Showing papers in "Journal of Propulsion and Power in 1993"

Research on supersonic combustion

Abstract: Introduction S OME qualifications to the pretentious nature of the title of this Paper are in order. Research in this context is applied, i.e., it is directed toward the design and development of devices. The devices are air breathing propulsion systems, wherein supersonic combustion is inherent or it provides an adjunct benefit. Principal applications of the former are manned hypersonic aircraft, transatmospheric accelerators, and missiles. Typical examples of the latter are external burning systems that sustain thrust or reduce drag when used in tandem with primary accelerator engines. This Paper is not at all representative of the complete body of the research on supersonic combustion. Instead, what is presented is primarily based on material with which the writer has had either direct involvement or a first-hand knowledge. It does not do justice to the exemplary works of many other investigators. Hopefully, this is somewhat rectified by the extensive list of references. The reader is encouraged to consult these manuscripts to obtain a more balanced perspective. This applied research must be viewed from the perspective of the engineer who is charged with expediting development and is in the need of design tools. Exact physics have frequently been abandoned to produce functional models. Much of the experimental verification upon which the models are based has been obtained in facilities which have known imperfections with respect to duplication of flight conditions. Moreover, compromises have been made in the analysis and interpretation of the experimental data, in part due to imperfect or incomplete diagnostic instrumentation, and in part due to a limited understanding of the underlying physics. An example that embodies all of these deficiencies would be a design model that is based on the use of integral techniques to describe combustion in supersonic flow from experiments conducted in an arc-heated tunnel.

Bare wire anodes for electrodynamic tethers

Abstract: The collection of electrons from the ionosphere is the major problem facing high-power electrodynamic tethers. This article discusses a simple electron-collection concept which is free of most of the physical uncertainties associated with plasma contactors in the rarefied, magnetized environment of an orbiting tether. The idea is to leave exposed a fraction of the tether length near its anodic end, such that, when a positive bias develops locally with respect to the ambient plasma, and for a tether radius small compared with both thermal gyroradius and Debye length, electrons are collected in an orbital-motion-limited regime. It is shown that large currents can be drawn in this way with only moderate voltage drops. The concept is illustrated through a discussion of performance characteristics for generators and thrusters.

Comprehensive review of liquid-propellant combustion instabilities in f-1 engines

Abstract: disparity in time and length scales exist in close proximity to one another. Although advances in the field have been made, the largest and most reliable source of information to date applicable to the design of improved combustion devices is the store of experimental data from full-scale engine tests. Consequently, the motivation behind the present work was to gain further insight into the mechanisms associated with combustion instability by providing a detailed, concise account and analyses of the design attributes which led to dynamic stability in F-l developmental injectors. Objectives were 1) to preserve the experience gained through development of the F-l engine; 2) to merge full-scale test results with corresponding theories and experiments; and 3) to analyze the effect of proposed solutions. To facilitate analysis, all available full-scale component and engine test data have been combined into a single data base. This compilation provides a complete genealogy of F-l developmental injector design configurations, and contains all available measured and observed test results. Table 1 lists the injector design parameters and test results acquired. The complete data base is available as an appendix to a separate technical report prepared by the authors.1 These data have been assembled from a variety of sources.2^ Reference 2 is a chronological tabulation of full-scale injector component test results recorded at the test site. This document lists the injectors tested along with the date, chamber pressure, thrust, run time, mixture ratio, bomb size, and damp times, as well as observations made during various tests. Reference 3 contains a set of 16 reports (four volumes of four reports each) which present a somewhat chronological account of the methodology leading to a dynamically stable injector design. Full-scale engine and component test results are discussed throughout this set of reports. Reference 4 provides a broad overview of the problems and solutions encountered with combustion instability in the F-l engine. Finally, Refs. 5 and 6 are weekly

Modeling of spray droplets deformation and breakup

Abstract: A droplet deformation and breakup (DDB) model is proposed to study shear-type mechanism of spray droplets in pure extentional flows. A numerical solution of the DDB model equation is obtained using a fourth-order Runge-Kutta initial-value solver. The predictions of the DDB model as well as semianalytical and the Taylor analogy models are compared with the experimental data (Krzeczkowski, 1980) for shear breakup, which depict the dimensionless deformation of the drop vs dimensionless time.

