Showing papers in "Journal of Propulsion and Power in 2009"

Chemical Properties of Aircraft Engine Particulate Exhaust Emissions

Abstract: The chemical properties of the particulate exhaust emissions from an in-use commercial aircraft engine were characterized in April 2004 as part of the Aircraft Particle Emissions Experiment. The test aircraft was the NASA DC-8 equipped with CFM56-2-C1 engines and the test matrix included 11 different engine throttle levels, three fuel compositions, and three sampling distances. The variations in particle emissions number, size, mass, and chemical composition were measured using a suite of instruments, including an aerosol mass spectrometer. The particle emissions were characterized by a trimodal size distribution. The largest mode was dominated by ambient accumulation mode particles mixed into the plume. The middle mode consisted of carbon soot with sulfate and organic coatings. The smallest mode was completely volatile and consisted of sulfate and organic components. The soot emission indices increased with power from 2―120 mg/kg fuel. The semivolatile components increased with distance and decreased with power from 33―5 mg/kg fuel. The sulfate emissions increased with distance and fuel sulfur content. The emissions under low power were dominated by organics, and the high-power conditions were dominated by soot. The CFM56 engine was less efficient at the low thrust levels typically used on the ground at an airport.

Acoustics, Vortex Shedding, and Low-Frequency Dynamics Interaction in an Unstable Hybrid Rocket

Abstract: This paper deals with an experimental investigation into the stability behavior of a hybrid rocket where gaseous oxygen is fed with either an axial conical subsonic nozzle or a radial injector. The influence of the oxidizer-injection configurations on the motor stability is thoroughly examined. These distinct oxidizer-injection techniques allowed unveiling key and so far unreported features of the hybrid rocket combustion stability, especially emphasizing the role of vortex shedding which occurs in both the pre- and postcombustion chamber. Axial and radial injectors caused completely stable and unstable combustor operations, respectively, and this fact has been attributed to the fluid dynamics and unsteady heat release at the entrance of the fuel grain port. In particular, the unstable combustion in the radial-flow injector motor was dominated by low-frequency pressure oscillations, around 10-20 Hz. These low-frequency pressure oscillations were always accompanied by longitudinal acoustic modes. In some cases, the pressure oscillations abruptly increased, reaching peak-to-peak amplitude close to 70% of the mean chamber pressure, which is somewhat unusual for hybrid engines. Vortex shedding in the aft-mixing chamber is considered as the main driving mechanism of this latter behavior.

Effect of Bulk Density on Reaction Propagation in Nanothermites and Micron Thermites

Abstract: The thermite reaction of nanoscale aluminum and molybdenum trioxide particles has revealed a paradoxical relationship between Al particle size and mixture bulk density. Specifically, with micron-scale Al particles, the thermite demonstrates an expected growth in flame speed with increased density, but nanoscale-Al-particle mixtures exhibit an opposing trend. This paper presents new experimental measurements of the thermal properties of this thermite as a function of Al particle size and applies a new oxidation mechanism in an effort to explain the paradoxical results between Al particle size and mixture bulk density. Results show that the nanocomposite's behavior is consistent with a new melt-dispersion oxidation mechanism and convective mode of flame propagation. Compaction-induced damage of the oxide shell and distortion of the shape of spherical particles, as well as reduced free space around Al nanoparticles suppress the melt-dispersion mechanism and reduce flame speed. An additional mode of energy transfer is proposed that is associated with molten Al clusters from the melt-dispersion mechanism that advance faster than the flame velocity. Micron-scale particle reactions may be governed by diffusion such that increased bulk density coincides with increased thermal properties and increased flame speeds.

Review of Recent Developments in Wave Rotor Combustion Technology

Abstract: For some decades, efforts have been made to exploit nonsteady combustion and gas dynamic phenomenon. The theoretical potential of nonsteady-flow machines has led to the investigation of various oscillatory flow devices such as pulse detonation engines, wave rotors, pulse jets, and nonsteady ejectors. This paper aims to provide a progress review of past and current research in developing a particular combustion concept: the wave rotor combustor. This pressure-gain combustor appears to have considerable potential to enhance the performance and operating characteristics of gas turbine and jet engines. After attempts in the mid-twentieth century were thwarted by mechanical problems and technical challenges identified herein, recent successes in Switzerland and efforts in the United States benefited from design expertise developed with pressure-exchange wave rotors. The history, potential benefits, past setbacks, and existing challenges and obstacles in developing these nonsteady combustors are reviewed. This review focuses on recent efforts that seek to improve the performance and costs of future propulsion and power-generation systems.

