Showing papers on "Supersonic speed published in 1996"

On the Two Components of Turbulent Mixing Noise from Supersonic Jets

Abstract: It is argued that because of the lack of intrinsic length and time scales in the core part of the jet flow, the radiated noise spectrum of a high-speed jet should exhibit similarity. A careful analysis of all the axisymmetric supersonic jet noise spectra in the data-bank of the Jet Noise Laboratory of the NASA Langley Research Center has been carried out. Two similarity spectra, one for the noise from the large turbulence structures/instability waves of the jet flow, the other for the noise from the fine-scale turbulence, are identified. The two similarity spectra appear to be universal spectra for axisymmetric jets. They fit all the measured data including those from subsonic jets. Experimental evidence are presented showing that regardless of whether a jet is supersonic or subsonic the noise characteristics and generation mechanisms are the same. There is large turbulence structures/instability waves noise from subsonic jets. This noise component can be seen prominently inside the cone of silence of the fine-scale turbulence noise near the jet axis. For imperfectly expanded supersonic jets, a shock cell structure is formed inside the jet plume. Measured spectra are provided to demonstrate that the presence of a shock cell structure has little effect on the radiated turbulent mixing noise. The shape of the noise spectrum as well as the noise intensity remain practically the same as those of a fully expanded jet. However, for jets undergoing strong screeching, there is broadband noise amplification for both turbulent mixing noise components. It is discovered through a pilot study of the noise spectrum of rectangular and elliptic supersonic jets that the turbulent mixing noise of these jets is also made up of the same two noise components found in axisymmetric jets. The spectrum of each individual noise component also fits the corresponding similarity spectrum of axisymmetric jets.

Coating or ablation applicator with a debris recovery attachment

Abstract: An environmentally compliant triboelectric applicator and process for coating or ablating a substrate and for retrieving excess or ejected material from the substrate. The applicator comprises an inner supersonic nozzle for accelerating triboelectrically charged projectile particles entrained in a supersonic gas to speeds sufficiently high to coat or ablate a substrate. The applicator further comprises an outer evacuator nozzle coaxially surrounding the inner supersonic nozzle for retrieving excess projectile particles, ablated substrate powders, or other environmentally hazardous materials. A fluid dynamic coupling uses the efficacy of the Mach turning angle associated with a supersonic boundary layer of carrier gas to aspirate the central core of the supersonic two-phase jet. This fluid coupling and spacing between the outlet of the supersonic nozzle and the substrate also permits the substrate to triboelectrically charge to levels which induce electrostatic discharges at the substrate simultaneous to the impacts. The aspiration feature reduces the outlet pressure in the central core of the nozzle below ambient pressure which allows the projectiles to travel unimpeded to the substrate, and reduces the inlet pressure required to achieve parallel and shock-free flow with the inner supersonic nozzle. The powders are injected into the carrier gas using powder feeders modified for high pressure and the ablated debris with excess projectile particles are collected in a particle precipitator and filter unit using a suction blower. A special nozzle applicator embodiment and process for coating or ablating the internal surface of a cylinder bore is also disclosed. This nozzle applicator comprises an axisymmetric cylindrical nozzle for conveying, accelerating, and triboelectrically charging projectile particles entrained in a carrier gas to speeds sufficiently high to coat or ablate the internal surface of the cylinder bore substrate when impacted by a triboelectrically charged jet comprising a central core.

