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In compressorless Diesel engines and in explosion engines using fuels with high boiling points it is difficult to effect a good combustion of the fuel mixture. This report presents different methods for calculating the size and uniformity of fuel droplets and mixtures.
Different methods of determining the size of fuel droplets in the atomization process are presented. The distribution of those drops is also presented in graphic form.
One of the technology goals of NASA for advanced space transportation is to develop highly efficient propulsion systems to reduce the cost of payload for space missions. Developments of rockets for the second generation Reusable Launch Vehicle (RLV) in the past several years have been focused on low-cost versions of conventional engines. However, recent changes in the Integrated Space Transportation Program to build a crew transportation vehicle to extend the life of the Space Shuttle fleet might suggest that air-breathing rockets could reemerge as a possible propulsion system for the third generation RLV to replace the Space Shuttle after 2015. The weight of the oxygen tank exceeds thirty percent of the total weight of the Space Shuttle at launch while the payload is only one percent of the total weight. The air-breathing rocket propulsion systems, which consume oxygen in the air, offer clear advantages by making vehicles lighter and more efficient. Experience in the National Aerospace Plane Program in the late 1980s indicates that scramjet engines can achieve high specific impulse for low hypersonic vehicle speeds. Whether taking a form of Rocket Based Combined Cycle (RBCC) or Turbine Based Combined Cycle (TBCC), the scramjet is an essential mode of operation for air-breathing rockets. It is well known that fuel-air mixing and rapid combustion are of crucial importance for the success of scramjet engines since the spreading rate of the supersonic mixing layer decreases as the Mach number increases. A factored form of the Gauss-Seidel relaxation method has been widely used in hypersonic flow research since its first application to non-equilibrium flows. However, difficulties in stability and convergence have been encountered when there is strong interaction between fluid motion and chemical reaction, such as multiple fuel injection problems. The present paper reports the results from investigation of the effect of modifications to the original algorithm on the performance for multiple injectors.
The feasibility of the generation of detonation waves moving periodically across high speed channel flow is experimentally demonstrated. Such waves are essential to the concept of compressing requirements and increasing the engine pressure compressor with the objective of reducing conventional compressor requirements and increasing the engine thermodynamic efficiency through isochoric energy addition. By generating transient transverse waves, rather than standing waves, shock wave losses are reduced by an order of magnitude. The ultimate objective is to use such detonation ducts downstream of a low pressure gas turbine compressor to produce a high overall pressure ratio thermodynamic cycle. A 4 foot long, 1 inch x 12 inch cross-section, detonation duct was operated in a blow-down mode using compressed air reservoirs. Liquid or vapor propane was injected through injectors or solenoid valves located in the plenum or the duct itself. Detonation waves were generated when the mixture was ignited by a row of spark plugs in the duct wall. Problems with fuel injection and mixing limited the air speeds to about Mach 0.5, frequencies to below 10 Hz, and measured pressure ratios of about 5 to 6. The feasibility of the gas dynamic compression was demonstrated and the critical problem areas were identified.
This report summarizes some results obtained with a single cylinder test engine at the Langley Field Laboratory during a preliminary investigation of the problem of applying fuel injection and compression ignition to aircraft engines. For this work a standard Liberty Engine cylinder was fitted with a high compression, 11.4 : 1 compression ratio, piston, and equipped with an airless injection system, including a primary fuel pump, an injection pump, and an automatic injection valve. The results obtained during this investigation have indicated the possibility of applying airless injection and compression ignition to a cylinder of this size, 8-inch bore by 7-inch stroke, when operating at engine speeds as high as 1,850 R. P. M. A minimum specific fuel consumption with diesel engine fuel oil of 0.30 pound per I. HP. Hour was obtained when developing about 16 B. HP. At 1,730 R. P. M.
An advance study of materials, fuel injection, and hot flow problems related to the gas core nuclear rocket is reported. The first task was to test a previously constructed induction heated plasma GCNR simulator above 300 kW. A number of tests are reported operating in the range of 300 kW at 10,000 cps. A second simulator was designed but not constructed for cold-hot visualization studies using louvered walls. A third task was a paper investigation of practical uranium feed systems, including a detailed discussion of related problems. The last assignment resulted in two designs for plasma nozzle test devices that could be operated at 200 atm on hydrogen.
A hybrid implicit/explicit approach for analyzing 2D ducted turbulent mixing problems with finite-rate chemistry is presented. The approach combines a fully-implicit parabolic mixing algorithm, an explicit viscous-characteristic based shock-capturing algorithm, and linearized implicit chemical kinetic algorithm. The resultant model provides spatial marching solutions of the parabolized Navier-Stokes (PNS) equations for supersonic combustion and viscous nozzle flowfields. Specialized procedures are incorporated to deal with the near wall sublayer and with small central pockets of subsonic flow. A parabolic option is available which can be utilized in the direct or inverse mode for design applications. The ability of the model to treat shock waves and wave/mixing layer interactions is assessed via comparisons of predictions with those of well established explicit shock-capturing Euler (SCIPPY) and PNS (SCIPVIS) models. Applications to a variety of supersonic combustion flowfield problems involving tangential or moderately inclined fuel injection, and, to a viscous nozzle flow problem, are presented. This paper serves to exhibit overall capabilities of the model developed and no comparisons with supersonic combustion data are presented. Such comparisons serve mainly to validate the modeling of the turbulence and turbulence/chemical interactions and will be the subject of a future paper.
