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Showing papers in "Journal of Propulsion and Power in 2012"


Journal ArticleDOI
Yen Yu1, James C. Sisco, Stanford Rosen1, Ajay Madhav1, William E. Anderson1 
TL;DR: In this paper, an experimental study of spontaneous longitudinal high-frequency combustion instabilities in a high-pressure model rocket combustor is described, where a traversing axial oxidizer inlet is used to vary the coupled resonances between the variable-length oxidizer tube and the fixed-length combustion chamber, thereby changing the temporal and spatial relationship between the fluid mechanical resonances in the oxidizer tubes and the local pressure oscillations at the head end of the combustor.
Abstract: An experimental study of spontaneous longitudinal high-frequency combustion instabilities in a high-pressure model rocket combustor is described. The experiment is designed to efficiently obtain stability data under variable resonance conditions in a single test. A traversing choked axial oxidizer inlet is used to vary the coupled resonances between the variable-length oxidizer tube and the fixed-length combustion chamber, thereby changing the temporal and spatial relationship between the fluid mechanical resonances in the oxidizer tube and the local pressure oscillations at the head end of the combustor. Marked transitions between pressure-fluctuation amplitudes and spectral contents are observed as the injector tube length changes. Stable-to-unstable and unstable-to-stable transitions are observed. Depending on the injector tube length, the first dominant injector-coupled resonance ranged from1200 to 1500Hz.The high-frequency pressure oscillation amplitudes varied from less than 10%ofmean pressure to greater than 30% of chamber pressure. Spectrograms are used to assess the global frequency contributions, and power spectral densities are used as a quantitative measure of the stability characteristics and to identify the most unstable geometry. The combustor was also operated in fixed-tube-length mode to measure linear growth rates and compare limit cycle amplitudes. The experimental results provide a basis for comparison with stability models, and their use in conjunction with reliable computational fluid dynamics models can potentially provide detailed insight into the physics of self-excitation and limit cycle behavior in a practical system.

119 citations


Journal ArticleDOI
TL;DR: In this paper, the performance of a nominal 200 W Hall effect thruster fueled by iodine vapor was evaluated, and the system included a laboratory propellant feed system, a flight-model Hall thruster, and breadboard power processing unit.
Abstract: The performance of a nominal 200 W Hall effect thruster fueled by iodine vapor was evaluated. The system included a laboratory propellant feed system, a flight-model Hall thruster, and breadboard power processing unit. Operation of the Hall thruster with iodine vapor was stable on both short and long time scales, enabling measurements of thruster performance across a broad range of operation conditions. Performance was found to be comparable with xenon. At 200 W, thrust is 13 mN, anode specific impulse is 1500 s, and anode efficiency is 48%. Plume-current measurements indicate a profile typical of xenon Hall thrusters, while E B probe measurements indicate the presence of ionic dimers.

97 citations


Journal ArticleDOI
TL;DR: A class of arc filament actuators called localized arc filament plasma actuators for high-speed and high-Reynoldsnumber flow and acoustic control has been developed at the Gas Dynamics and Turbulence Laboratory as mentioned in this paper.
Abstract: A class of plasma actuators called localized arc filament plasma actuators for high-speed and high-Reynoldsnumber flow and acoustic control has been developed at the Gas Dynamics and Turbulence Laboratory. Over the past several years, these high-bandwidth (0 to 200 kHz) and individually controlled actuators have been used successfully to excite the jet shear layer, jet column, and azimuthal instabilities in high subsonic and supersonic jets. The focus of this paper is to provide detailed information and sample results highlighting the capabilities and potential of the actuators and the control technique formixing enhancement, noise mitigation, and flow and acoustic diagnostics. The jet, using three different nozzles, is operated over a large range of jet Mach numbers (0.9 to 1.65), stagnation temperature ratios (up to 2.5), and Reynolds numbers (0:2 10 to 1:65 10). Over this space of operating conditions, the jet is found to respond to control with a large range of forcing Strouhal numbers and azimuthalmodes. The results reveal that the jet flowfield and acoustic farfield can be dramatically altered, providing a powerful control tool in these practical high-speed and high-Reynolds-number jets.

