Showing papers in "Journal of Propulsion and Power in 2007"

Reaction Propagation of Four Nanoscale Energetic Composites (Al/MoO3, Al/WO3, Al/CuO, and B12O3)

Abstract: Nanoscale composite energetics (also known as metastable intermolecular composites) represent an exciting new class of energetic materials. Nanoscale thermites are examples of these materials. The nanoscale thermites studied consist of a metal and metal oxide with particle sizes in the 30-200 nm range. They have potential for use in a wide range of applications. The modes of combustion and reaction behavior of these materials are not yet well understood. This investigation considers four different nanoaluminum/metal-oxide composites. The same nanoscale aluminum was used for each composite. The metal oxides used were molybdenum oxide (MoO 3 ), tungsten oxide (WO 3 ), copper oxide (CuO), and bismuth oxide (Bi 2 O 3 ). The reaction performance was quantified by the pressure output and propagation velocity using unconfined (or open burn) and confined (burn tube) experiments. We examine the optimization of each composite in terms of pressure output and propagaton speed (or burn rate) for the open burn experiment. We find that there is a correlation between the maximum pressure output and optimum propagation speed (or burn rate). Equilibrium calculations are used to interpret these results. We find that the propagation speed depends on the gas production and also on the thermodynamic state of the products. This suggests that condensing gases or solidifying liquids could greatly enhance heat transfer. We also vary the density of these composites and examine the change in performance. Although the propagation wave is likely supersonic with respect to the mixture sound speed, the propagation speed decreases with density. This behavior is opposite of classical detonation in which propagation (detonation) speed increases with density. This result indicates that the propagation mechanism may differ fundamentally from classical detonations.

Laminar Flame Speeds of Preheated iso-Octane/O2/N2 and n-Heptane/O2/N2 Mixtures

Abstract: Laminar flame speed measurements are carried out for premixed i so-octane/air and n-heptane/air mixtures under conditions of atmospheric pressure, equivalence ratios ranging from 0.7 to 1.4, and unburned mixture temperatures of 298, 360, 400, and 470 K using the counterflow flame technique. These experiments employ the digital particle image velocimetry technique to characterize the two-dimensional flow field upstream of the flame. As such, the reference stretch-affected flame speed and the imposed stretch rate can be simultaneously determined. By systematically varying the imposed stretch rate, the corresponding laminar flame speed is obtained by linearly extrapolating to zero stretch rate. In addition, the effect of nitrogen dilution level on the laminar flame speed is investigated by varying the nitrogen molar percentage in the oxidizer mixture from 78.5 to 80.5%. These results are further used for the determination of overall activation energies at different equivalence ratios. The experimental laminar flame speeds are subsequently compared with the computed values using two iso-octane reaction mechanisms and two n-heptane reaction mechanisms available in the literature, followed by discussion and sensitivity analysis.

Combustion of Nanoscale Al/MoO3 Thermite in Microchannels

Abstract: : Microscale combustion is of interest in small-volume energy-demanding systems, such as power supplies, actuation, ignition, and propulsion Energetic materials can have high burning rates that make these materials advantageous, especially for microscale applications in which the rate of energy release is important or in which air is not available as an oxidizer In this study we examine the combustion of mixtures of nanoscale aluminum with molybdenum trioxide in microscale channels Nanoscale composites can have very high burning rates that are much higher than typical materials Quartz and acrylic tubes are used Rectangular steel microchannels are also considered We find that the optimum mixture ratio for the maximum propagation rate is aluminum rich We use equilibrium calculations to argue that the propagation rate is dominated by a convective process where hot liquids and gases are propelled forward heating the reactants This is the first study to report the dependence of the propagation rate with a tube diameter for this class of materialsWefind that the propagation rate decreases linearly with 1=d The propagation rate remains high in tubes or channels with dimensions down to the scale of 100 m, which makes these materials applicable to microcombustion applications

