Showing papers in "Journal of Propulsion and Power in 1997"

Breakup Processes of Liquid Jets in Subsonic Crossflows

Abstract: The breakup processes of liquid jets injected into subsonic air crosse ows were experimentally studied. Test liquids, injector diameters, and air Mach numbers were varied to provide a wide range of jet operation conditions. Results indicate that for larger injection velocity conditions liquid jets penetrate relatively far into the crosse ows and exhibit surface breakup processes before the column breaks. Liquid column trajectories were correlated by liquid/air momentum e ux ratios based on a force analysis of a cylindrical liquid element subjected to an aerodynamic drag force. Drag coefe cients were inferred from the column trajectories and were found to exhibit a weak dependence on liquid viscosity. The heights of the column fracture points were correlated using the time required for an analogous droplet to complete an aerodynamic secondary breakup process. The success of the resulting correlation justie es the assumption that the aerodynamic forces acting on a droplet and those acting on a liquid column have similar effects. This result, combined with the trajectory correlation, leads to the conclusion that the liquid column always breaks at the same streamwise location, in agreement with the present experimental observation.

Velocity Measurements of a Jet Injected into a Supersonic Crossflow

Abstract: This paper presents a quantitative, experimental study of a single, sonic, underexpanded, transverse, round jet injected into a Mach 1.6 crosse ow. This investigation is applicable to studies of supersonic combustors, thrust vector control of rocket nozzles, the cooling of nozzle walls, and jet reaction force prediction. Schlieren/shadowgraph photography and two-component, frequency preshifted laser Doppler velocimetry are used to visualize the e ow and to measure three mean velocity components, e ve of the six kinematic Reynolds stresses, and turbulent kinetic energy at over 4000 locations throughout the e owe eld. The study focuses on the transverse, midline plane and on two crosse ow planes. These measurements are used to study the size and orientation of the recirculation regions upstream and downstream of the jet; the structure and strength of the bow shock, barrel shock, and Mach disk; the structure, strength, and development of the kidney-shaped, counter-rotating vortex pair; the growth of the annular shear layer between the jet plume and the crosse ow; and the growth of the boundary layer beneath the jet. In addition, the present study provides validation data for analytical and numerical predictions of the transverse jet e owe eld.

Properties of Laminar Premixed CO/H/Air Flames at Various Pressures

Abstract: Effects of positive  ame stretch on the laminar burning velocities of CO/H2/air mixtures were studied both experimentally and computationally for outwardly propagating spherical laminar premixed  ames having concentrations of hydrogen in the fuel mixture of 3 – 50% by volume, fuel-equivalence ratios of 0.6 – 5.0, and pressures of 0.5 – 4.0 atm. Both measured and predicted ratios of unstretched to stretched laminar burning velocities varied linearly with Karlovitz numbers, yielding constant Markstein numbers for each reactant mixture and pressure. Effects of stretch on laminar burning velocities were modest at low hydrogen concentrations, but approached earlier results for hydrogen/air  ames as hydrogen concentrations increased. Predicted and measured  ame properties were in reasonably good agreement using several contemporary chemical reaction mechanisms.

Vortex Shedding in a Large Solid Rocket Motor Without Inhibitors at the Segment Interfaces