84 citations


Journal ArticleDOI
TL;DR: In this article, an operating point approach is used to create amapping of the coupling effects between the isolator geometry, inlet flow conditions, and fuel injector behavior, and the resulting isolator/injector coupling map provides a description of the response of the isolateator to particular injector performance.
Abstract: Isolator–combustor interactions are measured in a direct-connect dual-mode ramjet-scramjet experiment. An operating point approach is used to create amapping of the coupling effects between the isolator geometry, inlet flow conditions, and fuel injector behavior. The resulting isolator/injector coupling map provides a description of the response of the isolator to particular injector performance and the effective blockage it induces on the isolator flow. Existing models and correlations predicting the pressure rise across a pseudoshock and its resultant length were evaluated through comparisonwithmeasurementsmade in a heated-flow isolator duct that is coupled to a hydrogenair combustor. The observation of a normal-to-oblique shock-train transitionmechanism has led to the development of a revised shock-train operating regime description that takes into account the impact of Mach number and maximum pressure recovery on the shock configurations present in the isolator.

78 citations


Journal ArticleDOI
TL;DR: In this article, the authors show that the total thruster efficiency is independent of the cathode flow over the range from 5 to 10% of the propellant injected into the thruster body through the anode.
Abstract: The cathode coupling voltage in Hall thrusters, which is the voltage difference between the cathode and the thruster beam plasma potential, is considered an indicator of the ease with which electrons flow from cathode to anode.Historically, the coupling voltage has beenminimizedby increasing the amount of propellant injected through the hollow cathode due to early observations that thismaximizes the discharge (or anode) efficiency.However, recent experiments described here show that the total thruster efficiency is independent of the cathode flow over the range from 5 to 10%of the propellant injected into the thruster body through the anode. For this reason, cathode flow rates can be reduced closer to the classic plume mode limit characteristic of the hollow cathode design without impacting the total thruster efficiency. Such reductions in cathode flow rate can significantly extend the cathode life, especially for higher-power Hall thrusters with larger discharge currents, where the normal Hall thruster cathode flow split will significantly exceed the optimum level for cathode operation and life.

69 citations


Journal ArticleDOI
TL;DR: In this article, the start flow of a hypersonic inlet was experimentally studied at a freestream Mach number of 4.5-6.0 with the aid of high-speed schlieren and time-accurate pressure measurements.
Abstract: Unstart flows of a hypersonic inlet were experimentally studied at freestream Mach number of 4.5–6.0. With the aid of high-speed schlieren and time-accurate pressuremeasurements, the unsteady flow processes of the entire inlet, including the shock systemmotions and the surface pressure fluctuations, were recorded and discussed. The started flowfield analysis was conducted first, and then the unstarted flowfield was analyzed by using the unsteady pressure signals and schlieren pictures. Results indicate that two novel oscillatory patterns were observed in comparison with the past reported inlet buzz patterns. One is a mixed oscillatory pattern that mixes the “big buzz”with “little buzz,” and the other is a nonoscillatory violent pattern. These novel findings on oscillatory patterns of hypersonic inlet can provide more insight on inlet buzz mechanism, prediction, and control.

60 citations


Journal ArticleDOI
TL;DR: In this paper, a new wind-tunnel flowfield has been proposed that captures much of the key shock boundary-layer interaction physics of supersonic external compression inlets.
Abstract: T HE interaction of a shock wavewith a turbulent boundary layer constitutes a fundamental problem of high-speed fluid mechanics. A detailed survey of past work on high-speed interactions has been carried out by Settles and Dolling [1] and Smits and Dussauge [2]. The shock interaction problem is particularly germane to the design of supersonic inlets. In such supersonic inlets, deceleration of the flow is achieved through a succession of oblique shock waves followed by a terminal normal shock. Boundary layers form on the inlet surfaces and interact with the shock system, giving rise to various shock/boundary-layer interactions (SBLIs). Each interaction of oblique/normal shock waves with the boundary layer causes stagnation pressure losses and downstream spatial distortions seen by the engine. An inlet must be carefully designed to minimize these losses and distortions during the compression process since they affect overall propulsion performance. In mixed-compression inlets, shock-induced separation can lead to engine unstart, which requires that the entire propulsion system undergo a restart sequence during flight. In external compression inlets, specifically axisymmetric configurations, a thick hubside boundary layer increases blockage and can decrease compressor performance. Thus, successfully controlling SBLIs has the potential to significantly improve supersonic inlet performance. As will be discussed in the following, various techniques of flow control for SBLIs have been proposed. However, it is often difficult to interpret the results because the flowfield may be too specific to an individual inlet configuration or too basic such that a relationship to inlet performance is not clear. To address this issue, a newwind-tunnel flowfield has been proposed [3] that captures much of the key shock boundary-layer interaction physics of supersonic external compression inlets. Thisflowfieldwill be used to study the novel flow control methods introduced herein. The conventionalflow control technique for SBLI conditions in an engine inlet employs a bleed of the boundary layer [4,5]. This bleed Presented as Paper 2010-4464 at the 40th AIAA Fluid Dynamics Conference and Exhibit, Chicago, IL, 28 June–1 July 2010; received 31 January 2011; revision received 26 July 2011; accepted for publication 11 August 2011. Copyright©2011 by theAmerican Institute ofAeronautics and Astronautics, Inc. All rights reserved. Copies of this paper may be made for personal or internal use, on condition that the copier pay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 0748-4658/12 and $10.00 in correspondence with the CCC. Ph.D. Candidate, Aerospace Engineering, 104 S. Wright Street. Member AIAA. Professor of Aerodynamics, Department of Engineering, Trumpington Street. Associate Fellow AIAA. Professor, Mechanical and Aerospace Engineering, 122 Engineer’s Way. Associate Fellow AIAA. JOURNAL OF PROPULSION AND POWER Vol. 28, No. 1, January–February 2012