Starting characteristics of supersonic inlets

Abstract: The starting characteristics of a small-scale rectangular inlet with a thick ingested boundary layer were investigated at nominal Mach 3 conditions. Parameters investigated included Reynolds number, cowl length, and cowl height. Measurements of the maximum and restart contraction ratios were made. Depending on the test configuration, the unstarts were classified into two broad categories as either "hard" or "soft." The hard unstarts appear to occur when the flow at the inlet throat chokes. The soft unstarts occur as large-scale separation develops within the inlet. The ability of the classical Kantrowitz limit to predict the restart contraction ratio was assessed, and it was shown to be applicable for the hard unstart/ restart configurations. The role of fluid injection upstream of the unstarted inlet was also assessed. The use of this injection may ultimately lead to improving the starting characteristics of inlets. Nomenclature A Area hc Cowl height Lc Cowl length rh Mass flow M Mach number M Mass-averaged Mach number P Pressure Pt Freestream total pressure x, y Cartesian coordinates y Ratio of specific heats 6 Boundary layer thickness 6* Boundary layer displacement thickness 0 Boundary layer momentum thickness 9C Cowl angle p Density * Principal staff engineer, senior member AIAA ** Senior staff engineer, member AIAA ^ Associate staff engineer, member AIAA Copyright © American Institute of Aeronautics and Astronautics, Inc., 1996. All rights reserved. Subscripts 0 Freestream 2 Entrance to cowl 4 Inlet throat cl Wall conditions at cowl lip inj Injectant Introduction Airbreathing engines that operate at supersonic and hypersonic speeds require inlets to capture and compress air for processing by the remainder of the engine. The goal in the design of any inlet is to define a minimum weight geometry that provides an efficient compression process, generates minimum drag, produces nearly uniform flow entering the compressor or combustor, and provides these characteristics over a wide range of flight and engine operating conditions. For efficient operation and moderate induced drag, most inlets use a combination of external and internal compression. The introduction of internal contraction in an inlet adds complexity in the design and analysis process in that the starting of the inlet must be ensured. For efficient operation, supersonic and hypersonic inlets must operate in a started mode. The process of inlet starting and unstarting is well understood at a conceptual level, although significant details remain to be resolved. Some variation exists in the very definition of a started inlet. One convention states that a started inlet is one with supersonic flow in the inlet throat, but it is well known that some unstarted inlets can have complex internal flowfields with a significant fraction of supersonic flow in the inlet throat. In the present work, the term "started" is used to denote operation under conditions where flow phenomena in the internal portions of the inlet do not alter the air capture characteristics of the inlet. (Reduction in the captured mass flow through the use of bleed holes or bypass channels is not considered in assessing whether an inlet is started.) An inlet can be unstarted by either over-contracting to the point where the flow chokes at the inlet throat or by raising the back pressure beyond the level that can be sustained by the inlet. Currently, a significant uncertainty exists regarding the conditions under which an inlet will unstart or restart. Part of this uncertainty is due to the large variety of 1 American Institute of Aeronautics and Astronautics geometries that have been considered in designing engines. The design of an inlet is strongly affected by vehicle considerations, and a variety of two-dimensional planar, axisymmetric, and three-dimensional inlet designs have been investigated. The diversity found in inlet designs can be seen in the sample inlets shown in Fig. I.'" Preliminary estimates of the internal contraction that will self-start can be obtained from the Kantrowitz limit. This limit is determined by assuming a normal shock wave at the beginning of the internal contraction and calculating the one-dimensional, isentropic internal area ratio that will produce sonic flow at the inlet throat. For a perfect gas, the Kantrowitz limit can be calculated as follows: Oswatisch inlet HRE-type inlet A2] /KANTROWITZ M, l)MJ

Dynamic crack propagation in piezoelectric materials—Part II. Vacuum solution

Abstract: In Part I of this work, antiplane dynamic crack propagation in piezoelectric materials was studied under the condition that crack surfaces behaved as though covered with a conducting electrode. Piezoelectric surface wave phenomena were clearly seen to be critical to the behavior of the moving crack. Closed form results were obtained for stress and electric displacement intensities at the crack tip in the subsonic crack speed range; the major result is that the energy release rate vanishes as the crack speed approaches the surface (Bleustein-Gulyaev) wave speed. In this paper, an alternative assumption is made that between the growing crack surfaces there is a permeable vacuum free space, in which the electrostatic potential is nonzero. By coupling the piezoelectric equations of the solid phase with the electric charge equation in the vacuum region, a closed form solution is again obtained. In contrast to the electrode case of Part I, this case allows both applied charge and applied traction loading. In addition, the work of Part I is extended to examine piezoelectric crack propagation over the full velocity range of subsonic, transonic and supersonic speeds. Several aspects of the results are explored. The energy release rate in this case does not go to zero when the crack propagating velocity approaches the surface wave speed, even if there is only applied traction loading. When the crack exceeds the Bleustein-Gulyaev wave speed, the character of the crack-tip singularities of the physical fields depends on both speed regime and type of loading. At the other extreme, the quasi-static limit of the dynamic solution furnishes a set of new static solutions. The general permeability assumptions made here allow for fully coupled conditions that are ruled out by the a priori interfacial assumptions made in previously published solutions.