Abstract: The internal acoustic response of the Titan IV Solid Rocket Motor Upgrade (SRMU) is analyzed using pressure oscillation time histories measured during four static e ring tests. Pressure oscillations for other large solid rocket motors are caused by vortex shedding at annular inhibitors and acoustic feedback resulting from impingement of the vortices on other inhibitors or on the solid rocket motor nozzle. The SRMU does not have inhibitors, but it is shown that acoustic feedback also dee nes its pressure oscillations. Vortices are shed around the cavity between the center and aft segments and impinge on the nozzle entrance. The frequencies of the pressure oscillations vary about that of the motor fundamental acoustic mode and generally agree with a simple empirical relationship that has been used to model acoustic feedback. The effect of the SRMU pressure oscillations on dynamics and control of the Titan IV system is minor. I. Introduction A NEW solid rocket motor design called the Solid Rocket Motor Upgrade (SRMU) will be integrated with the Titan IV system beginning in late 1996. The SRMU has a lower inert weight and higher propellant weight than the current Titan IV solid rocket motor, and therefore, augments system performance. The SRMU has three segments (forward, center, and aft) and is the only large solid rocket motor in production that does not have annular inhibitors at the segment interfaces to prevent combustion and to support the propellant grain during burning. A. SRMU Static Firing Tests Five static e ring tests were conducted for qualie cation of the SRMU. The test articles included one preliminary qualie cation motor (called PQM19) and four qualie cation motors (called QM1, QM2, QM3, and QM4 ). The tests were conducted vertically, nozzle down at the Phillips Laboratory 1-125 Firing Complex, Edwards Air Force Base. 1 The test cone guration and the attachments of the solid rocket motor to the test stand simulated e ight. The physical differences between the e ve test motors were minor and are believed to have had a negligible effect on combustion stability. The static e ring tests were conducted at propellant bulk temperatures in the range 36.5‐ 937F (2.5‐ 33.97C) to assess the effect on motor performance. The static e ring test instrumentation included Condec gauges (CCC-406) of absolute pressure on the forward closure (measurement PCF1 ) and on the aft dome (measurement PCA1), and a Kistler gauge (202M122) of oscillatory pressure (measurement PD4 ) on the forward closure. 2 Measurements PCF1 and PCA1 were recorded by a Neff digital acquisition system, while measurement PD4 was recorded by a Metraplex FM system. The Neff system applied a 100-Hz low-pass e lter and recorded measurements PCF1 and PCA1 at 250 samples/s.

Comparison of Scramjet Engine Performance in Mach 6 Vitiated and Storage-Heated Air

Abstract: To investigate the sensitivity of combustion to the test gas, an H2-fueled scramjet engine was tested at a Mach 6  ight condition with the air supplied by a combustion heater (V mode) and a storage heater (S mode). The fuel self-ignited without the assistance of igniters in the V mode. However, self-ignition was difŽ cult in the S mode. The easier ignition with vitiated air was caused by radicals supplied from the combustion heater. The combustion behavior was also affected by the test air, which suggests that the combustion was not fully mixing-controlled. As the fuel  ow rate increased, the combustion changed from a weak mode, delivering a lower thrust, to an intensive mode, with a higher thrust. Gas sampling showed that the weak combustion was caused by autoignition in the boundary layer on the engine walls. In the intensive combustion mode, the  ame was anchored near the backward-facing step on the sidewalls. However, the  ame partially detached from the step on the top wall in the combustor. The detached  ame may make the combustion kinetically controlled to produce the sensitivity to the test air.

Computational Study of Rotating-Stall Inception in Axial Compressors

Abstract: Rotating-stall inception in axial-e ow compressors has been studied computationally using a quasi-threedimensional time-marching Navier ‐ Stokes method with a mixing-length turbulence model. Of particular interest was the effect of external circumferential disturbances corresponding to inlet stagnation pressure distortions and rotor/stator blade interactions. The present results show that rotating stall onset patterns in terms of number of stall cells and rotating speeds were ine uenced by small external circumferential stationary or rotational disturbances. First mode circumferential disturbances had the most destabilizing effect, resulting in a single-cell pattern rotating in the absolute frame at about 50% rotor speed, as is observed in most experiments. Short-scale multiple-cell patterns rotating at a higher absolute speed could also be excited by a disturbance with the same circumferential length scale. Short-scale multiple-cell patterns tended to be more persistent in an isolated blade row than in a stage. In the latter case, a shortscale pattern initiated by a rotor ‐ stator interaction would quickly change into a long-scale single-cell pattern, associated with a distinct change of the rotating speed.