Characterization of Nanothermite Material for Solid-Fuel Microthruster Applications

Abstract: Nanothermite composites containing metallic fuel and inorganic oxidizer have unique combustion properties that make them potentially useful for microthruster applications. The thrust-generating characteristics of copper oxide/aluminum nanothermites have been investigated. The mixture was tested in various quantities (9―38 mg) by pressing the material over a range of densities. The testing was done in two different types of thrust motors: one with no nozzle and one with a convergent―divergent nozzle. As the packing density was varied, it was found that the material exhibited two distinct impulse characteristics. At low packing pressure, the combustion was in the fast regime, and the resulting thrust forces were ∼75 N with a duration of less than 50 μs full width at half-maximum. At high density, the combustion was relatively slow and the thrust forces were 3―5 N with a duration 1.5―3 ms. In both regimes, the specific impulse generated by the material was 20―25 s. The specific impulse and short thrust duration created by this unique nanothermite material makes it promising for micropropulsion applications, in which space is limited.

Chemical Erosion of Refractory-Metal Nozzle Inserts in Solid-Propellant Rocket Motors

Abstract: An integrated theoretical/numerical framework is established and validated to study the chemical erosion of refractory-metal (tungsten, rhenium, and molybdenum) nozzle inserts in solid-rocket-motor environments, with a primary focus on tungsten. The formulation takes into account multicomponent thermofluid dynamics in the gas phase, heterogeneous reactions at the surface, energy transport in the solid phase, and nozzle material properties. Typical combustion species of nonmetallized ammonium-perchlorate/hydroxyl-terminated-polybutadiene propellants at practical motor operating conditions are considered. The erosion rates calculated by employing three different sets of chemical kinetics data available in the literature for the tungsten-steam reaction have been compared. The effect of considering either of two different tungsten oxides, WO2 or WO3, as the final product of surface reactions is also investigated. The predicted erosion rates compare well with experimental data. The oxidizing species of H2O proved more detrimental than CO2 in dictating the tungsten nozzle erosion. The material recessionrateiscontrolledbyheterogeneouschemicalkineticsbecausethediffusionlimitisnotreached.Theerosion rate increases with increasing chamber pressure, mainly due to higher convective heat transfer and enhanced heterogeneous surface reactions. The tungsten nozzle erodes much more slowly than graphite, but at a rate comparable with that of rhenium. The molybdenum nozzle exhibits the least erosion for flame temperatures lower than 2860 K. Its low melting temperature (2896 K), however, restricts applications for propellants with high flame temperatures.

Methodology and Historical Perspective of a Hall Thruster Efficiency Analysis

Abstract: to ionization processes and losses that manifest as Joule heating, and contains no information about the vector properties of the jet. Propellant efficiency incorporates losses from dispersion in the jet composition and is unity for 100% ionization to a single ion species. The effect of neutrals on dispersion of the jet velocity distribution function in propellant efficiency is introduced in the neutral-gain utilization. The beam efficiency accounts for divergence of the jet and is ideal when the ion velocity vectors are parallel to the thrust axis. Plume divergence is defined as a momentum-weighted term, and the approximation as a charge-weighted term is characterized. The efficiency architecture is derived from first principles and is applicable to all propulsion employing electrostatic acceleration, including Hall thrusters and ion thrusters. Distinctions and similarities to several past methodologies are discussed, including past ion thruster analyses, early Russian performance studies, and contemporary architectures. To illustratethepotentialforenhancedunderstandingoflossmechanismsandionizationprocesseswithanarrayoffarfield plume diagnostics, a case study is presented of low-discharge voltage operation from a 6 kW laboratory Hall thruster.