Near-Blowoff Dynamics of a Bluff-Body Stabilized Flame

Abstract: Thispaperdescribesanexperimental investigation of thedynamicsof anear-blowoff, bluff-body stabilized flame. This work is motivated by a number of prior observations showing that near-blowoff flames exhibit enhanced unsteadiness. Laser sheet imaging studies and particle image velocimetry velocity field measurements show that the transient dynamics of these flames occur in two distinct stages before blowoff. The first stage is manifested by the presence of localized “holes” in the flame sheet, at locations where the instantaneous stretch rate exceeds the extinction stretch rate. During this stage, the overall flame and wake dynamics appear essentially unaltered and, moreover,the flamecanpersistindefinitely,althoughwithenhancedunsteadiness.Astheequivalenceratioisfurther decreased, the size of the flame region affected and the duration of these events increases monotonically. As the blowoff point is approached further, this leads to the second stage, large-scale alterations of the wake dynamics, violent flapping of the flame front, and even larger straining of the flame. In some cases, the flow in this second stage bears striking resemblance to the asymmetric von Karman type flowfield.

LaB6 Hollow Cathodes for Ion and Hall Thrusters

Abstract: Deep space missions and satellite station-keeping applications continue to demand higher power ion thrusters and Hall thrusters capable of providing high thrust and longer life. Depending on the thruster size, the hollow cathodes may be required to produce discharge currents in the 10-100 A range with lifetimes in excess of 10 years. A lanthanum hexaboride (LaB 6 ) hollow cathode has been developed for space applications to increase the current capability from the cathode and ease the handling and gas purity requirements. This cathode uses a LaB 6 insert in an all-graphite hollow cathode structure with an integral graphite keeper. Three different sizes of the LaB 6 cathode have been successfully operated at discharge currents of up to 100 A to date. Although the LaB 6 cathode insert operates at a higher temperature than the conventional BaO dispenser cathode, LaB 6 offers the capability of long life and orders of magnitude less sensitivity to propellant impurities and air exposure than conventional dispenser cathodes.

Processing and Ignition Characteristics of Aluminum-Bismuth Trioxide Nanothermite System

Abstract: During the past few years, significant progress has been made in the development of new nanoenergetic materials consisting of mixtures of metal and oxidizer nanopowders. It has been found that such reacting mixtures release energy by 2 to 3 orders of magnitude faster than similar systems consisting of micron-size reactants. In some cases, combustion-front velocities reach hundreds of meters per second. These new reacting systems find applications in both civilian and military sectors, including fast vaporization of active chemical components, fast heating of thermal batteries and main ingredients of new environmentally benign percussion primers or electric matches. This paper presents experimental results on ignition and combustion front propagation characteristics in the Al-Bi 2 O 3 nanothermite system. The effect of different coatings and inhibitors of the reaction of aluminum with water in the presence and without the presence of bismuth trioxide is discussed. In addition, thermodynamic analysis of the Al-Bi 2 O 3 reacting system and reaction kinetics measurements using differential scanning calorimetry are presented. Electrostatic discharge sensitivities of Al-Bi 2 O 3 , Al-MoO 3 , and Al-Fe 2 O 3 nanothermite systems were determined. It was found that all investigated nanothermite systems showed very high levels of electrostatic discharge sensitivity in the form of dry and loose powder.

Powered flight of electron cyclotron resonance ion engines on hayabusa explorer

Abstract: The electron cyclotron resonance ion engine has long life and high reliability because of electrodeless plasma generation in both the ion generator and the neutralizer. Four μ10s, each generating a thrust of 8 mN, specific impulse of 3200 s, and consuming 350 W of electric power, propelled the Hayabusa asteroid explorer launched on May 2003. After vacuum exposure and several baking runs to reduce residual gas, the ion engine system established continuous acceleration. Electric propelled delta-V Earth gravity assist, a new orbit change scheme that uses electric propulsion with a high specific impulse was applied to change from a terrestrial orbit to an asteroid-based orbit. In 2005, Hayabusa, using solar electric propulsion, managed to successfully cover the solar distance between 0.86 and 1.7 AU. It rendezvoused with, landed on, and lifted off from the asteroid Itokawa. During the 2-year flight, the ion engine system generated a delta-V of 1400 m/s while consuming 22 kg of xenon propellant and operating for 25,800 h.