Spray characteristics of liquid jet traversing subsonic airstreams

Abstract: The spray characteristics of a liquid jet traversing subsonic airstreams were experimentally investigated. The disintegration phenomena of liquid jets were observed by instantaneous photographs and high-speed video movies. The waves on the liquid jet surface are three dimensional and show very complicated behaviors. These waves are an important cause of the liquid jet disintegration. Droplet mass e uxes were measured using an isokinetic sampling probe. Empirical equations of their distributions that are expressed by the standard normal function were deduced. Droplet sizes and droplet velocities were measured by a phase Doppler particle analyzer. At low air velocity, the mean droplet diameter reaches its maximum in the peripheral mixing region. At high air velocity, however, the mean droplet diameter reaches its maximum in the core region. The droplet velocity peaks in the peripheral mixing region over the whole range of the air velocity.

Thrust Losses in Hypersonic Engines Part 1: Methodology

Abstract: Expressions for the thrust losses of a scramjet engine are developed in terms of irreversible entropy increases and the degree of incomplete combustion. A method is developed that allows the calculation of the lost engine thrust or thrust potential caused by different loss mechanisms within a given e owe eld. This method allows the performance-based assessment of the trade between mixing enhancement and resultant increased e ow losses in scramjet combustors. An engine effectiveness parameter for use in optimization of engine components is dee ned in terms of thrust losses.

Normal, Sonic Helium Injection Through a Wedge-Shaped Orifice into Supersonic Flow

Abstract: Helium was injected normally to a Mach 3 airstream to simulate hydrogen fuel injection in a scramjet combustor. Two geometries were evaluated, a wedge-shaped wall orie ce and a circular wall orie ce. Injection was sonic in both geometries, and the expansion ratios and mass e ow rates were matched to isolate the effects of the geometric difference. Surface oil e ow patterns were inspected to determine the extent of boundary-layer separation upstream of each injector, shadowgraphs were used to visualize the e owe elds, and probe measurements were utilized to determine local helium concentrations. The wedgeshaped injection scheme demonstrated more rapid penetration into the freestream and increased mixing when compared to the baseline circular orie ce. In addition, the oil e ow photography showed that the wedge-shaped injector had no upstream separation zone, whereas the circular injector had a large separation zone. The wedge cone guration would therefore be expected to exhibit reduced wall heat transfer in an actual combustor. It is concluded that wedge-shaped, normal, fuel injectors should provide generally better performance than circular normal injectors in supersonic combustors.

Mach 6 Testing of a Scramjet Engine Model

Abstract: Testing of a subscale scramjet research engine model was carried out in the Mach 6 Ramjet Engine Test Facility of the National Aerospace Laboratory, Kakuda Research Center. The engine had a sidewall compression-type inlet. The fuel was hydrogen. With the attachment of a short strut on the top wall, intensive combustion with high-combustion efe ciency was attained, and the engine-produced thrust canceled the drag. The e ame was held in the low-velocity region around the step, even after the ignitors had been turned off. When the fuel e ow rate was small, there was a different combustion mode with weak combustion and little thrust. The unstart condition seemed to begin around the cowl. Tangential injection of fuel inhibited intensive combustion.

Experimental Study on Rotating Cavitation of Rocket Propellant Pump Inducers

Abstract: Research on rotating cavitation progressed during the development of a liquid oxygen (LOX) turbopump for the LE-7 engine of the H-II rocket In some ranges of cavitation numbers, supersynchronous shaft vibrations were observed in the LE-7 LOX main pump inducer From a comparison with the results of our previous studies it was concluded that such shaft vibrations were caused by rotating cavitation in the inducer A simple modie cation of the inducer upstream housing almost completely extinguished such shaft vibrations Some characteristics of rotating cavitation have been fairly well elucidated However, we were not able to fully explain the mechanism of the ine uence of the simple modie cation of the inducer upstream housing on the rotating cavitation We thus commenced further experimental studies to investigate rotating cavitation in more detail In the present study, some visual observations of rotating cavitation were conducted using the same inducer test facility as that used in our previous work