Counterflow thrust vectoring of supersonic jets

Abstract: Fluidic thrust vector control is examined in a supersonic rectangular jet having a 4 :1 aspect ratio and for jet stagnation temperatures between 300 and 670 K. Experiments conducted in a nominally ideally expanded Mach 2 jet reveal that thrust can be continuously vectored up to at least 16 deg by creating a secondary counterflowing stream between the primary jet and an adjacent curved surface. Thrust vector control using counterflow is shown to be effective in both cold and moderately heated supersonic jets and to perform free of bistable or hysteretic effects. Measurements indicate that proportional thrust vector control can be achieved with less than 4% thrust loss and requiring secondary mass flow rates less than approximately 2% of the primary jet.

Supersonic-inlet boundary-layer bleed flow

Abstract: Boundary-layer bleed in supersonic inlets is typically used to avoid separation from adverse shock-wave/boundary-layer interactions and subsequent total pressure losses in the subsonic diffuser and to improve normal shock stability. Methodologies used to determine bleed requirements are reviewed. Empirical sonic flow coefficients are currently used to determine the bleed hole pattern. These coefficients depend on local Mach number, pressure ratio, hole geometry, etc. A new analytical bleed method is presented to compute sonic flow coefficients for holes and narrow slots and predictions are compared with published data to illustrate the accuracy of the model. The model can be used by inlet designers and as a bleed boundary condition for computational fluid dynamic studies.

Bow shock/jet interaction in compressible transverse injection flowfields

Abstract: In flowfields created by transverse injection into supersonic cross-flows, the bow and separation shocks formed upstream of the injectant plume are dominant features. In the present investigation, the interaction between these features and the large-scale eddies that develop at the jet/freestream interface has been examined.

Investigation of Large-Scale Structures in Supersonic Planar Base Flows

Abstract: Time-resolved, planar imaging was employed to study the spatial organization of large-scale structures within the shear layers, at reattachment, and in the wake region of a supersonic base flow. Side and end views were obtained at several streamwise locations to characterize the evolution and three-dimensionality of the large-scale motions. From statistically significant ensembles, spatial correlation fields were computed to quantify the mean size, eccentricity, and orientation of the large structures. Visualizations confirm that large-scale turbulent structures exist at all stations in the shear layers and interact vigorously with the recirculation region. Mach and/or shock waves are frequently seen emanating from within the shear layer, which may be indicative of eddy shocklet formation. The embedded turbulent structures are elliptical in shape and usually appear inclined to the mean flow direction. A distinct flattening and tilting in the streamwise direction occur as the coherent eddies negotiate shear layer formation, recompression, and reattachment processes. Spatial statistics indicate that the structures have a streamwise and spanwise spatial extent on the order of the local shear layer thickness.

Parietal vortex shedding as a cause of instability for long solid propellant motors - Numerial simulations and comparisons with firing tests