In-Stream Hypermixer Fueling Pylons in Supersonic Flow

Abstract: This paper presents results from both computational fluid dynamic and wind-tunnel experiments of in-stream fueling pylons injecting air, ethylene, and methane gas into Mach number 2.0 cold airflow. Three fuel-injection pylons studied include a basic pylon, a ramp pylon, and an alternating-wedge pylon. The latter two pylons introduce streamwise vorticity into the flow to increase mixing action. The computational fluid dynamic solution was accomplished using the commercial code FLUENT®. Three wind-tunnel experimental techniques were used: aerothermal probing, Raman spectroscopy, and nitric-oxide planar laser-induced fluorescence. Four measures reported include streamwise vorticity, total-pressure-loss, mixing efficiency, and flammable plume extent. The ramp and alternating-wedge pylons show decisive increases in mixing capability compared with the basic pylon for a finite distance downstream of the injector. The alternating-wedge pylon exhibits a measurable increase in total pressure loss compared with the basic pylon, and the ramp pylon exhibits a negligible increase in total pressure loss compared with the basic pylon. For comparison, the downstream mixing effectiveness of the three pylons is compared with the downstream mixing effectiveness of a transverse circular wall injector studied in past research. In addition, a qualitative comparison between the computational fluid dynamic and wind-tunnel experimental results is made.

Reaction Mechanism of Aluminum-Particle-Air Detonation

Abstract: Both in-tube and unconfined experimental evidence showed strong dependence of micrometric aluminum-air detonability on initial pressure and highly nonlinear behavior of abrupt deflagration-to-detonation transition, thus indicating dependence of the aluminum reaction mechanism of the detonation waves on chemical kinetics. On the other hand, the observed aluminum―air detonation manifested itself in a weak transverse wave structure, as revealed by the small-amplitude oscillation that rapidly degenerates behind the shock front in the pressure histories. This suggests a functional dependence that is weaker than the nonlinear Arrhenius kinetic behavior for the later aluminum combustion. Hence, a surface kinetic oxidation and diffusion hybrid reaction model with a degree of condensed detonation products was suggested, and the unsteady two-phase fluid dynamics modeling showed the success of the hybrid reaction model, capable of capturing both the kinetics-limited transient processes of detonation initiation, abrupt deflagration-to-detonation transition and detonation instability, and the diffusion-limited combustion of aluminum in the long reaction zone, supporting the weak transverse wave structure.

Nanoaluminum as a Solid Propellant Fuel

Abstract: Experimental studies on the burning of nanoaluminum-based solid rocket propellants are carried out. Data on the properties of condensed combustion products, mechanisms of their formation, and burning-rate law are obtained. Based on these data, a physical picture is developed of the considered burning-propellant classes. Mathematical modeling of burning nanoaluminum in composite solid rocket propellants is carried out. The influence of nanoaluminum on ignition temperature of the metal fuel and burning-rate law is shown. The results of this study allow carrying out the analysis and selection of good-quality propellants using nanoaluminum.

Swept-Leading-Edge Pylon Effects on a Scramjet Pylon-Cavity Flameholder Flowfield

Abstract: This study explores the effect of adding a pylon to the leading edge of a cavity flameholder in a scramjet combustor. Data were obtained through a combination of wind-tunnel experimentation and steady-state computational fluid dynamics. Wind-tunnel data were collected using surface pressure taps, static and total probe data, shadowgraph flow visualization, and particle image velocimetry. Computational fluid dynamics models were solved using the commercial FLUENT software. The addition of an intrusive device to the otherwise low-drag cavity flamebolder offers a potential means of improving combustor performance by enabling combustion products to propagate into the main combustor flow via the low-pressure region behind the pylon. This study characterized the flowfield effects of adding the pylon as well as the effect of changing Reynolds numbers over the range of approximately 33 x 10 6 to 55 × 10 6 m ―1 at a Mach number of 2. The addition of the pylon resulted in approximately 3 times the mass flow passing through the cavity compared with the cavity with no pylon installed. Reynolds number effects were weak. The addition of the pylon led to the cavity fluid traveling up to the top of the pylon wake and significantly increasing the exposure and exchange of cavity fluid with the main combustor flow.

Shock-Tunnel Experiments with a Mach 12 Rectangular-to-Elliptical Shape-Transition Scramjet at Offdesign Conditions

Abstract: A shock-tunnel investigation of a scramjet with a rectangular-to-elliptical shape-transition inlet and an elliptical combustor has been conducted at conditions simulating flight at Mach 8.7. The inlet was designed using a quasistreamline-tracing method to have a design point of Mach 12.0 and operation down to Mach 6.0. The elliptical combustor began with a rearward-facing step around its perimeter and was followed by a constant-area and diverging section. The flowpath was completed by a short thrust nozzle. Gaseous hydrogen fuel was injected either through multiple portholes on the intake, a series of 48 portholes on the rearward-facing step at the combustor entrance, or a combination of the two. All fueling configurations resulted in a positive thrust coefficient at equivalence ratios above 0.3 without the use of ignition aids. Fuel injection in the intake produced robust combustion and good internal thrust levels, but led to inlet unstart at fuel equivalence ratios above 0.61. Stable mixing-limited combustion was observed for fuel injection at the step at all fuel equivalence ratios up to 1.23. Combined intake and step injection was observed to have the best performance. These experimental results demonstrate that rectangular-to-elliptical shape-transition scramjets designed for access-to-space applications can operate efficiently at conditions below the design Mach number.