Multiple-Scales Solution to the Acoustic Boundary Layer in Solid Rocket Motors

Abstract: The acoustic boundary-layer structure is investigated in a cylindrical tube where steady sidewall injection is imposed upon an oscillatory e ow. Culick’ s steady, rotational, and inviscid solution is assumed for the mean e ow. The time-dependent velocity is obtained by superimposing the acoustic (compressible, inviscid, irrotational ) and the vortical (incompressible, viscous, rotational ) velocity vectors. A multiplescales perturbation technique that utilizes proper scaling coordinates is applied to the axial momentum equation by retaining the viscous terms and ignoring the axial convection of vorticity. A closed-form expression for the time-dependent axial velocity is derived that agrees well with the corresponding numerical solution, cold-e ow experimental data, and Flandro’ s near-wall analytic expression. A similarity parameter that controls the thickness of the rotational region is identie ed. The role of the Strouhal number in controlling the wavelength of rotational waves is established. An accurate assessment of the amplitude and phase relation between unsteady velocity and pressure components is obtained. Increasing viscosity is found to reduce the depth of penetration of the rotational region.

Mechanism of Burning Rate Enhancement of Composite Solid Propellants by Ferric Oxide

Abstract: This paper reports a series of experimental studies performed on sandwich propellants, wherein a matrix lamina of particulate oxidizer and polymeric binder is sandwiched between two ammonium perchlorate (AP) laminae. The catalyst (ferric oxide ) is incorporated in the matrix lamina. The variables are pressure (0.345‐ 6.9 MPa), matrix lamina thickness, catalyst concentration, matrix mixture ratio, types of oxidizer and binder, and the dispersion ability of the catalyst. The combined results indicate that, under the conditions tested, near-surface reactions associated with the particulate AP/binder contact lines on the burning surface assume signie cance in the presence of the catalyst. These reactions are further augmented by the presence of the leading-edge portion of the diffusion e ame above the interface of the matrix and AP laminae.

Stability Testing of Full-Scale Tactical Motors

Abstract: The U.S. Naval Air Warfare Center has participated in a program to develop an improved understanding of linear and nonlinear combustion instability in solid propellant rocket motors. One goal of this program was to develop a systematic database of motor and stability data. This paper describes the linear aspects of the motor e rings and analysis. The motors that were used had diameters of 127 mm and were 1.7 m long. The majority were loaded with an 88% solids reduced-smoke ammonium perchlorate propellant with a nominal burning rate of 6.1 mm/s at 6.9 MPa. In addition, motors have been e red that contain 1% 8- mm aluminum oxide, 90- mm aluminum oxide, and 3- mm zirconium carbide as stability additives in place of 1% ammonium perchlorate. Motor pressures ranged from 3.45 to 10.34 MPa and various grain geometries were tested. Pressure-coupled combustion response measurements were made at the nominal motor operating pressures. Motor performance and stability calculations were made using the Air Force Solid Performance Program and the Standard Stability Prediction Program for the motor cone gurations that were e red. The stability predictions were compared to the data obtained from the motor e rings. The results indicate that it is possible to theoretically predict linear motor stability for the class of motors discussed in this paper. It was also learned that higher-pressure motors tend to be less

Evaluation of Performance Loss Methods for High-Speed Engines and Engine Components

Abstract: A comparative study of high-speed engine performance assessment techniques based on exergy (available work) and thrust potential (thrust availability) is summarized. Simple one-dimensional  owŽ elds utilizing Rayleigh heat addition and friction are used to demonstrate the inability of conventional exergy techniques to predict engine component performance, aid in component design, or accurately assess  ow losses. The thrust-based method yields useful information in all of these categories for these  ows. The conventional deŽ nition of exergy includes work that is inherently unavailable to an aerospace Brayton engine. An engine-based exergy is developed that accurately accounts for this inherently unavailable work; performance parameters based on this quantity yield design and loss information identical to the thrustbased method.