Overview on the Aircraft Particle Emissions Experiment

Abstract: and10-m sampling rakes contained multiple gasandparticle inlet probes to facilitate astudy of the spatial variation of emissions across the engine exhaust plane. Gas-phase emission indices measured at 1 m were in good agreement withtheenginecertificationdataaswellaswithpredictionsprovidedbytheenginecompany.However,atlowpower settings, trace-species emissions were observed to be highly dependent on ambient conditions and engine temperature. Nonvolatile particles emitted by the engine exhibited a log-normal size distribution that peaked between 15 and 40 nm, depending on engine power. Samples collected 30 m downstream of the engine exit plane exhibited a prominent nucleation mode, indicating that secondary aerosols composed of sulfuric acid and low- volatility organic species formed rapidly within the plume as it expanded and cooled. Black carbon emissions were a minimum at approach and a maximum at climb and takeoff engine power settings. Black carbon dominated total mass emissions at high thrust, whereas volatile particles contributed an equal or perhaps greater fraction at low- to midpower settings. Although variations in fuel aromatic content had no discernible impact on particle emissions, volatile particle number and mass concentrations in aged exhaust plumes were highly sensitive to the fuel sulfur content.

Measurements in an Acoustically Driven Coaxial Jet under Sub-, Near-, and Supercritical Conditions

Abstract: An experimental investigation was conducted on a coaxial jet, similar to those used in cryogenic liquid rockets, under sub-, near-, and supercritical pressures, with the intent of gaining a better understanding of an aspect of combustion instability pertaining to interactions of an externally imposed acoustic field with the jet. Past research on this subject has shown both the relevance and importance of geometrical changes in an injector's exit-area and its nearby physical and fluid mechanical processes. Special attention is paid in collecting spatially resolved time averaged temperatures and documenting the aforementioned interactions at the exit of this injector. Short-duration and high-speed framing digital images provided information on the behavior of this jet under various conditions. Mean and root mean square values of the "dark-core" length fluctuations were measured from the acquired images via a computer-automated method, and results are discussed. There appears to be a good correlation between this length and the outer-to-inner-jet momentum ratio, but the form of this dependence was found to be different at subcritical pressures than the rest of the conditions. The root mean square values of the dark-core length fluctuations suggested possible explanation for the observed improvement in instability limit at increasingly higher outer-to-inner-jet velocity ratios.

Combustion instability with a single-element swirl injector

Abstract: A single injector element liquid rocket combustion experiment was designed and conducted to investigate the combustion dynamics of a gas-centered, liquid-swirled coaxial injector element. The oxidizer was a mixture of superheated water and oxygen, and kerosene was used as the fuel. The mean chamber pressure ranged from 2.14 to 2.38 MPa. The combustion chamber length was discretely varied between 25.4 and 88.9 cm to determine the dependence of combustion stability characteristics on resonant frequency and mode shape. Strong spontaneous instabilities were measured with peak-to-peak amplitudes of 0.69 to 1.38 MPa, and wave slopes on the order of 1000 MPa/s. The frequencies of the strongest instabilities ranged from 1184 to 1721 Hz. The most amplified modes ranged from the first longitudinal for the 38.1-cm chamber to the third longitudinal for the 88.9-cm chamber. One test, with a 25.4-cm chamber, was classically stable with pressure oscillation amplitudes less than 5% of the mean pressure. Resonant frequencies calculated with a model of the chamber acoustics compared well with measured values. For this injector, the data suggest that the observed stability behavior is a result of the combined effects of chamber mode shape and a driving combustion mechanism that limits the frequency range over which instability occurs.

Acoustic Simulation of Coaxial Hot Air Jets Using Cold Helium-Air Mixture Jets

Abstract: This work examines the ability of small-scale helium-air mixture coaxial jets to simulate the acoustics of large-scale hot air jets representing the exhaust of separate-flow turbofan engines. Experiments employed a one-eighth-scale model of a separate-flow nozzle used in hot tests at NASA John H. Glenn Research Center. Comparisons were conducted for two set points using the following methods: matching velocity and density, and matching velocity and Mach number. For both methods, the helium-air data compare well with the hot data in all measures of noise: spectral shapes, spectral levels, and overall sound pressure levels. The method of matching velocity and Mach number gives slightly better agreement in the spectral shapes at angles close to the jet axis and in the overall sound pressure levels. The overall agreement between the hot air and helium-air mixture data is within 1.2 dB.