Quasi-One-Dimensional Model of Hydrogen-Fueled Scramjet Combustors

Abstract: A computationally efficient, quasi-one-dimensional, supersonic combustion ramjet (scramjet) propulsion model has been produced for use in hypersonic system design studies. The model solves a series of ordinary differential equations using a fourth-order Runge–Kutta method to describe the gas dynamics within the scramjet duct. Additional models for skin friction and wall heat transfer are also included. The equations are derived assuming an open thermodynamic system with equilibrium or simplified-chemistry combustion models. The combustion is also assumed to be mixing-limited rather than kinetically limited. This assumption allows simplification of the modeling and is justified when the model is compared against experimental results. Three test cases are used to validate the performance of the scramjet propulsion model: 1) a reflected-shock-tunnel hydrogen-fueled scramjet experiment, 2) a continuous-flow hydrogen-fueled scramjet ground test, and 3) a segment of the HyShot II flight test. The results show that the model simulates scramjet propulsion with a reasonable degree of accuracy.

Thermal Cracking and Heat Sink Capacity of Aviation Kerosene Under Supercritical Conditions

Abstract: pressurerangeof3–4.5MPa,andaresidencetimerangeof0.6–3s.Thechemicalheatsinkwasdeterminedthrougha control volume analysis of the fuel flow. Compositions of the cracked gaseous and liquid products were analyzed via gas chromatography. Based on the results of fuel conversion, the temperature range for the active cracking was observedtobeapproximately800–1000K,beyondwhichthecrackingapproachescompletion.Itwasalsofoundthat the variation of the chemical heat sink with temperature can be nonmonotonic and a maximum endothermicity was seen to occur in the temperature range of 900–960 K, depending on the residence time. For the current operation conditions, the maximum chemical heat sink reached 0:5 MJ=kgat a fuel conversion of 45%. Composition analysis ofthegaseousproductindicatedthatthesaturatedhydrocarbonssuchasmethanebecamedominantastemperature

Investigation of Self-Sustaining Waves in Metastable Systems: Deflagration-to-Detonation Transition

Abstract: Self-sustaining waves can propagate in metastable media; energy needed to support such waves is released by the wave itself. The examples are waves of combustion and waves of boiling in overheated liquids. As a rule, two regimes of propagation exist: subsonic and supersonic. The difference is based on the different mechanisms of medium activation. Processes of transition between those regimes were less studied up to now, in comparison with pure subsonic or supersonic modes. Knowing mechanisms of controlling detonation initiation is important to work out effective preventive measures, such as suppressing deflagration-to-detonation transition in the case of combustible mixture ignition, and mitigation of a detonation wave in case it is already developed. On the other hand, the advantages of burning fuel in a detonation regime in comparison with slow burning at constant pressure attract increasing attention to pulse detonation burning chambers and to their possible application to new generation engines. The deflagration-to-detonation transition can be a principal stage of the work cycle in a pulse detonation engine, and the knowledge of details ofthis process and means of control can significantly decrease the predetonation distance and optimize the device. This work contains a review of the results obtained in theoretical and experimental investigations of deflagration-to-detonation transition processes in gases. Influence of internal geometry and flow turbulization on the detonation onset is considered; the influence of temperature and fuel concentration in the unburned mixture is discussed. Transitional processes of overheated liquid boiling up are also analyzed.

Production, Characterization, and Combustion of Nanoaluminum in Composite Solid Propellants

Abstract: non- and microaluminized propellants are washed out with the addition of nano-Al, but the burning rates of nanoaluminized propellants register low pressure exponents in the elevated pressure range. The results suggest that the heat feedback from the diffusion-limited combustion of nano-Al particles near the propellant burning surface controls the propellant burning rate when sufficiently large parts of the burning surface are made up of the nanoaluminized fine ammonium-perchlorate/binder matrix.