Challenge to Create the Space Drive

Abstract: To travel to our neighboring stars as practically as envisioned by science fiction, breakthroughs in science are required. One of these breakthroughs is to discover a self-contained means of propulsion that requires no propellant. To chart a path toward such a discovery, seven hypothetical space drives are presented to illustrate the specific unsolved challenges and associated research objectives toward this ambition. One research objective is to discover a means to asymmetrically interact with the electromagnetic fluctuations of the vacuum. Another is to develop a physics that describes inertia, gravity, or the properties of space-time as a function of electromagnetics that leads to using electromagnetic technology for inducing propulsive forces. Another is to determine if negative mass exists or if its properties can be synthesized. An alternative approach that covers the possibility that negative mass might not exist is to develop a formalism of Mach's principle or reformulate ether concepts to lay a foundation for addressing reaction forces and conservation of momentum with space drives.

Improvements to steady-state combustion modeling of cyclotrimethylenetrinitramine

Abstract: Many improvements have been made to models of cyclotrimethylenetrinitramine combustion in the past few years. The one-dimensional model presented in this paper models the solid, two-phase, and gas regions using complex kinetics and concentrationand temperature-dependent thermophysical properties. Calculated values agree well with experimentally determined burning rate sp, melt layer thickness, surface temperature, and species concentration proŽ les. When including laser-assisted burning in the model, a dark zone appeared that was similar to that seen experimentally. With the laser-assisted case, the chemistry controlling the burning rate is signiŽ cantly different from cases without the laser heat  ux. Calculations show that the melt-layer thickness is determined primarily by the liquid thermal conductivity and the surface temperature is controlled by the vapor pressure correlation. All other model predictions are relatively insensitive to these parameters. The condensed-phase decomposition and evaporation/condensation submodel are the weakest areas of the model.

Thrust losses in hypersonic engines part 2 : applications

Abstract: The methodology developed in Part 1 of this investigation (Riggins, D. W., McClinton, C. R., and Vitt, P. H., ‘‘Thrust Losses in Hypersonic Engines Part 1: Methodology,’’ Journal of Propulsion and Power, Vol. 13, No. 2, 1997, pp. 281 – 287) for the identiŽ cation and evaluation of the thrust losses in a scramjet engine is applied to one-dimensional scramjet engine  ows with coupled loss mechanisms. Thrust and thrust potential losses are related directly to increases in irreversible entropy caused by friction, heat transfer, mixing, nonequilibrium reaction, and shocks. This method is extended to enable the evaluation of thrust losses in multidimensional  ows. The fundamental relationship between performance assessment utilizing multidimensional  owŽ elds and cycle analysis performance prediction is shown and discussed.

Nonlinear Stability Testing of Full-Scale Tactical Motors

Abstract: The U.S. Naval Air Warfare Center has participated in a program to develop an improved understanding of linear and nonlinear combustion instability in solid propellant rocket motors. One goal of this program was to develop a systematic database of motor stability data. This paper describes the nonlinear aspects of the motor e rings and analysis. The motors used had diameters of 127 mm and were 1.7 m long. The majority were loaded with an 88% solids reduced-smoke ammonium perchlorate propellant with a nominal burning rate of 6.1 mm/s at 6.9 MPa. Motor pressures ranged from 3.45 to 10.34 MPa and various grain geometries were tested. In addition, motors have been e red that contain 1% 8- mm aluminum oxide, 90- mm aluminum oxide, and 3- mm zirconium carbide as stability additives in place of 1% ammonium perchlorate. This paper discusses experimental methods of motor pulsing and various theoretical approaches to predict pulse amplitudes in motors. The paper also examines nonlinear acoustic motor response to various amplitudes of acoustic pulsing and the resultant characteristics of sustained nonlinear oscillations. Finally, some theoretical interpretations are presented from the experimental motor data. The results show that there is a direct relationship between the dc pressure shift and the magnitude of the acoustic oscillations.

Plug Nozzle Flowfield Analysis.