51 citations


Journal ArticleDOI
TL;DR: In this article, an experimental investigation of boron combustion in an ethanol spray flame was carried out using x-ray diffraction, porosimetry, thermogravimetric analysis, and scanning electron microscopy.
Abstract: Biofuels such as ethanol have lower energy density than conventional petroleum-based fuels; therefore, enhancing their energy density by addition of higher-energy-density components is an attractive option. Boron is considered to be a good choice as a fuel additive because it has almost the highest volumetric heating value among potentially suitable additives. The present study deals with an experimental investigation of boron combustion in an ethanol spray flame. Measurements have been made of the emission of the intermediate suboxide BO2 using spectroscopy and chemiluminescence imaging with interference filters. The size, surface structure, and chemical composition of the injected boron nanopowders were characterized by x-ray diffraction, porosimetry, thermogravimetric analysis, and scanning electron microscopy. The postburn particles were also characterized by x-ray diffraction and thermogravimetric analysis. The effects of boron nanoparticles on the hydrocarbon combustion in terms of heat releasewere determined usingCHchemiluminescence and temperaturemeasurements. The chemiluminescence and spectroscopic signatures indicate that boron underwent combustion and simultaneously increased the CH-radical emission and flame temperatures. The x-ray diffraction and thermogravimetric analysis measurements of the particles collected at the exhaust indicate that complete conversion of the boron was achieved. The effect of particle size on the combustion of the boron nanoparticles was examined, comparing small commercial particles to sintered particles. A demonstrable size effect was observed with the larger particles, which exhibited their initial BO2 signatures at distances further above the dump plane, indicating greater ignition delay times.

50 citations


Journal ArticleDOI
TL;DR: In this paper, the transition from the weak mode to the intensive mode occurred at Φ ~ 0.4, accompanied by a sudden increase in thrust at the mode transition, where the combustion region expanded toward the cowl with boundary-layer separation.
Abstract: A sidewall compression scramjet engine operated in two combustion modes under Mach 6 flight condition, weak- and intensive-combustion modes. The weak mode occurred below the overall fuel equivalence ratio (Φ) of around 0.4. Transition from the weak mode to the intensive mode occurred at Φ ~ 0.4, accompanied by a sudden increase in thrust. Mechanisms of the transition were numerically investigated in this study. Our simulations captured the sudden increase in thrust at the mode transition. In the weak mode, combustion occurred in only a region near the topwall where an igniter was installed. The combustion region expanded toward the cowl with boundary-layer separation at the mode transition. Our simulations demonstrated that low ignition capability resulted in the weak mode. We demonstrated that the presence of additional igniters on the sidewalls improved the ignition capability and achieved the intensive mode in the entire Φ range.

48 citations


Journal ArticleDOI
TL;DR: In this paper, the authors examined the near field of a Mach 1.3 jet subjected to eight plasma actuators mounted on the periphery of the nozzle exit and compared several quantities with experiments, including mean and fluctuating streamwise velocities, as well as visualizations of coherent features.
Abstract: Simulations are employed to examine the near field of aMach 1.3 jet subjected to eight plasma actuators mounted on the periphery of the nozzle exit. The simulations are validated by comparing several quantities with experiments, including mean and fluctuating streamwise velocities, as well as visualizations of coherent features. The generation, evolution, and breakdown of coherent structures with first (m 1), second (m 2), and third (m 3) azimuthal mode excitation are described and set in the context of previous simulations with axisymmetric and mixed modes at themost amplified jet columnmode (Strouhal number 0.3). Hair-pin-like vortices, observed in instantaneous, as well as phase-averaged data, are noted as key building blocks in the evolution process. The tips/heads arise in the outer region of the jet shear layer and legs extend forward, slightly inclined in the direction of the jet axis where the velocity is higher. The interaction of these structures through self and mutual induction with those generated by adjacent actuators, or later by the same actuator, yields a rich variety of features depending on excitationmode. In the second part of the paper, the effect of Strouhal number on the near field is explored with m 0 and m 1 excitation. The results show that, at higher frequencies, large structure formation is inhibited and peak sound pressure level amplitudes diminish in a manner consistent with experimental results.