Hydraulic and Mechanical Performance of LE-7 LOX Pump Inducer

Abstract: A liquid oxygen turbopump has been developed for the main engine (LE-7) of the H-II rocket. The LE-7 LOX pump requires an inducer with quite high suction performance and high head, because a low-speed and low-pressure pump is not used ahead of the main pump in the LE-7 engine. The inducer was designed using the customary method, and its hydraulic and mechanical performances were investigated in tests of LE-7 LOX turbopumps. The original combination of an inducer and an inducer housing satisfied the required hydraulic performance criteria. However, this combination was found to result in supersynchronous shaft vibrations due to rotating cavitation which occurred in the inducer. This problem was almost completely solved by a simple modification of the inducer upstream housing. Furthermore, the rotating cavitation of the present inducer was investigated using a new theory of such cavitation. I. Introduction A LIQUID oxygen turbopump has been developed for the main engine (LE-7) of the H-II rocket, the next generation of Japanese launch vehicle. This turbopump requires a high-flow, high-pressure liquid oxygen pump. Because a low-speed, low-pressure pump is not used ahead of the main pump in the LE-7 engine, it is very important to operate the main pump at higher speed to obtain a smaller-size, lighterweight turbopump. The operational speed of the present turbopump was restricted by the suction performance of the main pump inducer. In this article, we report both the design and test results of the inducer. The main pump of the LE-7 LOX turbopump has a singlestage centrifugal impeller with an inducer. 1 The inducer is characterized by a low flow coefficient, a small inlet angle, a sharp leading edge, etc., to achieve higher suction performance. The inducer was designed using the customary method.2-3

Effects of fuel-air unmixedness on NO(x) emissions

Abstract: Unmixedness effects of lean fuel-air mixtures on emissions of nitrogen oxides (NOV) are experimentally studied. The objective is to study the relative influence of temporal fuel concentration fluctuations and spatial nonuniformities of mean concentration profiles on NO V emissions from lean premixed combustion. The NO2 laserinduced fluorescence (LIF) technique is used to quantify Unmixedness levels at the flameholder. Unmixedness data are related to NOr emissions measurements from downstream of the flames. The results show that temporal unmixedness contributes significantly to higher NOX emissions. It is not sufficient to eliminate only spatial nonuniformities of mean concentration in order to minimize NO_V emissions from lean premixed combustion. Nomenclature c = volumetric fuel concentration c' = rms fluctuation component of c c - time-mean fuel concentration Dc = combustor diameter Z)cf = coflow diameter Dj = jet diameter Recf = coflow Reynolds number Rej - jet Reynolds number U = unmixedness parameter Xj - distance from jet to flameholder z - vertical coordinate; origin at combustor centerline = fuel-air equivalence ratio

Time-marching method for the prediction of two-dimensional, unsteady flows of condensing steam

Abstract: A time-accurate, two-dimensional time-marching technique is presented which can predict unsteady phenomena in condensing steam flows. Conservation equations for the mixture are solved using a variation of a well-established Euler solver, while nucleation and droplet growth calculations are performed in a Lagrangian framework by tracking particle pathlines. A special averaging technique is used to retain a polydispersion of droplet sizes, necessary for the accurate modeling of the condensation processes, without consuming excessive storage or CPU time. The basic Euler solution technique has been validated by comparison with predictions from an independent source for the unsteady flow of air in a channel. The full scheme has been used to compute nucleating flows in converging-diverging nozzles for which agreement with experiment for both steady and unsteady cases is extremely good. All the results presented are for flows in nozzles for which experimental data are available, but the scheme may also be applied to turbine cascade geometries. 26 refs.

Unified Navier-Stokes flowfield and performance analysis of liquid rocket engines

Abstract: To improve the current composite solutions in the design and analysis of liquid propulsive engines, a computational fluid dynamics model capable of calculating the nonreacting and reacting flows from the combustion chamber, through the nozzle to the external plume, was developed The Space Shuttle Main Engine (SSME) fired at sea level, along with the flowfields of several other nozzles were investigated The bell-shaped SSME nozzle was run at 100 percent power level at various flow conditions, the computed flow results and performance compared well with those of other standard codes and engine hot fire test data

Plasma detachment from a magnetic nozzle

Abstract: Magnetic nozzles are used to convert plasma pressure into directed flow, thereby providing thrust for propulsion. The detachment of the flow from the magnetic field must be highly directed or the thrust efficiency will be low. A simple model of flowing, cold plasma in an axisymmetric nozzle is considered to determine the fundamental physics and scaling of separation in the limiting case that it is determined by plasma inertia. Ion and electron motions are coupled by the ambipolar electric field. In the limit of negligible meridional electric current, separation from the field is determined by the hybrid (electron-ion) Larmor radius evaluated at the initial magnetic flux and flow velocity. It is shown that the detachment is thereby significantly constrained, but that significant improvement can be obtained if electric currents can flow across the magnetic field either in walls or as plasma currents, or if dissipative effects are sufficiently strong.