Abstract: Results of numerical simulations of plug-nozzles are presented, and a phenomenological insight into the e owe eld development at different ambient pressures is given. Therefore, a e nite volume program using unstructured grids was adapted to the special boundary conditions of plug nozzles. Calculations were performed solving the Euler and Navier ‐ Stokes equations under ideal, perfect gas assumptions. Turbulence is taken into account with a k-e turbulence model. Numerical simulations are compared with experimental results of hot-run tests of a toroidal subscale model plug engine. Principal physical processes like expansion waves, compression shocks, and the recirculating base e ow region are in good agreement with available experimental data, and can therefore be predicted well. The simulations of a full-size plug nozzle, dee ned for a post Ariane 5 launcher, reveal a e owe eld behavior similar to the one observed with the toroidal subscale plug engine. Nomenclature I = impulse Mmix = molar mass p = pressure rO/F = mass ratio oxidizer/fuel mixture T = temperature a = angle « = nozzle area ratio k = isentropic coefe cient

Jet Engine Component Maps for Performance Modeling and Diagnosis

Abstract: This paper describes an effort to model the performance maps of compressors and turbines (i.e., the relation between mass e ow, pressure ratio, and efe ciency ), using analytical functions. Analytical functions are e tted to the available experimental data using a least-squares-type approach for determining the parameters of the e tting function. The success of using a particular function for an application is assessed through a suitably dee ned mean error of the model. Apart from presenting the method for setting up these analytical representations, applications to performance modeling and fault diagnosis are discussed. The change in model parameters is used to characterize changes of the engine condition and possibly diagnose occurring faults. The impact of introducing analytical component models into overall engine computer models, replacing a tabulated form of the component maps, is also discussed.

Study of a Supersonic Combustor Employing Swept Ramp Fuel Injectors

Abstract: A computational study of a scramjet combustor with swept ramp fuel injectors was conducted with the SPARK Navier ‐ Stokes computer code. Turbulence was modeled with a two-zone algebraic turbulence model, while combustion was modeled with two e nite rate H 2 ‐ air chemistry models. The calculated reacting e owe eld was mixing-limited. However, the extent of mixing was affected by the chemical reactions. Heat release from the chemical reactions signie cantly reduced the mixing between the fuel and airstreams. Comparisons between the calculations and experimental measurements of the wall pressure, surface heat e ux, in-stream pitot pressure, as well as Mie scattering e ow visualization were made and good overall agreement was observed.

Theoretical Formulation for Sauter Mean Diameter of Pressure-Swirl Atomizers

Abstract: This work derives a general formula for determining the Sauter mean diameter of sprays generated by pressure-swirl atomizers. This is done by extending the theory of the aerodynamic instability and disintegration of viscous liquid e at sheets to describe the dynamic behavior of the hollow conical sheet generated by these atomizers. The derived theoretical equation for the conical spray droplets mean diameter includes atomizer geometrical parameters as well as liquid fuel and atomizing gas e ow parameters. The results compare well with several existing empirical and semiempirical correlation formulas and available experimental data.

Effects of Mixing Schemes on Kerosene Combustion in a Supersonic Airstream

Abstract: A study of kerosene combustion in a supersonic vitiated aire ow at Mach 4.75 e ight enthalpy was conducted in direct-connect tests at Mach 1.8 at a stagnation temperature of 1000 K. The effects of shockand vortex-enhanced mixing mechanisms on the combustion efe ciency were evaluated. Also included in this study were the effects of fuel heating and jet penetration. The experimental conditions corresponded to the low end of the hypersonic e ight regime. The following geometric cone gurations were employed: 1) a generic, rearward-facing step, 2 ) a modie ed rearward-facing step with beveled edges to facilitate vortex-enhanced mixing, and 3 ) a rearward-facing wedge (15 or 30 deg) placed downstream of the rearward-facing step to induce shock-enhanced mixing. In all cone gurations, a gaseous hydrogen ‐ pilot jet was injected parallel to the main e ow from the base of the rearward-facing step and the liquid kerosene was injected normal to the main e ow at three or e ve step heights downstream of the step (the step height was 10 mm). Stable kerosene combustion was obtained for a maximum injected kerosene equivalence ratio of 0.86. For efe ciency evaluation, the pilot ‐ hydrogen equivalence ratio was selected between 0.02 ‐ 0.04, while the kerosene equivalence ratio was maintained at 0.325. In all experiments, locally rich stratie ed kerosene combustion took place in a layer close to the injection wall. The wedge e ameholder contributed to an increased kerosene combustion efe ciency by the generation of shock ‐ jet interactions. The beveled-edge step improved far-e eld mixing, thereby reducing the local kerosene equivalence ratio, resulting in the highest kerosene combustion efe ciency among all cone gurations tested. Fuel heating below levels required for e ash vaporization (one-third of the e ash vaporization energy, in this case ) did not contribute to increased combustion efe ciency. On the contrary, this level of heating reduced the fuel density with adverse effects on penetration and mixing.