48 citations



Journal ArticleDOI
TL;DR: In this article, phase-locked intensified charge coupled device imaging of CH chemiluminescence shows that during combustion instability there aremainly variations in the chemilumininescence intensity rather than in the spatial distribution of heat release.
Abstract: The present paper reports flame structure and combustion instability characteristics in a turbulent liquid-fueled swirl-stabilized lean direct fuel injection combustor. Because of the complexities in droplet size and distribution, evaporation, droplets–vapor–air entrainment and mixing, and droplets–flame–turbulence interactions, the flame structure is remarkably different from lean premixed gas-fueled combustion. With the preheat temperature above 423K, heat release is completed within a compact doughnut ring, about 12mmdownstream of the dump plane; with decreasing equivalence ratios at the approach of lean blowout, the doughnut ring shrinks in diameter and gradually converges to the axis; after that, the heat release zone becomes elongated and tilted off the axis. Chemiluminescence only provides qualitative information of the heat release rate, suggesting the limitations of global chemiluminescence measurements for spray combustion. Phase-locked intensified charge coupled device imaging of CH chemiluminescence shows that during combustion instability there aremainly variations in the chemiluminescence intensity rather than in the spatial distribution of heat release. Depending on the working conditions, the one-wave mode, the half-wave mode, or both of the combustion chambers can be excited. Although the combustion chamber is highly blocked at both ends, no pressure antinodes are found at the chamber exit and the chamber inlet. Simultaneous excitation of the one-wave mode and the half-wave mode is intrinsically unsteady and unstable, which can be attributed to the nonlinear response of the reacting swirling shear layer to acoustic oscillations. Both the amplitude and the frequency of thermoacoustic oscillations are constantly time-varying. With decreasing oscillation intensity, the limit-cycle oscillator becomes increasingly vulnerable to external disturbances, as indicated by a larger neighborhood of the state trajectory.

Journal ArticleDOI
TL;DR: In this article, a study was conducted to predict carbon-carbon nozzle erosion behavior in full-scale solid-rocket motors for wide variations of motor operating conditions, including variable chamber pressure over the burning time and the effect of nozzle shape change on the erosion rate.
Abstract: The erosion of nozzle protection materials during solid-rocket-motor burning needs to be accounted for to get reliable performance predictions, especially for long-durationfirings.A study is conducted topredict carbon–carbon nozzle erosion behavior in full-scale solid-rocket motors for wide variations of motor operating conditions. The numerical model considers the solution of Reynolds-averaged Navier–Stokes equations in the nozzle, heterogeneous chemical reactions at the nozzle surface, ablation species injection in the boundary layer, variable multicomponent transport and thermodynamic properties, and heat conduction in the nozzlematerial. Two different ablationmodels are considered: a diffusion-limited approach and a finite-rate approach. The numerical model is used to study the erosion of carbon–carbon nozzle inserts for the secondand third-stage solid-rocket motors of the European Vega launcher. The effect of variable chamber pressure over the burning time and the effect of nozzle shape change on the erosion rate are taken into account in the numerical analysis. The obtained results show a very good agreement with the measured final eroded profile along the entire carbon–carbon nozzle throat insert for both motors. The shapechange effect is shown to be an important factor that has to be taken into account to get a goodprediction of the throat erosion for long-duration firings.

Journal ArticleDOI
TL;DR: In this article, the authors investigate the operation condition of a fluidic thrust vector using injection of the control flow tangential to the main jet direction; co-flow injection is used to analyze the dynamic characteristics of fluidic control of jet vectoring up-and downward from the nozzle axis, so that the response time of jet deflection to control flow injection and the pressure dispersion on the nozzle wall were investigated.
Abstract: The purpose of this research is to investigate the operation condition of fluidic thrust vector using injection of the control flow tangential to the main jet direction; co-flow injection. The physical model of concern includes a chamber and a supersonic nozzle for supersonic main jet injection, and two chambers with slots for control flow injection. Steadystate numerical and experimental studies were conducted to investigate operating parameters; detailed flow structures, jet deflection angles, and shock effects were observed near the nozzle exit. An unsteady numerical calculation was conducted to analyze the dynamic characteristics of fluidic control of jet vectoring up- and downward from the nozzle axis, so that the response time of jet deflection to control flow injection and the pressure dispersion on the nozzle wall were investigated. Internal nozzle performance was predicted for total pressure range of the jet from 300 kPa to 1000 kPa to the control flow pressure from 120 to 200 kPa. To take into account the important features of high-speed flows, including shock-boundary layer interactions, a low Reynolds number k-e turbulence model with compressible-dissipation and pressure-dilatation effects was applied.