Mixing and combustion characteristics of kerosene In a model supersonic combustor

Abstract: In this numerical study, supersonic combustion of kerosene in three model combustor configurations is investigated. To this end, 3-D, compressible, turbulent, nonreacting, and reacting flow calculations with a single step chemistry model have been carried out. For the nonreacting flow calculations, the droplet diameter distribution at different axial locations, variation of the Sauter mean diameter, and the mixing efficiency for three injection pressures are presented and discussed. In addition, the effect of turbulent dispersion on the mixing efficiency is studied using a stochastic model in conjunction with the two-equation shear stress transport k-ω turbulence model. For the reacting flow calculations, contours of heat release and axial velocity at several axial locations are used to identify regions of heat release inside the combustor. Combustion efficiency predicted by the present results is compared with earlier predictions for all the combustor models. Furthermore, the predicted variation of static pressure along the combustor top wall is compared with experimental data reported in the literature. Calculations show that the penetration and spreading of the fuel increases with an increase in the injection pressure. Predicted values of the combustion efficiency are more realistic when the spray model is used for modelling the injection of the fuel. The importance of the mixing process, especially for a liquid fuel such as kerosene, on the prediction of heat release is discussed in detail.

Scaling and Evaluation of Pt/Al2O3 Catalytic Reactor for Hydrogen Peroxide Monopropellant Thruster

Abstract: A scaling methodology of hydrogen peroxide monopropellant thruster is described. As the decomposition process of the hydrogen peroxide on the surface of catalyst bed is extremely complex, empirical method was taken for design purposes. A small-scale thruster was fabricated and important design parameters, including temperature at different locations of the catalyst bed, were measured. Based on the measurement, the catalyst bed size as a function of the propellant flow rate was estimated. Using the scaling methodology, a catalyst bed configuration for a thruster capable of delivering 50 N was estimated. The thruster built on this design produced 42 N at sea level and specific impulse of 123 s.

Performance Validation of a Single-Tube Pulse Detonation Rocket System

Abstract: A pulse detonation engine (PDE) can be operated even if there are no compression mechanisms such as compressors or pistons, and a rocket engine with an extremely low combustor fill pressure (pulse detonation rocket, PDR) thus becomes possible. In this research, we made a model PDR system with increased specific impulse by partial fill. The performance predicted by this model was then confirmed experimentally. The thrust can be calculated by using the simplified PDE model of Endo et al. and the partial filling effect models of Sato et al. The mass flow rate of the propellant supplied from the pressurized cylinders is considered in this calculation. As a result, the thrust performance can be determined by the kind of propellant, the initial conditions of the gas in the cylinders, the supply-valve orifice and PDE-tube volume, and the operation frequencies. We fabricated a pulse detonation rocket (PDR) named “TODOROKI” and verified the thrust calculation model via a horizontal sliding test. We confirmed that the stability of the PDE operation depends on the ratio between the purge-gas thickness and the tube diameter. The thrust predicted by the model was identical to experimental results within 4%.

Multiple-Point Adaptive Performance Simulation Tuned to Aeroengine Test-Bed Data

Abstract: Adaptive simulation technology enables the calibration of a performance simulation code to a given in-service gas turbine and provides correct prediction of its performance. This is a fundamental prerequisite for reliable gas-path diagnostics and performance health monitoring. In this paper, a new offdesign performance adaption algorithm is introduced. Cranfield University's consolidated engine performance simulation code PYTHIA is enhanced with the capability of offdesign performance adaptation to model available field data. The software minimizes, via a genetic algorithm, an objective function that measures the error between an initial engine model output and the real engine data by varying some characteristics' scaling factors. In this study, a multiple-point adaptation procedure was applied to a two-shaft aeroengine. This generated an optimized engine model that minimized its deviations from a set of test-bed data. The adapted model was then tested against different real data, resulting in an average error, over 8 measured parameters, of less than 0.35%.