Effect of pressure and propellant composition on graphite rocket nozzle erosion rate

Abstract: The objective of this work is to study the nozzle erosion rates at a broad range of pressures from 7 to 55 MPa with two baseline propellants: one is a nonmetallized propellant and the other is a metallized propellant, called propellants S and M, respectively. A comprehensive model for graphite nozzle erosion minimization and a numerical code has been advanced to predict the nozzle throat recession rates at high pressures. Four different kinetic schemes for heterogeneous graphite oxidation reactions were compared. The recession rate was found to increase almost linearly with pressure. The magnitudes of recession rates depend on the chemical kinetic scheme and the propellant composition. Contrary to popular belief, at lower pressures (P < 14 MPa), the heterogeneous kinetic rates showed a pronounced effect on the erosion rates, though at higher pressures, the nozzle throat erosion is mainly diffusion controlled. This observation stresses the importance of more accurate and definitive kinetic parameters for graphite oxidation reactions, especially at lower pressures. It was also observed that, besides H 2 O, the OH species affects the nozzle recession rate greatly. For the metallized propellant, the concentrations of major oxidizing species such as H 2 O, OH, and CO 2 are substantially reduced in comparison with the nonmetallized propellant, resulting in significant reduction of the erosion rates. A comparison of experimental data and predicted results from the graphite nozzle erosion minimization code shows excellent agreement especially for the nonmetallized propellant. To substantially reduce the throat recession rates at high pressures, it is suggested that the boundary-layer control at the throat region could be an effective method for future nozzle design considerations.

Development of scalable space-time averaged regression rate expressions for hybrid rockets

Abstract: DOI: 10.2514/1.19226 The fuel regression rate expressions reported in the hybrid literature often depend explicitly on the physical dimensions of the system such as the fuel port diameter. Typically, when these dimensional formulas are applied to systems with significantly different scales, they produce grossly inaccurate results. This paper addresses the development of scalable space–time averaged regression rate formulas for hybrid rockets. The derivation process hingesonthe assumption thatthe local instantaneous regression rateisa function of the local mass fluxandthe axial port distance in the power law format as predicted by the classical theory developed by Marxman. In this study, we have developed physics-based nondimensional formulas for the space–time averaged regression rate and used these expressions to develop a scalable regression rate law for a selected propellant combination, paraffin-based SP-1a/ GOX,fromlimitedmotordata.Initially,spaceandtimeaveragingaretreatedseparately,whichwerelatercombined to develop a technique that allows for the coupling between the spatial and time variations to predict the port diameter and mass flow rate profiles as functions of time. Finally, a comprehensive technique to estimate the systematic and random errors on the regression rate and mass flux data is also outlined.

Experimental Study of the Unstable-Unstarted Condition of a Hypersonic Inlet at Mach 6

Abstract: Wind-tunnel tests were conducted to better understand the flow pattern of a hypersonic inlet at an unstable-unstarted condition that is caused by the downstream mass flow choking. High-speed shadowgraph flow visualization and instantaneous pressure measurements are employed to reveal the oscillations of the shock system and the unsteady process of the duct flow. Results show that the shock system both inside and in front of the inlet duct oscillates intermittently when the inlet is unstarted. With the increase of the downstream throttling ratio, the fundamental frequency of the unsteady flow process goes up and the shock-on-lip time goes down, but the process whereby the shock system is disgorged and swallowed remains unchanged, lasting for 15 ms. When the throttling ratio is 89 and 91%, the fundamental frequency is 8 and 23 Hz, respectively. At the smaller throttling ratio, the aerodynamic oscillatory phenomenon of the inlet can be divided into three stages: namely, mass filling up, shock system disgorging and swallowing, and near-throat flow pattern establishing and backpressure propagating. At the larger throttling ratio, the first stage disappears and the third one shortens.

Design and Characterization of a Hypervelocity Expansion Tube Facility

Abstract: We report on the design and characterization of a 152 mm diameter expansion tube capable of accessing a range of high enthalpy test conditions with Mach numbers up to 7.1 for aerodynamic studies. Expansion tubes have the potential to offer a wide range of test flow conditions as gas acceleration is achieved through interaction with an unsteady expansion wave rather than expansion through a fixed area ratio nozzle. However, the range of test flow conditions is in practice limited by a number of considerations such as short test time and large amplitude flow disturbances. We present a generalized design strategy for small-scale expansion tubes. As a starting point, ideal gas dynamic calculations for optimal facility design to maximize test time at a given Mach number test condition are presented, together with a correction for the expansion head reflection through a non-simple region. A compilation of practical limitations that have been identified for expansion tube facilities such as diaphragm rupture and flow disturbance minimization is then used to map out a functional design parameter space. Experimentally, a range of test conditions have been verified through pitot pressure measurements and analysis of schlieren images of flow over simple geometries. To date there has been good agreement between theoretical and experimental results.