Abstract: This paper describes a new mechanism leading to vortex shedding instabilities in long (large L/D) solid propellant motors. This mechanism is termed "parietal vortex shedding" and has been discovered thanks to numerical simulations of the unsteady, 2D, compressible, Navier-Stokes equations. It seems to involve hydrodynamic instabilities of the mean flow velocity profiles, corresponding to injection induced internal flow (so called Culick or Taylor profiles), that couple with the acoustic frequencies of the chamber. Although this mechanism is found to be very powerful, it seems to need some background noise to feed it. The presented simulations can explain observed instabilities in configurations, without segmentation or without protruding inhibitor rings, of a simplified subscale setup of the Ariane 5 MPS P230 solid propellant motor. Detailed comparisons are proposed and the influence of the propellant combustion response and of the turbulence of the flow field are analyzed by means of recently developed models. INTRODUCTION-OBJECTIVES not precisely known if inhibitor rings are or not destroyed or completely slack; on the other hand non segmented motors have also given rise to a same type of instability as vortex shedding : this is the case of one of the configuration (LP3D) of the LP3 set up presented in a preceding paper. As it is known ' ' , classical linear acoustic balance computations do not give reliable stability predictions in complex internal geometries (such as in the P230 motor), and then an effort was carried out to perform full numerical simulations of the unsteady, compressible internal flow fields. On the other hand, motor internal flows are mostly non-observable, and without numerical simulations it is not possible to describe the path of the aerodynamic instability development. The object of the present paper is to show hydrodynamic instabilities (which drive pressure oscillations) in configurations without shear flows induced by inhibitor rings, and to explain how these instabilities occur by means of numerical simulations. LP3D AND LP3E TEST CASES This work is part of the overall research effort, supported by CNES, accompanying the development of the Ariane 5 P230 MPS solid propellant motor (program ASSM for Aerodynamics of Segmented Solid Motors) and makes use of experimental results obtained during the combustion stability assessment program carried out for BPD and CNES, by delegation of ESA. In this scope, it is the continuation of earlier works about numerical simulations in solid propellant rocket motors ". The Ariane 5 motor, as other large segmented motors (U.S. Space Shuttle and Titan SRM) has been reported to exhibit pressure and thrust oscillations. Until recently, it was believed that such instability was exclusively due to the segmented design : inhibitor rings induce shear layers and vortex shedding driven oscillations. Nevertheless it is * Research scientist, Energetics Dept. 1 Project manager, Energetics Dept., Member AIAA Copyright © 1996 by the American Institute of Aeronautics and Astronautics, Inc., All rights reserved. The test cases are based on two configurations of the LP3 motor. The LP3 motor, which has been presented in reference , is a simplified 1/15 subscale set-up of the Ariane 5 P230 motor used to test several inter-segment arrangements. The configurations of interest here, called LP3D and LP3E, have no prominent obstacles at the mid chamber point, see figure 1. Both configurations have a cylindrical chamber of length 1632 mm and inner diameter 203 mm. At ignition, the main propellant burning surface is cylindrical, inner diameter 90 mm, with a chamfered surface near the aft end. The supersonic outlet nozzle is submerged, with a throat diameter of 56.5 mm, then the total length of the motor is 1650 mm. LP3D and LP3E have a small forward segment of 225 mm length, cylindrical for LP3D (169.5

Algebraic turbulence model simulations of supersonic open-cavity flow physics

Abstract: A time-accurate double thin-layer Navier-Stokes computation is performed for an unsteady supersonic open cavity with a length-to-depth ratio of 2 The results are used to determine the flow-physics mechanisms responsible for the cavity oscillation cycle A new cycle is described and compared to previous descriptions It is found that a shed vortex impinges on the cavity aft lip and forms a pressure pulse that augments or forces, at the vortex shedding frequency, an internal upstream moving wave that has been reflected from the aft corner This upstream moving wave eventually reflects off the cavity forward wall and forces the shedding of a new vortex It was found, however, that the reflected wave dissipates before it reaches the aft wall Instead, a second wave forms beneath the shed vortex and eventually reflects from the aft corner and is forced at the shedding frequency by the shed vortex wave, completing the cycle

Complete three-dimensional multiparameter mapping of a supersonic ramp fuel injector flowfield

Abstract: Planar laser-induced iodine fluorescence is used to map out the nonreacting mixing flowfield of an unswept ramp fuel injector using air injected at Mach 2.0 into a Mach 2.9 freestream. A fully automated test setup is used to measure time-averaged pressure, temperature, velocity, and injectant mole fraction on 21 crossflow planes and 7 axial planes. The measurement uncertainties are 5-8% for temperature, 4-10% for pressure, 10-20 m/s for velocity, and 2-3% for injectant mole fraction depending on the thermodynamic conditions. The measurements allow any desired gasdynamic quantity to be determined on a three-dimensional grid that spans the entire wind-tunnel test section. The experimental data set is comparable to the completeness of results normally available only from a computational fluid dynamics simulation. Results showing detailed flow features on specific planes, as well as overall quantities, such as global conservation checks, mixing performance, and flowfield losses, are presented. Mass, momentum, and energy flux, determined at the crossflow plane locations of the data set, show about a 2% standard deviation. The results are compared to a simulation using a three-dimensional Navier-Stokes solver. Agreement is reasonable with the exception of measurements in regions very close to walls, where the intensity of scattered light is high or where optical access is limited. The ability to generate extensive data sets, such as the one presented here, demonstrates that the planar laser-induced iodine fluorescence technique can be used 1) to generate detailed test cases for the validation of computational fluid dynamics codes and 2) as an alternative to computational fluid dynamics for performing design studies and performance evaluation in complex compressible flows.

