is superior to existing systems because of simplicity of operation, low cost fabrication, installation, operational and maintainability features, and excellent reliability in all phases of required operations.
Polymer degradation to small fragments is treated as a rate controlling step in hybrid rocket combustion. Both numerical and approximate analytical solutions of the complete energy and polymer chain bond conservation equations for the condensed phase are obtained. Comparison with inert atmosphere data is very good. It is found that the intersect of curves of pyrolysis rate versus interface temperature for hybrid combustors, with the thermal degradation theory, falls at a pyrolysis rate very close to that for which a pressure dependence begins to be observable. Since simple thermal degradation cannot give sufficient depolymerization at higher pyrolysis rates, it is suggested that oxidative catalysis of the process occurs at the surface, giving a first order dependence on reactive species concentration at the wall. Estimates of the ratio of this activation energy and interface temperature are in agreement with best fit procedures for hybrid combustion data. Requisite active species concentrations and flux are shown to be compatible with turbulent transport. Pressure dependence of hybrid rocket fuel regression rate is thus shown to be describable in a consistent manner in terms of reactive species catalysis of polymer degradation.
Data from prototype GO2-GH2 injection elements were obtained and analyzed. The bulk of the testing was conducted with nonreacting propellants, N2 to simulate the O2 and H2. A limited number of tests were conducted in combusting environment, with the purpose of this testing being to evaluate the effects of combustion on cold flow mixture ratio and mass profiles. The cold flow testing was completed in 223 tests. Analytical model correlations of these data are under way. The hot flow testing was completed in 36 tests. These data have been reduced and are documented in the report.
A preliminary analysis has been made of a supersonic-combustion rocket engine concept using hydrogen and oxygen propellants. The ejector action of a separate small rocket motor is employed to pump the propellants to high stagnation pressures and supersonic velocities. Therefore complicated heavy turbopumps are eliminated and cooling problems of a sonic throat are reduced. The results of the study show that vacuum specific impulse levels as high as a conventional rocket having the same chamber pressure as the drive motor are possible. The supersonic-combustion rocket would be an attractive alternate for a high-altitude low-thrust conventional rocket operating with a pressure feed propellant system. It would also be a convenient technique for obtaining extremely high thrusts without the need for developing corresponding large turbopumps.
An analytical technique was developed to solve nonlinear longitudinal combustion instability problems. The analysis yields the transient and limit cycle behavior of unstable motors and the threshold amplitude required to trigger a linearly stable motor into unstable operation. The limit cycle waveforms were found to exhibit shock wave characteristics for most unstable engine operating conditions. A method of correlating the analytical solutions with experimental data was developed. Calculated results indicate that a second-order solution adequately describes the behavior of combustion instability oscillations over a broad range of engine operating conditions, but that higher order effects must be accounted for in order to investigate engine triggering.
Experimental data were interpreted using two supersonic combustion computer programs. The P1 program is based on a conventional boundary layer treatment of the mixing of concentric gas streams and complete combustion chemistry. The H1 program is based on a modified boundary layer approach which accounts for radial pressure gradients in the flow and also incorporates a finite rate chemistry calculation. The objective of the investigation was to compare the experimental data with theoretical predictions of the two programs with special emphasis on the prediction of radial pressure gradients by the H1 program. A test of the H1 program was also desired through comparison with the experimental data and with the P1 program.
A two-year study of the combustion efficiency of single beryllium droplets burning in a variety of oxidizers (primarily mixtures of oxygen/argon and oxygen/nitrogen) is summarized. An advanced laser heating technique was used to acquire systematic quantitative data on the burning of single beryllium droplets at atmospheric pressure. The research confirmed the sensitivity of beryllium droplet combustion to the chemistry of environmental species and provides experimental documentation for the nitrogen-induced droplet fragmentation of burning beryllium droplets.
The ignition of a combustible gas mixture by a hot cylinder under the effect of a gravity field for steady state conditions is examined. For this purpose a horizontal cylinder is considered with gravity as a parameter together with a finite chemical reacting flow generated by free convection with the additional effect of diffusion. Both mass transfer and zero mass transfer cases are considered. By defining an ignition criterion the surface temperature and species are obtained from the analysis as a function of the gravity field. It is supposed that at the point of ignition the heat evolved in the gas is sufficiently high to attain a sustained combustion without any energy from the hot cylinder.
The solution of problems of combustion instability for more effective communication between the various workers in this field is considered. The extent of combustion instability problems in liquid propellant rocket engines and recommendations for their solution are discussed. The most significant developments, both theoretical and experimental, are presented, with emphasis on fundamental principles and relationships between alternative approaches.