The SIMPLE-based parabolic flow code, SHIP3D, was under development for use as a parametric design and analysis tool for scramjets. Some capabilities and applications of the code are demonstrated, and a report on its current status is given. The focus is on the combustor for which the code was mostly used. Recently, it was also applied to nozzle flows. Code validation results are presented for combustor unit problems involving film cooling, transverse fuel injection, and nozzle test. A parametric study of a film cooled or transpiration cooled Mach 16 combustor is also conducted to illustrate the application of the code to a design problem.
Problems concerned with the supercritical injection concept are discussed. Supercritical injection involves dissolving air into a fuel prior to injection. A similar effect is obtained by preheating the fuel so that a portion of the fuel flashes when its pressure is reduced. Flashing improves atomization properties and the presence of air in the primary zone of a spray flame reduces the formation of pollutants. The investigation is divided into three phases: (1) measure the solubility and density properties of fuel/gas mixtures, including Jet A/air, at pressures and correlate these results using theory; (2) investigate the atomization properties of flashing liquids, including fuel/dissolved gas systems. Determine and correlate the effect of inlet properties and injector geometry on mass flow rates, Sauter mean diameter and spray angles; (3) examine the combustion properties of flashing injection in an open burner flame, considering flame shape and soot production.
In this project on the first stage (2000-Ol), we continued to develop the previous joint research between the Fluid Mechanics and Acoustics Laboratory (FM&AL) at Hampton University (HU) and the Jet Noise Team (JNT) at the NASA Langley Research Center (NASA LaRC). At the second stage (2001-03), FM&AL team concentrated its efforts on solving of problems of interest to Glenn Research Center (NASA GRC), especially in the field of propulsion system enhancement. The NASA GRC R&D Directorate and LaRC Hyper-X Program specialists in a hypersonic technology jointly with the FM&AL staff conducted research on a wide region of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analyses for advanced aircraft and rocket engines. The last year the Hampton University School of Engineering & Technology was awarded the NASA grant, for creation of the Aeropropulsion Center, and the FM&AL is a key team of the project fulfillment responsible for research in Aeropropulsion and Acoustics (Pillar I). This work is supported by joint research between the NASA GRC/ FM&AL and the Institute of Mechanics at Moscow State University (IMMSU) in Russia under a CRDF grant. The main areas of current scientific interest of the FM&AL include an investigation of the proposed and patented advanced methods for aircraft engine thrust and noise benefits. This is the main subject of our other projects, of which one is presented. The last year we concentrated our efforts to analyze three main problems: (a) new effective methods fuel injection into the flow stream in air-breathing engines; (b) new re-circulation method for mixing, heat transfer and combustion enhancement in propulsion systems and domestic industry application; (c) covexity flow The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analyses for advanced aircraft and rocket engines (see, for example, Figures 4). The FM&AL Team uses analytical methods, numerical simulations and experimental tests at the Hampton University campus, NASA and IM/MSU.
A survey is presented of the major small-engine manufacturers and governmental users. A consensus was undertaken regarding small-combustor requirements. The results presented are based on an evaluation of the information obtained in the course of the study. The current status of small-combustor technology is reviewed. The principal problems lie in liner cooling, fuel injection, part-power performance, and ignition. Projections of future engine requirements and their effect on the combustor are discussed. The major changes anticipated are significant increases in operating pressure and temperature levels and greater capability of using heavier alternative fuels. All aspects of combustor design are affected, but the principal impact is on liner durability. An R&D plan which addresses the critical combustor needs is described. The plan consists of 15 recommended programs for achieving necessary advances in the areas of liner thermal design, primary-zone performance, fuel injection, dilution, analytical modeling, and alternative-fuel utilization.
A critique of the various characteristics of engine design influencing noise and attempts to indicator areas where attention is required to obtain noise acceptable engine for automobiles are discussed. It was concluded that the engine has a potential to be quiet beccause a ion rated speed is chosen. Problems with high gas pressure, the fuel injection pump, and the expander/compressor are discussed.
In the present treatise we will consider chiefly the problem of solid injection in comparison with air injection. On leaving the valve or nozzle through one or more small openings, the fuel is split up into innumerable fine drops, which penetrate the combustion chamber in divergent directions in the form of a conical jet. The efficiency of this jet is judged from the following three viewpoints: 1) with respect to the fineness of atomization; 2) with respect to the direction or distribution of sprayed particles; 3) with respect to the penetration of the particles.