Journal ArticleDOI
TL;DR: In this article, a full-annulus detached-eddy simulation (DES) is conducted to investigate stall inception for the axial transonic rotor, NASA rotor 67, which is observed that the rotating stall is initiated by the local spike flow disturbance, which quickly induces the rotor to stall roughly over two rotor revolutions.
Abstract: Full-annulus detached-eddy simulation (DES) is conducted in this paper to investigate stall inception for the axial transonic rotor, NASA rotor 67. A low-diffusion E-CUSP scheme with a third-orderMUSCL reconstruction is used to discretize the inviscid fluxes, and second-order central differencing is used for the viscous terms. An implicit line Gauss–Seidel iteration with dual time-stepping method and second-order temporal accuracy is employed for time integration of the spatially filtered unsteady Navier–Stokes equations in a rotating frame. It is observed that the rotating stall is initiated by the local spike flow disturbance, which quickly induces the rotor to stall roughly over two rotor revolutions. The stall cell covering more than six blade tip passages propagates at 48% of rotor speed in the counter-rotor rotation direction. The process of rotating stall is captured by the full-annulus DES, which indicates that the blockage created by the low-energy vortical flow structure pushes the tip leakage flow to the adjacent blade behind the detached sonic boundary. The unsteady Reynolds-averaged Navier–Stokes (URANS) simulation is also performed for comparison. DES predicts the stall inception roughly one rotor revolution earlier than the URANS model. Overall, DES predicts the stall inception more realistically.

Journal ArticleDOI
TL;DR: In this article, shape optimization has been performed with surrogate-assisted evolutionary algorithms to maximize the thrust generated by an axisymmetric scramjet nozzle configuration, including the base flow and external surface for cruise conditions at Mach 8 at two altitudes with and without fuel.
Abstract: Scramjet propulsion is a promising hypersonic airbreathing technology that offers the potential for efficient and flexible access to space and high-speed atmospheric transport. Robust nozzle design over a range of operating conditions is of critical importance for successful scramjet operation. In this paper, shape optimization has been performed with surrogate-assisted evolutionary algorithms to maximize the thrust generated by an axisymmetric scramjet nozzle configuration, including the base flow and external surface for cruise conditions at Mach 8 at two altitudes with and without fuel. The optimization results have been examined in a coupled numerical/analytical approach in order to identify the key design factors and investigate the effects of design parameters. It has been found that the optimum nozzle geometries are characterized by bell-type shapes for the fuel-on conditions, whereas the optima for the fuel-off case feature nearly conical shapes. Their robustness in thrust production has been demonstrated by cross- referencing the optimum geometries at off-design altitudes. The nozzle length and radius have been found to be the most influential parameters in all considered conditions, with their optimum values determined based on the balance between inviscid and viscous force components, whereas the other parameters have minor impact on the total axial force. Copyright

Journal ArticleDOI
TL;DR: The T6 ion engine is a 22-cm diameter, 4.5kW Kaufman-type ion thruster produced by QinetiQ, Ltd., and is baselined for the European Space Agency BepiColombo mission to Mercury and is being qualified under ESA sponsorship for the extended range AlphaBus communications satellite platform.
Abstract: *† * ‡ § ** The T6 ion engine is a 22-cm diameter, 4.5-kW Kaufman-type ion thruster produced by QinetiQ, Ltd., and is baselined for the European Space Agency BepiColombo mission to Mercury and is being qualified under ESA sponsorship for the extended range AlphaBus communications satellite platform. The heritage of the T6 includes the T5 ion thruster now successfully operating on the ESA GOCE spacecraft. As a part of the T6 development program, an engineering model thruster was subjected to a suite of performance tests and plume diagnostics at the Jet Propulsion Laboratory. The engine was mounted on a thrust stand and operated over its nominal throttle range of 2.5 to 4.5 kW. In addition to the typical electrical and flow measurements, an EB mass analyzer, scanning Faraday probe, thrust vector probe, and several near-field probes were utilized. Thrust, beam divergence, double ion content, and thrust vector movement were all measured at four separate throttle points. The engine performance agreed well with published data on this thruster. At full power the T6 produced 143 mN of thrust at a specific impulse of 4120 seconds and an efficiency of 64%; optimization of the neutralizer for lower flow rates increased the specific impulse to 4300 seconds and the efficiency to nearly 66%. Measured beam divergence was less than, and double ion content was greater than, the ring-cusp-design NSTAR thruster that has flown on NASA missions. The measured thrust vector offset depended slightly on throttle level and was found to increase with time as the thruster approached thermal equilibrium.