Analytical studies of the transient and steady-state combustion processes in a hybrid rocket system are discussed. The particular system chosen consists of a gaseous oxidizer flowing within a tube of solid fuel, resulting in a heterogeneous combustion. Finite rate chemical kinetics with appropriate reaction mechanisms were incorporated in the model. A temperature dependent Arrhenius type fuel surface regression rate equation was chosen for the current study. The governing mathematical equations employed for the reacting gas phase and for the solid phase are the general, two-dimensional, time-dependent conservation equations in a cylindrical coordinate system. Keeping the simplifying assumptions to a minimum, these basic equations were programmed for numerical computation, using two implicit finite-difference schemes, the Lax-Wendroff scheme for the gas phase, and, the Crank-Nicolson scheme for the solid phase.
Computations are presented describing the mixing and combustion of swirling jets in a coaxial stream. It is demonstrated that the boundary layer equations represent the flow reasonably well until reversed flow is imminent. For the range of parameters investigated indications are that the edge velocity has little effect on the behavior of the flow. Furthermore, confining the flow with a constant pressure wall, or impressing a favorable pressure gradient on the coaxial flow, acts to reduce the severity of the centerline adverse pressure gradient created by the swirl decay. A simple scalar eddy viscosity model, including a potential core formulation, is shown to described the behavior of weak swirling flow in the far region but is only in fair agreement with observations in the near region. The effects of swirl on a burning hydrocarbon jet exhausting into a cold coaxial stream are shown to be intensified by the reduction of the density due to combustion. The enhanced mixing properties of high swirl flow produce rapid diffusion of the burning gases into the cold edge flow causing early cessation of the NO producing reactions. Computations show that doubling the initial jet swirl could reduce the NO production by 25 percent.
A chemical analysis of the emission and absorption spectra in the combustion chamber of a nitrogen tetroxide/aerozine-50 rocket engine was conducted. Measurements were made under conditions of preignition, ignition, and post combustion operating periods. The cause of severe ignition overpressures sporadically observed during the vacuum startup of the Apollo reaction control system engine was investigated. The extent to which residual propellants or condensed intermediate reaction products remain after the engine has been operated in a pulse mode duty cycle was determined.
A program was conducted to develop and experimentally evaluate an improved version of a modified machine gun for use as a device for rating the relative combustion stability of various rocket combustors. Following the results of a previous study involving a caliber .30 machine gun, a caliber .50 machine gun was modified in order to extend the charge-size range of the device. Nitrocellulose charge sizes ranging from 1.004 to 9.720 grams were fired at rates up to four shots per second. Shock pressures up to 25,512 kN/sq m were measured near the end of a shortened gun barrel. A minimal resistance type of check valve permitted the gun to fire into pressurized regions; back pressures up to 3448 kN/sq m abs were tested. The final modified assembly was evaluated during combustion stability tests on rocket combustors burning a FLOX-methane propellant combination.
Review of the current state of analyses of combustion instability in solid-propellant rocket motors, citing appropriate measurements and observations. The work discussed has become increasingly important, both for the interpretation of laboratory data and for predicting the transient behavior of disturbances in full-scale motors. Two central questions are considered - namely, linear stability and nonlinear behavior. Several classes of problems are discussed as special cases of a general approach to the analysis of combustion instability. Application to motors, and particularly the limitations presently understood, are stressed.
A previously developed computer program was used to model two transient hybrid combustion processes involving tubes of solid Plexiglas. In the first study, representing combustion of a hybrid rocket, the oxidizing gas was oxygen, and calculations were continued sufficiently long to obtain steady-state values. Systematic variations were made in reaction rate constant, mass flow rate, and pressure, alternatively using constant and temperature dependent regression rate models for the fuel surface. Consistent results were obtained, as is evidenced by the values for the mass function of the reaction product and the flame temperature, for which plots are supplied. In the second study, fire initiation in a duct was studied, with an air mixture as the oxidizing gas. It was demonstrated that a satisfactory flame spread mechanism could be reproduced on the computer. In both of the above applications, the general, transient, two-dimensional conservation equations were represented, together with chemical reactions, solid-fuel interface conditions, and heat conduction in the solid fuel.
Noise generation by a subsonic flow discharging from a combustion chamber is examined with regard to the relative importance of combustion as a source of noise in such a flow system. Measurements of pressure fluctuations inside the combustor are compared with far-field noise measurements by direct cross-correlations. The cross-correlations and derived cross-spectral densities verify that much of the noise originates inside the combustor. A first-order fluid-mechanical perturbation model is used to predict exit plane velocity fluctuations due to internal pressure fluctuations. Far-field noise levels estimated on this basis are in good agreement with measured values. The overall noise level from the combustor/jet is found to be 10 to 20 dB higher than for an equivalent clean, cold jet at the same exit velocity.