Performance and Acceleration Process of Quasisteady Magnetoplasmadynamic Arcjets with Applied Magnetic Fields

Abstract: A quasisteady magnetoplasmadynamic (MPD) arcjet with an applied magnetic Ž eld was investigated to improve the thruster performance and understand the complex acceleration mechanisms with both the self-induced and applied magnetic Ž elds. The MPD arcjet was operated with hydrogen, a mixture of nitrogen and hydrogen simulating hydrazine, and argon at discharge currents of 3 – 18 kA in high speciŽ c impulse levels around a critical discharge current predicted from the rules of Alfven’s ionization velocity or minimum input power. The application of axial magnetic Ž elds achieved higher thrust efŽ ciencies than those for only the self-induced magnetic Ž eld at constant speciŽ c impulses, and still achieved stable operations at higher speciŽ c impulses with less electrode erosion. The following guidelines were suggested to achieve higher thruster performance: 1) the axial magnetic Ž eld strength must be smaller than the azimuthal self-Ž eld strength in the main discharge region near the cathode tip, and 2) the applied magnetic Ž eld lines must expand gradually downstream for smooth expansion of plasma. Furthermore, the measured pressures on the electrodes and the current distributions in the discharge chamber showed that the overall thrust measured by a pendulum method increased, in spite of a decrease in the electromagnetic pumping thrust and a small contribution of Hall acceleration. Thus, an additional thrust component because of the axial magnetic Ž eld, such as that caused by swirl acceleration, is expected to exist.

Counterflow Thrust Vector Control of Subsonic Jets: Continuous and Bistable Regimes

Abstract: Fluidic thrust vectoring of a subsonic jet was examined by the asymmetric application of countere ow to the jet shear layers. When countere ow is applied to one side of a 4:1 aspect ratio rectangular jet in the proximity of a control surface, the shear-layer entrainment characteristics are altered asymmetrically giving rise to a cross-stream pressure gradient and e ow vectoring. Thrust vector control up to 20 deg was possible at Mach numbers up to 0.5 using countere ow. Two regimes of e uidic control were identie ed, a continuous regime and a bistable one. During continuous vectoring, proportional control could be achieved between jet response and the pressure conditions established in the countere owing stream. However, under certain circumstances, proportional control was lost leading to jet bistability. Since such bistability is undesirable for aircraft control applications, the phenomenon was studied in detail. A parametric study of the collar geometry was used to minimize jet attachment and expand the operating domain over which proportional control could be maintained.

Mixing Characteristics of Forced Mixers with Scalloped Lobes

Abstract: An experimental investigation has been undertaken to study the mixing characteristics of forced mixers with scalloped lobes using a two-component e ber-optic laser Doppler anemometer at a Reynolds number of 2.27 3 10 4 and at a velocity ratio 1:2 across the lobes. The trailing edge of the mixers, without scalloping, had the shape of a square wave, a semicircular wave, and a triangular wave. Scalloping was achieved by eliminating up to 70% of the sidewall area at the penetration region of each lobe. The aspect ratio of each lobe (lobe height-to-wavelength ratio ) was at unity and the half-angles at the penetration region were 22 and 35 deg for the scalloped and the aggressively scalloped mixers, respectively. The wavelength and lobe height were the same for both types of mixer. The results showed that lobe cone gurations appeared to be more important than the penetration angles for the benee t of scalloping to occur. Strengths of the streamwise circulation near the trailing edge for respective scalloped mixers were higher than the nonscalloping cases, largely because of the formation of two pairs of streamwise vortices at each lobe. The subsequent decay rates with downstream distance were also found to be more rapid, indicating that a faster mixing rate can actually be achieved by the scalloped mixers. Additional production of turbulent kinetic energy appeared at about two to three wavelengths downstream of the trailing edge and then asymptote to lower level in the far-e eld region. Spatial uniformity of the mass e ux at the wake region could be achieved at two wavelengths further downstream of the high turbulent kinetic energy region.