Atomization and Flames in LOX/H2- and LOx/CH4- Spray Combustion

Abstract: Hydrogen is a widely used rocket fuel and methane is particularly of interest in Europe as a promising substitute for H2. Experimental investigation of cryogenic reactive coaxial sprays with oxygen as an oxidizer and hydrogen and methane as fuels is conducted to prove whether concepts from LOX/H2 injector design can be transferred to LOX/CH4 injection. The liquid oxygen has been atomized in shear co-axial atomizers, the sprays and the flames have been investigated by visualization methods like shadowgraphy and imaging of the flame emission. LOX-sprays are characterized for both propellants by the intact core lengths and droplet numbers and the combustion is analyzed in terms of the flame anchoring mechanism and the flame spreading angle. The results for LOX/H2- and LOX/CH4-spray combustion are compared, also the influence of the injection conditions of the propellants on atomization and spray flame is discussed. Significant differences of the sprays and flames have been observed for the two propellant combinations at similar injection conditions as defined by Weber number and momentum flux ratio. The flame stabilization process has shown a strong influence on the atomization and flame characteristics.

Solid Rocket Motor Internal Ballistics Simulation Using Three-Dimensional Grain Burnback

Abstract: Internal ballistics simulations of solid rocket motors have been conducted with the propellant grain's 3-D burning surface geometry described by a new minimum distance function approach and the internal flowfield represented by 1-D, time-dependent, single-phase compressible flow equations. The combustion model includes erosive burning and unsteady, dynamic burning corresponding to transient energy storage in the heated surface layer of the propellant. The integrated internal ballistics code (Rocballist) is used to investigate the role of these two burning rate augmenting mechanisms in solid rocket motor performance. Two tactical motors are used as test cases. Results indicate that dynamic burning can be the dominant factor in producing a short-duration ignition pressure spike in low-L* motors, particularly if the L/D ratio is not too large and the port cross section is nonrestrictive (e.g., center perforated grain). However, when L/D is large and the port cross section is noncircular in the aft section (aft fins/slots), erosive burning can take over in dominating the burning rate to the extent that an otherwise progressive pressure-time trace becomes regressive/neutral. That is, erosive burning can effectively prolong the initial pressure spike in some star-aft motors. The results also show that with sufficiently accurate models of dynamic burning and erosive burning, it is reasonable to expect reliable internal ballistics predictions with suitable simplified flowfield models, thereby realizing significant reductions in computation time compared with 3-D, multiphase reacting flow simulations.

Experimental Investigation of a Pressure Swirl Atomizer Spray

Abstract: The fuel injector has an important role in the process for an efficient combustion because it increases the specific surface area of the fuel and it allows one to reach high rates of mixing and evaporation. This paper has focused on the behavior of kerosene Jet A-1 spray produced by commercial pressure swirl atomizers in terms of mean diameter distributions, velocity component profiles, and cone angle variations over time. The analysis has been carried out experimentally with the aid of a phase-Doppler anemometer system, which provides drop sizes and velocities. The fluctuating behavior of the spray has been analyzed also with a fast imaging technique by means of a high-speed charge-coupled device camera. The instantaneous spray cone angle, estimated with a digital image analysis, has shown that the average cone angle roughly equals the one calculated with phase-Doppler anemometer data, and that it is different from its nominal value. A preliminary investigation into the frequency domain has shown two oscillation modes at low frequency around 100 Hz and at high frequency around 1800 Hz. This issue is a key result because spray oscillations affect the flame stability and the resulting combustion efficiency because a change in the local air/fuel mixture ratio is induced.