Demonstration of mode transition in a scramjet combustor

Abstract: Direct-connect combustor hardware has been assembled at the Avery Propulsion Research Laboratory (APRL) of the Johns Hopkins University Applied Physics Laboratory (JHU/APL) to investigate a hydrogen-fueled scramjet combustor. The test hardware was designed to perform tests at simulated Mach 5 to Mach 8 flight conditions. This is done using a combustion heater with H2 fuel and makeup O2. The air, H2 and O2 flow rates are all supplied through computer-controlled digital valves. This system allows rapid changes in conditions and very steady flow rates can be maintained throughout the test. Recently, tests were performed in which the flow rates were systematically varied during the test to simulate an acceleration from M = 5.9 to 6.2. During this acceleration the fuel-air equivalence ratio was held constant and the combustor transitioned from a dual mode ramjet with a precombustion shock system creating subsonic flow at the injection plane, to a scramjet with no precombustion shock system. The results of these tests are presented along with descriptions of the hardware and control systems.

Magnesium and Carbon Dioxide - A Rocket Propellant for Mars Missions

Abstract: A rocket engine for Mars missions is proposed that could utilize CO2 accumulated from the Martian atmosphere as an oxidizer. For use as possible fuel, various metals, their hydrides, and mixtures with hydrogen compounds are considered. Thermodynamic calculations show that beryllium fuels ensure the most impulse but poor inflammability of Be and high toxicity of its compounds put obstacles to their applications. Analysis of the engine performance for other metals together with the parameters of ignition and combustion show that magnesium seems to be the most promising fuel. Ballistic estimates imply that a hopper with the chemical rocket engine on Mg + CO2 propellant could be readily developed. This vehicle would be able to carry out 2-3 ballistic flights on Mars before the final ascent to orbit.

Quasi-Three-Dimensional Nonreflecting Boundary Conditions for Euler Equations Calculations

Abstract: This article presents a theory for the construction of steady-state quasi-three-dimensional nonreflecting boundary conditions for the Euler equations. These allow calculations to he performed on truncated domains without the generation of spurious nonphysical reflections at the far-field boundaries. The theory is based upon Fourier analysis and eigenvectors applied to the linearized Euler equations. It is presented within the context of transonic axial flow turbomachinery computations. The effectiveness of the new boundary conditions is demonstrated by comparing results obtained using this new formulation and calculations performed with the standard onedimensional approach. c ht k,l P P Pt R

Some governing parameters of plasma torch igniter/flameholder in a scramjet combustor

Abstract: Effects of various operational parameters of plasma torch igniters/flameholders were experimentally studied in a hydrogen-fueled supersonic combustor with a rectangular cross section. The stagnation temperature of the airstream at the ignition limit almost linearly decreased as input electric power increased, although it did not significantly change with the flow rate of feedstock. Effectiveness of the argon plasma torch igniter was remarkably improved by adding a small fraction of hydrogen, but it showed rather modest improvement for further increase of the hydrogen contents. Modification of the combustor top wall design led to successful ignition of fuel jets injected from opposing side walls by a single plasma torch igniter and great reduction of airstream temperature at the ignition limit.

Computations and experiments for a multiple normal shock/boundary-layer interaction

Abstract: Results from a numerical investigation of a Mach 1.61 multiple normal shock wave/turbulent boundary-layer interaction are compared to wall static pressure and laser Doppler velocimeter measurements. The computations used the explicit, time-dependent, second-order accurate MacCormack scheme to solve the mass-averaged Navier-Stokes equations. Turbulence was modeled by means of the Baldwin-Lomax algebraic model and the Wilcox-Rubesin two-equation model. The computation with the Wilcox-Rubesin model was able to capture the major features of the normal shock train and accurately predicted the flow reacceleration mechanisms which occur between shocks. However, this computation failed to accurately predict the level of flow separation under the first shock. The Baldwin-Lomax computation displayed a more limited ability to capture the features of this shock train flow.

