Mixture Ratio Distribution Research Articles

Optimization of Length of Mixing Zone in GH<sub>2</sub>-GO<sub>2</sub> Based Torch Igniter, Applying Numerical Techniques

Initiation of chemical reactions liberating thermal energy in non hypergolic propellants require ignition systems which are reliable as well as robust in nature. This necessitates the igniter and its subsystems be on par or exceed every quality criteria required for the main stage. The ignition systems which provide pilot action and flame holding in supersonic environment adds new challenges in its structural and thermal stability. The paper aims to conduct a numerical analysis of the Augmented Spark Torch igniter required for an air breathing rocket (SCRAMJET) engine with the emphasis on the optimization of the mixing length ratio required to achieve a uniform mixture ratio distribution in the zone of the igniter combustion chamber. Nomenclature

가스발생기 연료 및 산화제 매니폴드 유동해석을 통한 유량균일성 파악

펌프공급방식 액체로켓엔진 가스발생기의 매니폴드를 대상으로, 연료 및 산화제의 유량 균일성을 파악하기 위하여 유동해석을 수행하였다. 3가지의 연료링과 3가지의 분사기 헤드 형상에 대하여 유동해석을 수행함으로써, 유량균일성을 검토하였다. 이를 통해 최적의 유량균일성을 가지는 연료링과 분사기 헤드의 형상을 결정하였다. 또한 분사기 출구에서의 혼합비 분포를 유동해석 결과로부터 도출하였다. 이를 통하여, 연료와 산화제 모두 유량균일성이 우수한 형상을 결정할 수 있었다. Flow analyses have been performed to investigate the uniformity of propellant flow through the fuel and oxidizer manifolds of a full-scaled gas generator for a pump-fed liquid rocket engines. Injectors were simulated as porous medium layers having equivalent pressure drops. The uniformity of propellants has been analyzed for 3 fuel rings and 3 injector head configurations. The mixture ratio distribution at the exit of injectors has been estimated from the mass flow rates of fuel and oxidizer. The best configuration of fuel ring and injection head was selected through these flow analyses.

Mixing and combustion in a spinning combustor

The effect of spin on mixing and combustion in a circular diict with a primary air jet and secondary gas injection through the duct wail was studied using a spin rig, with spin rates up to 10,000 rpm. The spinning ducted flowfield was compared to stationary conditions. Spin enhanced the spreading rate of the pipe and orifice jets, resulting in a larger radial velocity component and an onset of a tangential component in the jet's core for the pipe jet and in the jet's shear layer for the orifice jet. The turbulence amplification rate was enhanced predominantly in the pipe jet. The tests showed that the duct-spin affects the flow even for a relatively short entrance duct. Measurement of the mixture ratio distribution using gas sampling probes in nonreacting tests were done with inert gas injection. Centrifugal forces affected the mixing depending on the weight of the gas relative to air. The mixing was enhanced for lighter-than-air gas and suppressed for heavier. Combustion tests included both ethylene as secondary gas and solid-fuel at the wall, In gaseous fuel combustion, the spin increased the combustion temperature due to improved mixing caused by low-molecular -weight products and reduced density. For solid-fuel combustion the spin had an adverse effect. Combustion temperatures were decreased due to reduced mixing caused by centrifugal forces acting on soot and heavy hydrocarbon products.

Optimization of mixture ratio distribution in liquid propellant rocket thrust chamber

An optimization problem is set up for determining the propellant mixture ratio distribution in the liquidpropellant rocket thrust chamber that would yield maximum specific impulse. The problem is solved using the method of Lagrangian Multipliers. The results show that uniform mixture ratio distribution in the combustion chamber will give a maximum value of specific impulse for mixture ratios greater than a critical value. Normal operating mixture ratios for liquid-propellant rocket engines are well above this critical value.

An experimental technique for measurement of injector spray mixing

Introduction P studies have related liquid-rocket spray characteristics to injector mechanical and hydraulic parameters and/or combustion efficiencies from hot firings, i.e., to measured performance, but none has predicted combustion efficiency from spray measurements. This Note presents a technique that combines spray data obtained with nonreactive propellant simulants with an analytical model to predict the loss in combustion efficiency due to nonuniformities in the propellant mixture ratio distribution. Efficiencies thus predicted are compared with hot-firing results. The combustion chamber is considered to be divided into a primary mixing zone that includes the initial combustion region and that terminates when the gas velocity has essentially turned the spray in the axial direction, and a secondary zone, wherein turbulent vaporization and combustion occur within stream tubes, but with no thermal or mass transfer between stream tubes. c) V FROM 1-D COMBUSTION MODEL gas _ £ = 3.0; D 200/*

Mixture ratio distribution - Its impact on rocket thrust chamber performance.