Results of theoretical and applied research related to fissioning uranium plasmas. Topics examined include uranium plasma instabilities, diagnostic techniques, radiant heat transfer characteristics, nuclear pumping of lasers, and various fission engine concepts. Engineering aspects considered for open-cycle gas core engines include effects of buoyancy on fuel containment, flow and criticality problems, effects of injection conditions, and radiation hazards. The nuclear light bulb engine, the mini-cavity reactor, the dust-bed reactor, and the colloid core reactor are also examined in terms of design and control considerations. Individual items are abstracted in this issue.
Further progress toward the satisfactory solution of the difficult problem of the distribution and atomization of the injected fuel was made by extensive experimentation with various fuel valves, nozzles, and atomizing devices. Valuable information was also obtained through numerous experimental researches on the combustion of oils and the manner of introducing the combustion air into the cylinder, as well as on the physical processes of atomization, the determination of the size of drops, etc. These researches led to the conclusion that it is possible, even without producing great turbulence in the combustion chamber and at moderate pump pressure, if the degree of atomization and the penetrative power of the fuel jet are adapted to the shape of the combustion chamber and to the dimensions of the cylinder.
Three connected sub-projects were conducted under reported project. Partially, these sub-projects are directed to solving the problems conducted by the HU/FM&AL under two other NASA grants. The fundamental idea uniting these projects is to use untraditional 3D corrugated nozzle designs and additional methods for exhaust jet noise reduction without essential thrust lost and even with thrust augmentation. Such additional approaches are: (1) to add some solid, fluid, or gas mass at discrete locations to the main supersonic gas stream to minimize the negative influence of strong shock waves forming in propulsion systems; this mass addition may be accompanied by heat addition to the main stream as a result of the fuel combustion or by cooling of this stream as a result of the liquid mass evaporation and boiling; (2) to use porous or permeable nozzles and additional shells at the nozzle exit for preliminary cooling of exhaust hot jet and pressure compensation for non-design conditions (so-called continuous ejector with small mass flow rate; and (3) to propose and analyze new effective methods fuel injection into flow stream in air-breathing engines. Note that all these problems were formulated based on detailed descriptions of the main experimental facts observed at NASA Glenn Research Center. Basically, the HU/FM&AL Team has been involved in joint research with the purpose of finding theoretical explanations for experimental facts and the creation of the accurate numerical simulation technique and prediction theory for solutions for current problems in propulsion systems solved by NASA and Navy agencies. The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analysis for advanced aircraft and rocket engines. The F&AL Team uses analytical methods, numerical simulations, and possible experimental tests at the Hampton University campus. We will present some management activity and theoretical numerical simulation results obtained by the FM&AL Team in the reporting period in accordance with the schedule of the work.
A series of multispecies, multiphase computational fluid dynamics (CFD) analyses of the 24-inch diameter joint government industry industrial research and development (JIRAD) hybrid rocket motor is described. The 24-inch JIRAD hybrid motor operates by injection of liquid oxygen (LOX) into a vaporization plenum chamber upstream of ports in the hydroxyl-terminated polybutadiene (HTPB) solid fuel. The injector spray pattern had a strong influence on combustion stability of the JIRAD motor so a CFD study was initiated to define the injector end flow field under different oxidizer spray patterns and operating conditions. By using CFD to gain a clear picture of the flow field and temperature distribution within the JIRAD motor, it is hoped that the fundamental mechanisms of hybrid combustion instability may be identified and then suppressed by simple alterations to the oxidizer injection parameters such as injection angle and velocity. The simulations in this study were carried out using the General Algorithm for Analysis of Combustion SYstems (GALACSY) multiphase combustion codes. GALACSY consists of a comprehensive set of droplet dynamic submodels (atomization, evaporation, etc.) and a computationally efficient hydrocarbon chemistry package built around a robust Navier-Stokes solver optimized for low Mach number flows. Lagrangian tracking of dispersed particles describes a closely coupled spray phase. The CFD cases described in this paper represent various levels of simplification of the problem. They include: (A) gaseous oxygen with combusting fuel vapor blowing off the walls at various oxidizer injection angles and velocities, (B) gaseous oxygen with combusting fuel vapor blowing off the walls, and (C) liquid oxygen with combusting fuel vapor blowing off the walls. The study used an axisymmetric model and the results indicate that the injector design significantly effects the flow field in the injector end of the motor. Markedly different recirculation patterns are observed in the vaporization chamber as the oxygen velocity and/or spray pattern is varied. The ability of these recirculation patterns to stabilize the diffusion flame above the surface of the solid fuel gives a plausible explanation for the experimentally determined combustion stability characteristics of the JIRAD motor, and suggests how combustion stability can be assured by modifications to the injector design.
The development of a computer program for the analytical prediction of the distribution of liquid and vapor fuel in the premixing-prevaporizing passage by the direct injection method is described. The technical approach adopted for this program is to separate the problem into three parts each with its own computer code. These three parts are: calculation of the two-dimensional or axisymmetric air flow; calculation of the three-dimensional fuel droplet evaporation; and calculation of the fuel vapor diffusion. This method of approach is justified because premixing passages operate at lean equivalence ratios. Hence, a weak interaction assumption can be made wherein the airflow can affect the fuel droplet behavior but the fuel droplet behavior does not affect the airflow.