Multigrid calculation of three-dimensional viscous cascade flows

Abstract: A three-dimensional code for viscous cascade flow prediction has been developed. The space discretization uses a cell-centered scheme with eigenvalue scaling to weigh the artificial dissipation terms. Computational efficiency of a four-stage Runge-Kutta scheme is enhanced by using variable coefficients, implicit residual smoothing, and a full-multigrid method. The Baldwin-Lomax eddy-viscosity model is used for turbulence closure. A zonal, nonperiodic grid is used to minimize mesh distortion in and downstream of the throat region. Applications are presented for an annular vane with and without end wall contouring, and for a large-scale linear cascade. The calculation is validated by comparing with experiments and by studying grid dependency.

Study on vortex generator flow control for the management of inlet distortion

Abstract: The present study demonstrates that the reduced Navier-Stokes RNS3D code can be used very effectively to develop a vortex generator installation to minimize the engine face circumferential distortion by controlling secondary flow. The computing times required are small enough that studies such as this are feasible within an analysis-design environment with all its constraints of time and costs. This research study also established the nature of the performance improvements that can be realized with vortex flow control, and suggests a set of aerodynamic properties (called observations) that can be used to arrive at a successful vortex generator installation design. This study also indicated that scaling between flight and typical wind-tunnel test conditions is possible only within a very narrow range of generator configurations close to an optimum installation. Lastly, this study indicated that vortex generator installation design for inlet ducts is more complex than simply satisfying the requirement of attached flow, it must satisfy the requirement of minimum engine face distortion.

Erosive burning of solid propellants

Abstract: Presented here is a review of the experimental and modeling work concerning erosive burning of solid propellants (augmentation of burning rate by flow of product gases across a burning surface). A brief introduction describes the motor design problems caused by this phenomenon, particularly for low port/throat area ratio motors and nozzleless motors. Various experimental techniques for measuring crossflow sensitivity of solid propellant burning rates are described, with the conclusion that accurate simulation of the flow, including upstream flow development, in actual motors is important since the degree of erosive burning depends not only on local mean crossflow velocity and propellant nature, but also upon this upstream development. In the modeling area, a brief review of simplified models and correlating equations is presented, followed by a description of more complex numerical analysis models. Both composite and double-base propellant models are reviewed. A second generation composite model is shown to give good agreement with data obtained in a series of tests in which composite propellant composition and heterogeneity (particle size distribution) were systematically varied. Finally, the use of numerical models for the development of erosive burning correlations is described, and a brief discussion of scaling is presented.

Linear Burning Rate Dynamics of Solids Subjected to Pressure or External Radiant Heat Flux Oscillations

Abstract: Linearized analysis of burning solids subjected to pressure or external radiant heat flux oscillations results in relatively simple expressions for the burning rate response, particularly when combined with quasisteady gas and surface zone assumptions In this article, a linear expression for the radiant heat flux response function Rq as a function of primary experimental parameters is obtained The implications of this expression and its relationship to the pressure coupled response function Rp are examined In particular, the linearized effects of the mean radiant heat flux level and in-depth absorption on Rq are investigated The linear response to a series of radiant pulses is also presented to suggest an alternate method of experimentally measuring Rq The effects of nonlinearitie s are investigated using numerical calculations It is shown that, in general, the relationship between Rp and Rq is more complicated than a constant scaling factor The results demonstrate that in-depth absorption of the thermal radiant energy in the solid significantly affects Rq for many practical conditions Furthermore, even when the equivalence principle holds in the steady case, an equivalent change in the initial temperature does not have the same effect on Rq, in general, as an equivalent mean radiant flux Also, the mean radiant flux is seen to have a significant effect on Rq An attempt is made throughout this article to further clarify the relationship and differences between the flame modeling (FM) and the Zeldovich-Novozhilov (ZN) phenomenological approaches in the prediction of Rq and Rp

Pressure oscillations in post-Challenger Space Shuttle redesigned solid rocket motors