Irregular mixture ratio (Q/F) distribution across the face of an injector can result from unequal hydraulic resistances to injector elements and by chamber cooling requirements. The degree to which O/F distribution will influence specific impulse (7sp) depends upon the propellant system and operating conditions, the extent of the O/F deviations, and the degree of turbulent mixing in the combustion chamber; /sp losses up to 5% have been observed. The objective of this work is to show how these effects can be taken into account in rocket thrust chamber designs by use of a stream-tube analysis. Evaluations of three rocket thrust chamber systems utilizing either the liquid-liquid or gas-liquid injector designs are illustrated. Experimental results are shown which verify predicted performance gains due to control of the propellant distribution in the injector.

Monitoring the combustion process in large engines.

References 1 Valentine, R. S., Dean, L. E., and Pieper, J. L., An improved method for rocket engine performance prediction, J. Spacecraft Rockets 3, 1409-1414 (1966). 2 Crisman, P. A., Goldwasser, S. R., and Petrozzi, P. J., A general computer program for calculation of rocket performance parameters of propellants containing the atomic species, C, H, N, F and Cl, Proceedings of the ARS Propellant Thermodynamics and Handling Conference (American Rocket Society, New York, 1960), Special Rept. 12, pp. 293-313. 3 DeLeuze, N. C., Walsh, T. J., and Petrozzi, P. J., Calculation of chemical kinetic losses in a rocket nozzle, AerojetGeneral Corp., Thrust Chamber Engineering Rept. 9642:0046 (January 6, 1966). 4 Pieper, J. L., Dean, L. E., and Valentine, R. S., Mixture ratio distribution—its impact on rocket engine performance, Aerojet-General Corp., Thrust Chamber Engineering Rept. 9642: 0062/Revision B (December 21, 1966). 5 Rupe, J. H., A correlation between the dynamic properties of a pair of impinging streams and the uniformity of mixtureratio distribution in the resulting spray, Jet Propulsion Lab. Progress Rept. 20-209 (March 28, 1956). 6 Beltran, M. R., Breen, B. P., Kosvic, T. C., Sanders, C. F., Hoffman, R. J., and Wright, R. O., Liquid rocket engine combustion instability studies, Dynamic Science Final Report, Contract AF 04(611)-10542, AFRPL-TR-66-125 (July 1, 1966).

An improved method for rocket performance prediction.

The prediction of flight performance of rocket engines has, in general, been dependent upon sea-level data that must be extrapolated to the flight configuration. Errors up to 4% in predicted specific impulse (/sp) resulted because of the inability to identify the factors theoretically affecting the extrapolation. Furthermore, predicted sea-level /sp for new designs was often in error by up to 5%. This paper describes a technique that allows /sp to be estimated within 2% for conventional propellant systems — at any exit area ratio — with no test data. If sea-level test data are available, altitude performance can be estimated within 1%. It is shown that the effects of mixture ratio distribution and interactions between injector /chamber performance and the nozzle are responsible for most of the errors associated with evaluating performance by the c* — CF method. Data for several rocket systems indicate a maximum error of 0.8% in comparing measured mean altitude performance to values predicted from sealevel data.

The Sampling of Rocket Combustion Gases

Summary:—In the past, attempts have been made to evaluate injectors for rocket engine combustion chambers by the use of water analogy rigs and model combustion systems that simulated the injection and combustion phase change occurring in the actual engine. To confirm that conditions in the engine were being correctly simulated, a technique was evolved for determining the mixture ratio distribution achieved by the combustion system of a Spectre variable thrust rocket engine. Gas samples extracted from the rocket-efflux were analysed, and the technique has been applied to evaluate the Spectre's standard central mushroom type injector and also a multi-head injector.Tests have been conducted over a thrust range of 2000 lb to 8000 lb and at oxidant/fuel ratios from 7·5 to 13·0.In parallel with this external sampling, a probe has been designed and developed for extracting gas samples from selected points across a diameter of the combustion chamber itself. This probe has been successfully operated for several minutes under combustion conditions of 500 p.s.i.a. and 2600°K, without sustaining any damage.Analysis of the oxidant/fuel ratio pattern within the combustion chamber and in the efflux, at comparable operating conditions, indicates that little change in distribution occurs between these two points of the system. Also, the distribution found with the standard injector was that for which the combustion system was designed. It is demonstrated that loss of performance depends on the degree of non-uniformity of mixing. A 5 per cent loss in performance at full thrust and optimum mixture ratio occurs with the standard injector.