Journal ArticleDOI
TL;DR: In this paper, a method based on Riemann interactions is proposed for the analysis of two different nozzle geometries and supersonic flow, which is ideal for conceptual design, control design, or control evaluation studies.
Abstract: interaction between operating condition and plume shape complicates the analysis of such nozzles compared to traditional bell nozzles. A method that is based on Riemann interactions is proposed for the analysis of two such nozzle geometries. The method assumes two-dimensional geometries and supersonic flow. Unlike the method of characteristics, this method accounts explicitly for the presence of oblique shocks and curved shear layers. Comparisons to both experiment and computational fluid dynamics are shown. The solution method requires no grid generation and typically runs in less than a minute on a single desktop computer, which is ideal for conceptual design, control design, or control evaluation studies. It includes high-temperature gas modeling and finite-rate chemistry. Nomenclature A = area c = specific heat Ex = momentum conservation error H = height or length scale M = Mach number nexp = number of discrete waves in expansion nsp = number of species p = pressure r = length of characteristic R = normalized gas constant T = temperature u = magnitude of flow velocity W = molecular weight x, y = spatial coordinates Y = mass fraction = angle between wave and upstream flow = ratio of specific heats = deflection angle across a wave = angle of deflection caused by boundary layer = flowpath angle = momentum thickness " = ratio of static pressures = Mach angle = Prandtl-Meyer angle = streamwise coordinate = density = angle between wave and x-axis ˙ ! = molar rate of production


Journal ArticleDOI
TL;DR: In this paper, high-fidelity simulations with a validated methodology are employed to explore the physical processes associated with different injection strategies on supersonic combustion on a single-cavity flameholder.
Abstract: High-fidelity simulations with a validated methodology are employed to explore the physical processes associated with different injection strategies on supersonic combustion. The configurations consider a commonly employed open single-cavityflameholder. The effects of different injector locations and injection angles are examined under the constraint that the total fuel mass flow rate is the same. The numerical approach solves the full three-dimensional Navier–Stokes equations, supplemented with a two-equation k-! turbulence closure. The specific injection locations include 10 different arrangements that examine fuel injection upstream of the cavity, on the backward step, on the cavity bottom wall, and on the downstream ramp. The angles of the fuel port injection slots include combinations of parallel and 27 and 90 deg to the airflow inside the cavity. One case with a closed cavity is also examined for comparison. The simulations are employed to characterize the performancewith qualitative andquantitativemixing metrics. Detailed analysis of the results reveals both expected and unexpected findings. As anticipated, the closed cavity performs poorly relative to its open counterpart. However, an injection strategy that enhances the natural circulation pattern of the cavity is found to be superior to one that opposes it. Another counterintuitive finding is that although direct injection of the fuel upstream of the cavity into themain stream results in deeper penetration, the fuel from different injectors remains distinct, with relatively small spanwise mixing. On the other hand, injection on the floor of the cavity results in more diffused fuel distribution.

Journal ArticleDOI
TL;DR: In this article, the authors proposed a shock-unsteadiness model to account for the effect of unsteady shock motion in a steady-mean flow and showed that the model improved flow separation and reattachment in hypersonic flows.
Abstract: In hypersonic flows, the interaction of a shock wave with a turbulent boundary layer can result in flow separation and high aerothermal loads. In this paper, cone–flare configurationswith different flare angles and freestreamMach numbers are simulated using Reynolds-averaged Navier–Stokes method, and results are compared with experimental data. The standard Spalart–Allmaras and k-! turbulence models do not predict flow separation at the cone–flare junction, and therefore yield a large deviation from the surface pressure measurements. Sinha et al. (“Modeling Shock-Unsteadiness in Shock/Turbulence Interaction,” Physics of Fluids, Vol. 15, No. 8, 2003, pp. 2290– 2297) proposed a shock-unsteadiness model to account for the effect of unsteady shockmotion in a steadymean flow. The shock-unsteadiness correction damps turbulence amplification at the shock and results in significant improvement in predictingflow separation and reattachment. The flow topology in the interaction region, in terms of the pattern of shocks and expansion waves, is predicted correctly by the modified turbulence models. The resulting surface pressure distribution matches experimental data well.