Abstract: Data from the first seven solid rocket motor (SRM) static tests were used to establish upper bounds for the maximum acoustic pressure amplitudes in the latter half of firing Those bounds have since been used as a basis for worst-case simulation scenarios by specialists in structural dynamics at NASA and Rockwell International, and to provide a basis for evaluating data from individual motors which were tested subsequent to the original seven SRMs The purpose of this article is to compare current motor data with data from the early motors and determine whether or not the upper bounds need to be recalculated Comparisons of chamber pressure amplitudes in redesigned solid rocket motors (RSRMs) with the predicted upper bounds, based on the original seven motors, initially suggest that the original bounds are no longer suitable descriptors for the current motor designs However, a rigorous statistical analysis of the SRM and RSRM motor data indicates that the data are from similar statistical populations

Experimental Supersonic Hydrogen Combustion Employing Staged Injection Behind a Rearward-Facing Step

Abstract: An experimental investigation of a Mach 2 combustor has been conducted in order to characterize flow properties in a supersonic reacting flowfield. Hydrogen was injected transversely as staged, underexpanded jets behind a rearward-facing step into a ducted Mach 2 air freestream. The effects of the chemical reaction on the supersonic flowfield was investigated using shadowgraphs, broadband flame emission photography, and planar laser-induced fluorescence of OH. The shadowgraphs indicated that the wave pattern in the combustor along with flowfield unsteadiness was strongly affected by the heat release. The broadband flame emission photographs revealed large regions of no combustion in the vicinity of the fuel injectors where fuel/air mixing was insufficient to support combustion. These regions decreased in size as the freestream stagnation temperature was decreased for fixed hydrogen mass flow rate, consistent with an increase in the effective g-ratio with combustion. The size of the zones containing OH in the planar fluorescence images also increased as the main flow stagnation temperature was decreased. Reaction zones were found in the planar fluorescence images away from regions containing inject ant in a nonreacting study of the same geometry, indicating that the pressure rise associated with the reaction forced a large redistribution of the fuel.

Static tests of jet fuel thermal and oxidative stability

Abstract: Jet fuels and a jet fuel surrogate have been thermally stressed to simulate the time/temperature history of aircraft fuel handling systems. Surrogate fuels were used to develop quantitative measurement techniques to assess fuel stability in static tests and compare the results with flowing tests. A variety of experimental techniques including Fourier transform infrared (IR), gas chromatography with atom-sensitive atomic-emission detection and high-pressure liquid chromatography have been used to study stressed and unstressed fuels. Quantitative and qualitative measurements of the deposits and the fuels are presented. In general, the static tests described here indicate that there is good agreement between static and flowing tests concerning the quality of a fuel. However, to adequately assess fuel stability, the availability of oxygen must be limited. Arbitrarily increasing the oxygen availability is likely to yield results which are not applicable to oxygen-starved stressing processes. Furthermore, contrary to expectations, the rate at which a fuel oxidizes is shown to be inversely related to the rate of formation of insoluble products.

Numerical Simulation of Turbine "Hot Spot" Alleviation Using Film Cooling

Abstract: Experimental data have shown that combustor hot streaks can lead to pressure side "hot spots" on firststage turbine rotor blades. In previous numerical studies, it has been shown that unsteady Navier-Stokes procedures can be used to predict the rotor pressure surface temperature increase associated with these combustor hot streaks. In the current investigation, similar two-and three-dimensional unsteady Navier-Strokes simulations have been performed to demonstrate the use of numerical tools in the optimization of film cooling configurations. In this study, the addition of prudently placed film cooling holes along the rotor pressure surface is shown to significantly diminish the adverse effects of the hot streak. Using a two-dimensional Navier-Stokes procedure, a parametric study was performed to determine the impact of the location of the film cooling holes, fluid injection velocity, and fluid injection angle on the time-averaged rotor surface temperature. The experience gained from these two-dimensional simulations was then applied to a series of three-dimensi onal simulations in which the effects of the film cooling hole distribution on the rotor pressure surface temperature were studied. The results of these simulations indicate that computational procedures can be used to design feasible film cooling schemes which eliminate the adverse effects of combustor hot streaks.