Analytical Ion Thruster Discharge Performance Model

Abstract: *† ‡ A particle and energy balance model of the plasma discharge in magnetic ring-cusp ion thrusters has been developed. The model follows the original work of Brophy in the development of global 0-D discharge models that utilize conservation of particles into and out of the thruster and conservation of energy into the discharge and out of the plasma in the form of charged particles to the walls and beam and plasma radiation. The present model is significantly expanded over Brophy’s original work by closing the set of equations with selfconsistently calculations of the internal neutral pressure, electron temperature, primary electron density, electrostatic ion confinement (due to the ring-cusp fields), plasma potential, discharge stability, and time dependent behavior during recycling. The model only requires information on the thruster geometry, ion optics performance and electrical inputs such as discharge voltage and currents, etc. to produce accurate performance curves of discharge loss versus mass utilization efficiency. The model has been benchmarked against NSTAR and NEXIS thrusters, and successfully predicts the thruster discharge loss as a function of mass utilization efficiency for a variety of thrusters. The discharge performance model will be described and results showing ion thruster performance and stability presented.

Ignition and oxidation of lean CO/H2 fuel blends in air

Abstract: Ignition and oxidation characteristics of CO/H 2 fuel blends were studied using both experimental and computer simulation methods. Shock-tube experiments were conducted behind reflected shock waves at intermediate temperatures (890 < T < 1300 K) for three pressure regimes of approximately 1, 2.5, and 15 atm. Results of this study provide the first undiluted fuel-air ignition-delay-time experiments to cover such a wide range of CO/H 2 composition (5-80% H 2 ) over the stated temperature range. Emission in the form of chemiluminescence from the hydroxyl radical (OH*)A 2 Σ + → X 2 Π transition near 307 nm was used to monitor the reaction progress from which ignition delay times were determined. In addition to the experimental analysis, chemical kinetics calculations were completed to compare several chemical kinetics mechanisms with the new experimental results. The models were in excellent agreement with the shock-tube data, especially at higher temperatures and lower pressures, yet there were some differences between the models at the higher pressures and lowest temperatures, in some cases by as much as a factor of 5. Ignition-delay-time and reaction-rate sensitivity analyses were completed at higher and lower temperatures and higher and lower pressures to identify the key reactions responsible for ignition. The results of the sensitivity analysis indicate that the ignition-enhancing reaction H + O 2 = O + OH and hydrogen oxidation kinetics in general were most important, regardless of mixture composition, temperature, or pressure. However, lower-temperature, higher-pressure ignition-delay-time results indicate additional influence from HO 2 - and CO-containing reactions, particularly, the well-known H + O + M = HO 2 + M reaction and the CO + O + M = CO 2 + M and CO + HO 2 = CO 2 + OH reactions. Differences in the rates of the CO-related reactions are shown to be the cause of discrepancies among the various models at elevated pressures. Additional calculations were performed to show that the mixtures used are insensitive to small levels of water vapor, and the disagreement between experiment and model at the lowest temperatures and higher H 2 concentrations cannot be explained by possible impurities.

Regression rates study of mixed hybrid propellants

Abstract: The low regression rates of classic hybrid rocket fuels lead to large internal ports that limit potential applications. This experimental study investigated the increase in regression rate that results from adding a solid oxidizer and a catalyst to a hybrid fuel grain. The configuration is named a "mixed hybrid" hybrid to signify solid oxidizer and catalyst in the grain. A design of experiments approach guided fuel formulation to systematically control levels of ammonium perehlorate from 25% to 30%, ferric oxide from 0 to 5%, and hydroxyl-terminated polybutadiene from 70% to 75%. The 1.5-in. diam. port, 12-in. long center perforated grains were burned with gaseous oxygen at pressure levels from 150 to 550 psig and port flux levels from 0.1 to 0.4 lbm/s-in. 2 . The results show that the mixed hybrid propellants burn as a function of both pressure and mass flux. A grain formulation having 27.5 % ammonium perchlorate and 2.5% ferric oxide provided the maximum burning rate augmentation (447%) among the formulations tested.

Kinetic Analysis of Thermite Reactions in Al-MoO3 Nanocomposites

Abstract: Reactions in energetic Al-MoO 3 nanocomposites prepared by arrested reactive milling were investigated by scanning calorimetry and heated filament ignition experiments. The calorimetry data were processed to obtain kinetic parameters describing the reaction between Al and MoO 3 . The reaction was treated as a combination of four subreactions, which were described by a combination of a diffusion-controlled reaction model and first-order reactions. The activation energies determined in this study allowed the comparison to reference values for the decomposition of MOO 3 and the diffusion of oxygen through an Al 2 O 3 product layer. The kinetic model was extrapolated to high heating rates in the 10 3 -10 6 K/s range and compared to ignition data. It was concluded that ignition of Al-MoO 3 nanocomposites prepared by arrested reactive milling is primarily controlled by oxygen diffusion in Al 2 O 3 .