Journal ArticleDOI
TL;DR: In this paper, a small-scale aerospike thruster with a truncated spike was used to demonstrate side force ampli cation factors up to 1.4 and side force bursts up to 55 s.
Abstract: Results from computational and coldow experiments on uidic thrust vectoring of a small-scale aerospike thruster are presented. Thrust vectoring is produced by injection of a secondary uid into the primary ow eld normal to the nozzle axis. The experimental aerospike nozzle was truncated at 57% of its full theoretical length. For these tests, carbon dioxide is the working uid. Injection points near the end of the truncated spike produced the highest force ampli cation factors. Explanations are given for this phenomenon. For secondary injection near the end of the aerospike, side force ampli cation factors up to 1.4 and side force speci c impulses up to 55 s were demonstrated. By comparison, the main ow speci c impulse averaged approximately 38s. Secondary side-injection pulses were observed to crisply reproduce side forces with a high degree of delity. Side force levels approach 2.7% of the total thrust level at maximum e ciency. Higher side forces of 4.7% axial thrust were also achieved at reduced e ciency. The side force ampli cation factors were independent of operating nozzle pressure ratio for the range of chamber pressures evaluated in this test series.


Journal ArticleDOI
TL;DR: In this article, a simple model of a Hall-effect thruster in which the propellant is an ambient air is presented, and the required lengths of the thruster chamber, themagnetic fields, the thrust, and other parameters of an ideal air-breathing Hall effect thruster are calculated as a function of the flying altitude of the vehicle.
Abstract: The principle idea of using airbreathing electrical propulsion for a vehicle flying at orbital speed on the edge of Earth’s atmosphere is examined. In this paper, a simple model of a Hall-effect thruster in which the propellant is an ambient air is presented. A new mode of the airbreathing thruster operation is presented in which incoming air is fully ionizedwithout preliminary compression. The required lengths of the thruster chamber, themagneticfields, the thrust, and other parameters of an ideal airbreathing Hall-effect thruster are calculated as a function of the flying altitude of the vehicle. For instance, in the case of the 95-kmorbit, the lengths of thruster chamber of 47.8 cm, theHall thruster gap of 2.3 cm, and the applied voltage of 3 kV, themodel gives the thrust of about 9.1 N and the power range of about 700–800 kW depending of the strength of a magnetic field. The thruster might be powered using electromagnetic beams generated at ground power stations or crafts flying at higher orbital altitudes where the drag force is negligible. The estimates presented show that the beam receiver size will not exceed the size of the vehicle and should not generate a drag force larger than the calculated thrust.

Journal ArticleDOI
TL;DR: This paper argues that the issue of accuracy of the experimental measurements be addressed by cross-facility and cross-disciplinary examination of modern data sets along with increased reporting of internal quality checks in particle image velocimetry analysis.
Abstract: Engineers charged with making jet aircraft quieter have long dreamed of being able to see exactly how turbulent eddies produce sound, and this dream is now coming true with the advent of large-eddy simulation. Two obvious challenges remain: validating the large-eddy-simulation codes at the resolution required to see the fluid–acoustic coupling, and the interpretation of the massive data sets that are produced. This paper addresses the former, the use of advanced experimental techniques such as particle image velocimetry and Raman and Rayleigh scattering, to validate the computer codes and procedures used to create large-eddy-simulation solutions. This paper argues that the issue of accuracy of the experimental measurements be addressed by cross-facility and cross-disciplinary examination of modern data sets along with increased reporting of internal quality checks in particle image velocimetry analysis. Furthermore, it argues that the appropriate validation metrics for aeroacoustic applications are increasingly complicated statistics that have been shown in aeroacoustic theory to be critical to flow-generated sound, such as two-point space-time velocity correlations. A brief review of data sources available is presented along with examples illustrating cross-facility and internal quality checks required of the data before they should be accepted for validation of large-eddy simulation.