Aluminum Combustion Driven Instabilities in Solid Rocket Motors

Abstract: This work describes an unreported instability found out by numerical simulations on a solid rocket motor. A simple motor with a cylindrical port is considered and predicted to be stable by single-phase computational fluid dynamics computations. When aluminum combustion is modeled, the motor experiences a strong instability on the first longitudinal acoustic mode. However, no vortex shedding is observed, meaning that it is a genuine combustion instability. A detailed numerical study leads to the conclusion that the instability is thermoacoustic and results from a coupling between chamber acoustics and aluminum combustion heat release. A parametric study points out the importance of aluminum distributed combustion, particularly the thickness of the combustion zone and the aluminum heat of reaction. Theoretical assessment of this instability is also obtained by revisiting the acoustic balance theory. An additional thermoacoustic stability integral is derived and appears to be a driving term. This term also helps to shed light on this instability and explains some of the computational fluid dynamics results. In particular, the underlying mechanism is found to be primarily caused by the acoustic boundary layer which creates high acoustic velocities that enhance heat release from burning aluminum particles.

Experiments on Electrically Controlled Flameholding on a Plane Wall in a Supersonic Airflow

Abstract: We describe experiments on gaseous fuel ignition and flameholding controlled by an electrical discharge in high-speed airflow The geometrical configuration does not include any mechanical or physical flameholder The fuel is nonpremixed and injected directly into the air crossflow from the combustor bottom wall A multi-electrode, nonuniform transversal electrical discharge is excited, also on the bottom wall, between flush-mounted electrodes The initial gas temperature is lower than the value for autoignition of hydrogen and ethylene Results are presented for a wide range of fuel mass flow rate and discharge power deposited into the flow This coupling between the discharge and the flow presents a new type of flameholder over a plane wall for a high-speed combustor

Experimental Investigation of Thermoacoustic Coupling for Low-Swirl Lean Premixed Flames

Abstract: Combustion instabilities are frequently encountered in gas turbine engines. The main physical mechanism that leads to combustion instability is thermoacoustic coupling, where energy released as part of the combustion process is transferred into combustor acoustics, which can then affect the combustion heat release in a feedback process. This paper presents an experimental study ofthis coupling for a lean premixed methane―air low-swirl flame. Planar laser-induced fluorescence of the hydroxyl radical was employed to measure the behavior of this flame in an acoustically driven environment and the Rayleigh index was used as an indicator of the thermoacoustic coupling. The acoustic excitation ranged from 13 to 270 Hz. Experiments show that the coupling occurred within certain frequency subranges depending on the Reynolds and Strouhal number. The acoustic excitation can couple with the shear-layer vortices, which in turn wrinkle the flame with a periodicity similar to the acoustic driving. This effect, through the changes in local flame surface density, gives rise to intense toroidal structures in the Rayleigh index field.

Effect of Recess in High-Pressure Liquid Oxygen/Methane Coaxial Injection and Combustion

Abstract: The effect of a recessed liquid oxygen tube in shear coaxial injection has been investigated experimentally using an optically accessible subscale rocket combustor operated at pressures between 40 and 60 bar. Different single-shear coaxial injectors have been used to inject liquid oxygen and methane at relevant operating conditions covering sub-, near-, and supercritical pressures with respect to the critical point of oxygen. Liquid oxygen was injected at 120 K and the injection temperature of gaseous methane was about 275 K. Detection of spontaneous OH and CH chemiluminescene has been performed to characterize the flame-anchoring zone near the liquid oxygen post tip. In addition, the influence of the injector geometry on the combustion roughness and stability has been investigated during steady-state operating points. An increased flame expansion was observed with a recessed injector element. At low momentum flux ratio, the pressure drop accross the injector increases with a recessed liquid oxygen tube compared with a flush tube. Furthermore, a recessed liquid oxygen tube led to be a smoother combustion in general; however, this configuration also led to additional resonant frequencies in the chamber acoustics.

Thrust Measurement of a Multicycle Partially Filled Pulse Detonation Rocket Engine

Abstract: In the present research, we experimentally verified the partial-fill effect in a multicycle pulse detonation rocket engine. The intermittent thrust of a pulse detonation rocket engine was measured by using a spring-damper mechanism that smoothed this intermittent thrust in the time direction. The intermittent mass flow rates were assessed by gas cylinder pressure or mass difference measurement. The maximum specific impulse was 305 ± 9 s at an ethylene and oxygen propellant fill fraction of 0.130 ± 0.004. When the fill fraction was greater than 0.130, the specific impulse was increased as the partial-fill fraction was decreased. When the fill fraction was less than 0.130, the specific impulse was sharply decreased as the partial-fill fraction was decreased. This decrease was due to diffusion between propellant and purge gases and the short length of the transition from deflagration to detonation. The multicycle pulse detonation rocket engine had a partial-fill effect that may have been mainly due to the suctioned air and was consistent with the single-cycle partial-fill model of Endo et al. [Endo, T., Yatsufusa, T., Taki, S., Matsuo, A., Inaba, K., and Kasahara, J., "Homogeneous-Dilution Model of Partially-Fueled Simplified Pulse Detonation Engines," Journal of Propulsion and Power, Vol. 235, 2007, pp. 1033-1041.].