Combustion and Ignition of Thermally Cracked Kerosene in Supersonic Model Combustors

Abstract: A series of experiments were conducted to characterize the self-ignition and combustion of thermally cracked kerosene in both a Mach 2.5 model combustor with a combustor entrance height of 51 mm and a Mach 3.0 model combustor with an entrance height of 70 mm. A unique kerosene heating and delivery system was developed, which can prepare heated kerosene up to 950 K at a pressure of 5.5 MPa with negligible fuel coking. The extent of China no. 3 kerosene conversion under supercritical conditions was measured using a specially designed system. The compositions of gaseous products as a result of thermal cracking were analyzed using gas chromatography. The mass flow rates of cracked kerosene were also calibrated and measured using sonic nozzles. With the injection of thermally cracked kerosene, the ability to achieve enhanced combustion performance was demonstrated under a variety of airflow and fuel conditions. Furthermore, self-ignition tests of cracked kerosene in a Mach 2.5 model combustor over a range of fuel injection conditions and with the help of different amounts of pilot hydrogen were conducted and discussed.

Internal Flowfield Investigation of a Hypersonic Inlet at Mach 6 with Bleed

Abstract: To study the influence of viscous effects on the performance and, especially the internal flowfield, of scramjet engines, an experimental study has been carried out on a fixed geometry scramjet inlet operating at Mach 6. The tests have been performed using a hypersonic blowdown wind-tunnel facility. To reduce the risk of inlet unstart, a passive boundary-layer bleed has been integrated at the throat and successfully tested. The passive bleed reduces the lip shock-induced separation bubble on the ramp significantly. The obtained experimental results are discussed and compared with computational fluid dynamics calculations. To investigate the compression behavior, different backpressures have been applied and the effects on the internal flow structure are analyzed by means of schlieren change in pitot pressure and calculated Mach number profile base don the geometrical shape of the internal flowpath. The heat transfer coefrficient to the inner sidewall of the isolator is calculated using the time-dependent surface temperature, measured with an infrared system. The variation of the Stanton number based on different flow phenomena could be resolved successfully.

Experimental Characterization of a Micro-Hall Thruster

Abstract: The operation and performance of a micro-Hall thruster are characterized. The thruster is coaxial in design, with a 0.5 mm channel width and 4 mm outer diameter. The magnetic circuit includes a samarium cobalt permanent magnet generating approximately 0.7 T at the exit plane and 1 T inside the channel. Operation with a commercial hollow cathode neutralizer is achieved in the 10-40 W power range with an anode flow rate of 0.12-0.20 mg/s of xenon. The measured thrust is in the range of 0.6-1.6 mN for an anode flow rate of 0.17-0.20 mg/s and an applied voltage of 110-275 V. Anode thrust efficiency and specific impulse are in the range of 10-15% and 300-850 s, respectively, for the same conditions. Relatively broad ion energy distributions and large beam divergence are observed from an analysis of the plume using a retarding potential analyzer and ion current probe. The discharge exhibits the characteristic Hall thruster "breathing mode" instability in the 35-70 kHz frequency range.

Turbulence Measurements and Analysis in a Multistage Axial Turbine

Abstract: This paper presents turbulence measurements and detailed flow analysis in an axial turbine stage. Fast response aerodynamic probes were used to resolve aperiodic fluctuations along the three directions. Assuming incompressible flow, the effective turbulence level and Reynolds stress are retrieved by evaluating the stochastic velocity component out of the measured time-resolved pressure and flow angle fluctuations along the streamwise, radial, and circumferential direction. A comparison between turbulence intensity and measured total pressure shows that flow structures with higher turbulence level are identified in the region of loss cores at the exit of the second stator passage. Turbulence intensity is evaluated under isotropic and nonisotropic assumption in order to quantify the departure from isotropic conditions. The measurements show that locally the streamwise fluctuating component can be twice bigger than the radial and tangential component. The current analysis shows that multisensor fast response aerodynamic probes can be used to provide information about the mean turbulence levels in the flow and the Reynolds stress tensor, in addition to the measurements of unsteady total pressure loss.