Effects of injector geometry on scramjet combustor performance

Abstract: An experiment was conducted to investigate the effect of injector/combustor geometry on combustion-induced peak wall pressure and associated upstream influence, as well as on mixing/combustion characteristics at an entrance Mach number of 2.5. The length of the constant area section downstream of injection orifices had a strong influence on the above-mentioned characteristics. However, the sweep of the rearward-facing steps on both side walls had little effect on these characteristics, nor did reversing them have any effect. The peak wall pressure and the length of the upstream influence agreed qualitatively with predictions of an analytical model and an empirical formula developed at Johns Hopkins University. Fuel jets injected from the model with the longest constant area section and the fuel equivalence ratio of unity, coalesced at a very early stage downstream of the fuel injection orifices. This coalescence led to a decrease in mixing rate downstream, despite the higher degree of mixing near the injection orifices. The combustion efficiencies were higher than those obtained at NASA Langley in the upstream region due to the higher mixing rate near the injection orifices.

Photovoltaic receivers for laser beamed power in space

Abstract: One of the most promising beamed power concepts uses a laser beam to transmit power to a remote photovoltaic array. Large lasers can be located on cloud-free sites at one or more ground locations, and illuminate solar arrays to a level sufficient to provide operating power. Issues involved in providing photovoltaic receivers for such applications are discussed. >

Evaluation of brush seals for limited-life engines

Abstract: Brush seals are a relatively new concept for replacing labyrinth seals in gas turbine engines. An evaluation was performed to assess the potential of brush seals for limited-life gas turbine engines. A rotating rig was designed and built to test labyrinth and brush seals over a range of simulated engine conditions. An initial set of brush seals was rig-tested to determine leakage and wear performance and identify potential optimum configurations. The measured results showed that brush seals offer significant improvements over labyrinth seals with a factor of three or more reduction in leakage flow. Brush seals exhibit an initial wear-in period but retain significantly reduced leakage over labyrinth seals for times exceeding most limited-life engine applications. Consequently, brush seals offer the potential to precisely meter cooling/leakage air, thereby decreasing parasitic leakage and improving fuel consumption and thrust. Thus, brush seals are a definite candidate for replacing labyrinth seals in gas turbine engines.

Application of a quadruple probe technique to MPD thruster plume measurements

Abstract: A new quadruple Langmuir probe diagnostic technique is applied to the plume of a pulsed magnetoplasmadynamic (MPD) thruster with argon propellant. The probe permits simultaneous measurement of electron temperature Te, density ne, and ion flow velocity u in the exhaust stream. For distances of 3-25 cm from the thruster face, Te is 8-1.5 eV, ne is 4.8-1.2 x 10 m~, and u is 0.7-1.8 x 10 m/s. The measurements suggest that the plume is well-collimated, and that the ion temperature is significantly higher than the electron temperature, consistent with an ion-conduction model of the thruster.

New Two-Grid Acceleration Method for Unsteady Navier -S tokes Calculations

Abstract: A quasi-three-dimensional time-marching Navier-Stokes method for calculating unsteady viscous flows in turbomachines is presented. A major feature of the present work is that the time-step limitation in the NavierStokes solutions suffered by all explicit time-marching methods is effectively relaxed by using a time-consistent two-grid method. The spatial accuracy is subject to the basic fine mesh, while the coarse mesh, on which the temporal accuracy is guaranteed, is locally applied to the near wall and wake regions to increase the allowable time-step length. The loss of the time accuracy on the basic fine mesh can be easily controlled by choosing a suitable grid size of the coarse mesh according to the wavelength of physical unsteadiness to be dealt with. This two-grid method has been compared with the implicit residual-averaging method and the direct time-marching method for a transonic oscillating cascade flow. Numerical examples for a low-speed oscillating airfoil flow at a dynamic stall condition and a transonic airfoil flow with a self-excited shock oscillation are also presented, in which an increase in the time-step length by a factor of 20 has been achieved.

Some aspects of secondary atomization of aluminum/hydrocarbon slurry propellants

Abstract: A theoretical and experimental investigation of the secondary atomization of Al/RP-1 slurry propellants was conducted. Theoretical efforts examined the effects of aluminum particle size in particle shell/crust formation on the surface of a burning slurry droplet. A model was developed for the pressure buildup and mechanical stresses in the shell for the time period after the shell becomes impermeable. Experimental efforts focused on the ignition and combustion of a dilute stream of slurry droplets. Measurements of individual particle diameters and velocities were made at various axial locations in the stream. Radiant emission from the particles was also monitored to determine whether aluminum was actively burning. Experimental aluminum agglomerate ignition times were found to be comparable to theoretical estimates. The experimentally measured minimum initial slurry droplet diameter for secondary atomization was 20-25 /*m, compared to a theoretical diameter of 35 /an.