Journal ArticleDOI
TL;DR: Based on Lefebvre's correlation, a new physical model is established and a flame volume concept is proposed according to the experimental observations as discussed by the authors. But, it is argued that the influences of the variations of combustor's configurations upstream of dilution holes on lean blowout cannot be embodied in thismodel.
Abstract: Lean blowout limits play a critical role in the operational envelope of aircraft. Semiempirical correlations, a convenient and fast methodology to estimate the lean blowout limits for aircraft engine combustors, are extensively employed in the preliminary combustion design stage. Among the correlations, Lefebvre’s model is widely used. However, it is argued that the influences of the variations of combustor’s configurations upstreamof dilution holes on lean blowout cannot be embodied in thismodel. Based on Lefebvre’s correlation, a new physical model is established and a flame volume concept is proposed according to the experimental observations. Then, an improved correlation (i.e., flame volume lean blowout model) is derived to consider the effects of the variations of dome geometry and primary zone configurations. The flame volume lean blowoutmodel is verified bymany fuel-lean visual experiments. In the experiments, three kinds of swirl-stabilized assemblies and two different primary hole arrangements have been employed. It is concluded that the flame volume lean blowout model shows better agreement with the corresponding experimental values of different designs than the Lefebvre model. The prediction uncertainties of the two models are about 15 and 45% in the present combustion chamber configurations, respectively.

Journal ArticleDOI
TL;DR: In this paper, a methodology based on the degree of disequilibrium of chemical reactions is proposed to select kinetic constraints in the method of rate-controlled constrained equilibrium (RCCE).
Abstract: Amethodology based on the degree of disequilibriumof chemical reactions is proposed to select kinetic constraints in the method of rate-controlled constrained equilibrium (RCCE). Predictions of chemical relaxation of the combustion products within a supersonic nozzle are also made under the constraints identified using the proposed methodology. The quasi-one-dimensional steady Euler equations are employed. The RCCE method provides a general framework, which enables studying shifting equilibrium, frozen equilibrium, as well as nonequilibrium chemical kinetics under different sets of constraints, either a single or a linear combination of species. It is shown that the previously identified generalized constraints of the total number of moles, free valence, and free oxygen are implied as a consequence of this methodology for the kinetics considered and for the relaxation studied. The kinetic scheme involves anH=O reactingmixture with 24 reactions and eight species. The constraints identified areH H2, HO2 H2O2, andO OH H2O, which result in less than 0.06% local errors in temperature compared with the detailed kinetic model calculations throughout the nozzle.

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TL;DR: This paper brings together many important prediction methods and makes a systematic study of their accuracy and applicability and proposes a new method, called NOx: generic, which compares well with the most dependable method, viz., the P3 T3 method, but unlike the P2 T2 method the present formulation does not require any proprietary information.
Abstract: Protecting the environment from the consequences of human activity has become a major challenge and goal in recent years, and therefore such considerations have become a critical component of engineering design and operation as well as of the formulation of policies and legislation. Emission from aircraft engines is a major environmental issue. To assess and control aircraft emission, one needs an accurate tool for predicting it reliably. Manymethods of prediction are available forNOx emission index in the open literature, while somemethods used by the industries require proprietary information. This paper brings together many important prediction methods (listed in the Appendix) and makes a systematic study of their accuracy and applicability. Finally, a new method called NOx: generic is proposed here, which compares well with the most dependable method (preferred by the industries), viz., the P3 T3 method, but unlike the P3 T3 method the present formulation does not require any proprietary information.

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Lingyun Hou1, Zhen Jia1, Jingsong Gong1, Yifan Xhou1, Ying Piao1 
TL;DR: In this paper, a high-pressure flat plate reactor was used to determine the overall heat sinks and reforming products for hydrocarbon fuel, and the results showed that the resulting heat sinks are improved by using catalytic reforming reactions on the walls of the reactor.
Abstract: The endothermic potential of hydrocarbon-fuel catalytic reforming was developed and demonstrated for hightemperature wall cooling. A high-pressure flat plate reactor was used to determine the overall heat sinks and reforming products for hydrocarbon fuel. Tests were conducted in a catalyst-coated channel that simulates a single passageinapracticalcatalyticheatexchanger/reactorunderregenerative flowconditions.Electricheatingwasused as the heat source. The wall temperature ranged from 400 to 740C. The catalytic and noncatalytic endothermic processes were compared under both subcritical and supercritical pressure conditions. The results show that the overall heat sinks are improved by using catalytic reforming reactions on the walls of the reactor. Hydrogen volume concentrations up to 85–90% are obtained when the catalytic reforming reaction takes place at a temperature of approximately 400C. The catalytic reforming and thermal cracking reactions occur simultaneously from a temperature of 500C, and the hydrogen volume concentration drops to 30%. An increase in pressure results in a decrease in the heat sink and an intensification of the thermal cracking reaction, but it also reduces the extent of catalytic reforming reactions.