Combustion Behaviors Resulting from Bimodal Aluminum Size Distributions in Thermites

Abstract: Studies that replace a portion of the micron-size aluminum (Al) with nano-Al particles in an energetic formulation demonstrate significant performance enhancement Little is known, however, about the critical level of nano-sized fuel particles needed to enhance the performance of the energetic composite Ignition sensitivity and combustion velocity experiments were performed using a thermite composed of Al and molybdenum trioxide (MoO 3 ) Both loose powders and compressed pellets were examined A bimodal Al particle size distribution was prepared using 4 or 20-μm-diam Al fuel particles that were replaced in 10% increments by 80-nm-diam Al particles until the fuel was 100% nano-AI Results show that with only 20% nano-Al content, the mixtures showed reduced ignition delay times by up to 2 orders of magnitude The combustion velocity was shown to dramatically increase as more nano-Al particles replace micron-Al particles within the mixture This increasing trend was attributed to incomplete reactions of the micron-Al particles or significantly slower reactions such that the micron-Al particles promote cooling or quenching of the reaction

Effect of ambient gas density on spray characteristics of swirling liquid sheets

Abstract: The spray and breakup characteristics of a swirling liquid sheet were investigated by measuring the spray angle and breakup length as the axial Weber number We l was increased up to 1554 and the ambient gas pressure up to 4.0 MPa. As the We l and ambient gas density p increased, the disturbances on the annular liquid sheet surface were amplified by the increase of the aerodynamic forces, and thus the liquid sheet disintegrated from the injector exit. The measured spray angles according to the ambient gas density differed before and after the sheet broke up. Before the liquid sheet broke up, the spray angle was almost constant; however, once the liquid sheet started to break up, the spray angle decreased. As the ambient gas density and We l increased, the increasing aerodynamic force caused the breakup length to decrease. Finally, the measured breakup lengths according to the ambient gas density and We l were compared with the results of the linear instability theory. Considering the attenuation of sheet thickness in the linear instability theory, the corrected breakup length relation agreed well with our experimental results.

Solid Propellant Grain Design and Burnback Simulation Using a Minimum Distance Function

Abstract: DOI: 10.2514/1.22937 A fast computational method for simulating the evolution of the burning surface of a complex, three-dimensional solid rocket motor propellant grain has been developed using a signed minimum distance function. The minimum distance function is calculated using stereo-lithography surface information from a computer-aided-design file and propellant surface burnback is simulated by manipulation of the initial minimum distance function. Variable time steppingandmultiplespatialgridsfurtherreducecomputationtimerequirements. Resultsindicatethatthismethod gives adequate accuracy with acceptable computation time for time scales of the full motor burn. The resulting code (Rocgrain) allows for motor grain design by user-friendly commercial computer-aided-design programs and for coupling with internal flow codes. This enables a single geometric tool to be used for describing the propellant grain geometry for both grain design and internal flowfield analysis. The Rocgrain code can be coupled with a variety of flowfield codes ranging in complexity from simple zero-dimensional to more sophisticated computational fluid dynamics analysis (e.g., nonlinear acoustic instability).

Physical Characterization of Aerosol Emissions from a Commercial Gas Turbine Engine

Abstract: This paper discusses the results of the Aircraft Particle Emissions Experiment Project for the physical characterization of total (nonvolatile plus volatile) aerosol emissions (emission factors, hydration properties, and distribution shape parameters) by extractive sampling from an on-wing CFM56-2C 1 engine. Samples were extracted at the engine exit plane (1 m) as well as locations 10 and 30 m downstream. Three different fuels were used in this study: base fuel, high-sulfur fuel, and high-aromatic fuel. For the 1 and 10-m probe locations, strong and sometimes nonlinear dependencies were observed on fuel flow rate and no statistically significant dependencies were observed for fuel composition. At 30 m, the onset of gas-to-particle conversion was apparent for low- to medium-fuel flow rates. The soluble mass fraction was found to increase with distance from the engine exit plane and with increasing fuel aromatic and sulfur content. An intercomparison of gas and particle sampling trains showed that gas-to-particle conversion is a serious sample train artifact for gas sampling trains in which dilution cannot be achieved at the probe tip.