Zonal Approach for Prediction of Jet Noise

Abstract: A zonal approach for direct computation of sound generation and propagation from a supersonic jet is investigated. The present work splits the computational domain into a nonlinear, acoustic-source regime and a linear acoustic wave propagation regime. In the nonlinear regime, the unsteady flow is governed by the large-scale equations, which are the filtered compressible Navier-Stokes equations. In the linear acoustic regime, the sound wave propagation is described by the linearized Euler equations. Computational results are presented for a supersonic jet at M = 2. 1. It is demonstrated that no spurious modes are generated in the matching region and the computational expense is reduced substantially as opposed to fully large-scale simulation.

Tangential Injection from Overlaid Slots into a Supersonic Stream

Abstract: The results of an experimental study of tangential supersonic slot injection into a supersonic airstream with tangential subsonic injection through an additional slot above the main slot is presented. Such a  ow is of interest for at least two propulsion-related applications. First, one can simulate a near-separation boundary-layer proŽ le by subsonic injection through the upper slot, which is then to be re-energized by supersonic injection through the lower slot. Second, this  ow is also of interest for application to fuel/ oxidizer/pilot injection from multiple overlaid slots. The experiments were performed in an intermittent, vacuum wind tunnel at a freestream Mach number of 2.85. The supersonic injectant had a Mach number of 2.00, and the subsonic injection was at Mach numbers of 0.26 and 0.72. The results are presented in the form of spark schlieren photographs, interferograms, and wall-static pressure measurements. Density proŽ les at several axial locations determined from the interferograms are presented, as well as streamwise and spanwise static pressure distributions. The major conclusions drawn from the tandem injection results, are Ž rst, that a near-separation boundary-layer proŽ le can be simulated in this way. Next, tandemly injected subsonic and supersonic  ow can be divided into separate components that closely resemble the respective individual injections into an undisturbed freestream. Also, the effects of the subsonic injection were completely mixed out at a downstream location of six slot heights. Therefore, adverse-pressuregradient-inducing devices should be positioned at least six slot heights downstream of the supersonic injection station if the effects of supersonic injection into a near-separation boundary layer for the purpose of re-energizing it are to be studied.

Tip Clearance Flow-Shock Interaction in a Transonic Compressor Rotor

Abstract: A combination of experimental, computational, and analytical efforts were used to describe the phenomenological features of the interaction between the tip clearance e ow and the passage shock in a transonic, axial compressor rotor. Unsteady static pressure was measured at the casing over the rotor, and steady measurements of total pressure and total temperature were made at the rotor exit. A fully three-dimensional, steady, Navier ‐ Stokes computational e uid dynamics technique was used to obtain a solution to the e owe eld at design conditions that compared favorably with the experimental measurements. Finally, a simple model was developed that predicts the kinematic and thermodynamic properties within the pre- and postshock vortex. Analysis of the numerical results revealed a wake-like nature of the vortex both upstream and downstream of the shock. The upstream character was strongly ine uenced by the injection of mass into the vortex via the clearance e ow. The success of the simplie ed model’ s ability to predict the properties of the vortex within the interaction region allows two important conclusions to be drawn. First, the vortical character of the vortex upstream of the shock is not a signie cant factor in driving the interaction. Second, the interaction between the clearance e ow and the shock is fundamentally the result of the change in momentum brought about by the shock-induced pressure rise. The interaction can therefore be viewed as an inviscid phenomenon.