Validating a Plasma Momentum Flux Sensor to an Inverted Pendulum Thrust Stand

Abstract: The accuracy of a plasma impact force sensor was compared with that of the more commonly used inverted pendulum thrust stand using a 5 kW Xe Hall effect thruster. An improved plasma momentum flux sensor was designed and constructed based on a previous design. Real-time force measurements were made with both the plasma momentum flux sensor and the inverted pendulum thrust stand. The plasma momentum flux sensor measured the force exerted onto it by the Hall effect thruster exhaust plume with a resolution of 0.1 mN and an average discrepancy of 2 % compared with thrust stand measurements. Experiments were completed using a 9 m by 6 m cylindrical vacuum chamber. The total force from the Hall effect thruster was modulated from 34 to 356 mN by varying both the anode voltage, from 150 to 500 V, and the neutral Xe gas flow rate, from 5 to 15 mg/s.

Flame Stabilization in High-Pressure Liquid Oxygen/Methane Rocket Engine Combustion

Abstract: Flame stabilization in the near-injector region of a subscale liquid propellant rocket combustion chamber has been investigated using optical diagnostics. Several different single shear coaxial injectors have been used to inject liquid oxygen (LOX) and methane at relevant operating conditions. Three main operating points cover the range of sub-, near-, and superciritical conditions with respect to the thermodynamic cirtical point of oxygen. Whereas liquid temperatures of about 270 K. Similar to previous investigations performed with LOX/H2 combustion, spontaneous OH chemiluminescene has been detected to characterize the flame anchoring zone near the LOX post tip. Theoretical considerations have indicated that the binary mixture of oxygen and methane shows and entirely different behaviour compared with the oxygen/hydrogen system. This is believed to have an influence on the spray evolution and mixing characteristics at supercritical conditions. The experimental investigation includes both the ignition transient a well as the steady-state operating points. It has been detected that the LOX/CH4 flame shows very similar characteristics in comparison with LOX/H2 flame at similar operating conditions. Critical flame stabilization at the startup transient has been found during all hot runs. At steady-state conditions, the influence of the injection parameters on the flame shape is comparable to previous LOX/Hydrogen investigations.

Design of Quiet Propeller for an Electric Mini Unmanned Air Vehicle

Abstract: Designing a quiet and efficient propeller is a demanding task because these two goals often lead to contradictory design trends. The task becomes even more complicated when additional constraints are introduced. This paper presents a multidisciplinary design optimization process for designing a quiet propeller under various constraints. The acoustic model of the propeller is presented in detail, including its validation. The new process is used to design a quiet propeller for an electric mini unmanned air vehicle. This design is subjected to power, structural, and side constraints. The design variables include blade geometry, blade cone angle, propeller radius, number of blades, and rotational speed. It is shown that the motor characteristics have an important influence on the optimization, and it is therefore important to take them into account during the design process. The influences of the various parameters on the design and propeller's characteristics are presented and discussed in detail.

Performance Characteristics of an Annular Conical Aerospike Nozzle with Freestream Effect

Abstract: An experimental investigation has been carried out to study the performance and base pressure characteristics of a Mach 2.0 annular conical aerospike nozzle with and without freestream flow. The effect of cowl length, plug length, and plug contour variation on the nozzle performance and base pressure characteristics is studied. It is observed that the overexpansion shock from the internal nozzle, overexpansion shock on the spike surface, and the expansion fan from the cowl lip of the internal nozzle dominate the overall flowfield development. The presence of freestream flow reduces the nozzle performance by approximately 4% relative to static conditions. Base pressure characteristics are observed to be strongly influenced by the movement of these shocks on the plug surface, and their subsequent interaction with the inner shear layer controls the base-wake closure. Relative to the conical plug conflguration, the contoured plug shows considerably enhanced base pressure characteristics. Real-time pressure measurements on the spike reveal highly unsteady flow in the intermittent region of separation.