Application of absorption filter planar Doppler velocimetry to sonic and supersonic jets

Abstract: An absorption filter planar Doppler velocimeter has been constructed and tested on a pressure-matched sonic jet and an overexpanded supersonic jet (A/design = 1-9). The current system can acquire single-velocity-component single-shot planar images (15-ns exposures) at 30 Hz. A unique aspect of the system is the use of a single camera and lens, made possible by the application of an image splitter system. This development reduces the expense and experimental complexity of the technique. The most extensive of the three data sets taken comprises 1100 centerline planes of instantaneous velocity acquired in the overexpanded supersonic jet. Mean and rms velocity fields are presented for this data set; a mean centerline and a mean transverse profile are extracted and compared with profiles from a computational fluid dynamics solution generated using the LARCK computer code employing a k-e turbulence model.

An experimental study of supersonic microjets

Abstract: Miniature axisymmetric supersonic nozzles were produced with exit Mach numbers ranging from 10 to 28 by forming Pyrex® capillary tubing of 06 and 12 mm inside diameter into converging-diverging channels The nozzle contours were measured and were found to compare favorably to ideal solutions given by the axisymmetric method of characteristics In addition, the surfaces of these nozzles were quite smooth, providing featureless flows at perfect expansion Schlieren visualization and pitot pressure measurements of the resulting microjets were compared to the literature available for jets produced by larger-scale nozzles A postponed transition to turbulence is noted in these microjets due to their low Reynolds number The pitot pressure on centerline is nearly uniform at perfect expansion over core lengths up to 12 nozzle exit diameters Supersonic microjet nozzles thus provide a more effective small-scale high-pressure gas delivery device than do sonic nozzles of comparable scale at equivalent mass flow rates Supersonic microjets may therefore have several industrial applications

Spreading Characteristics and Thrust of Jets from Asymmetric Nozzles

Abstract: The spreading characteristics of jets from several asymmetric nozzles are studied in comparison to those of an axisymmetric jet, over the Mach number (M(sub J)) range of 0.3 to 1.96. The effect of tabs in two cases, the axisymmetric nozzle fitted with four tabs and a rectangular nozzle fitted with two large tabs, is also included in the comparison. Compared to the axisymmetric jet, the asymmetric jets spread only slightly faster at subsonic conditions, while at supersonic conditions, when screech occurs, they spread much faster. Screech profoundly increases the spreading of all jets. The effect varies in the different stages of screech, and the corresponding unsteady flowfield characteristics are documented via phase-averaged measurement of the fluctuating total pressure. An organization and intensification of the azimuthal vortical structures under the screeching condition is believed to be responsible for the increased spreading. Curiously, the jet from a 'lobed mixer' nozzle spreads much less at supersonic conditions compared to all other cases. This is due to the absence of screech with this nozzle. Jet spreading for the two tab configurations, on the other hand, is significantly more than any of the no-tab cases. This is true in the subsonic regime, as well as in the supersonic regime in spite of the fact that screech is essentially eliminated by the tabs. The dynamics of the streamwise vortex pairs produced by the tabs cause the most efficient jet spreading thus far observed in the study.

2 - Supersonic Beam Sources

Abstract: The supersonic expansions provide a means of preparing a molecular beam with a well-defined kinetic energy, which is particularly useful for crossed-beam and beam-surface scattering experiments This chapter discusses the supersonic expansions from circular nozzles The idealized continuum model assumes that collisions occur with sufficient frequency for equilibrium to be maintained throughout the expansion process This chapter also describes the interaction with background gases In any real apparatus, the idealized continuum model will also break down when the expanding gas becomes sufficiently rarefied that its density approaches that of the background gases This chapter also focuses on the supersonic expansions from slit nozzles (plane jets) In an idealized continuum slit supersonic expansion, it is still assumed that the gas flows without heat transfer to the walls and without viscosity, and an adiabatic, isentropic expansion is obtained

Spatial numerical simulations of linear and weakly nonlinear wave instabilities in supersonic boundary layers

Abstract: The spatial development of disturbances with small and moderate amplitudes in a two-dimensional (2-D) supersonic flat-plate boundary layer at Mach 48 is investigated using direct numerical simulations based on the compressible 3-D Navier-Stokes equations Disturbances are introduced into the boundary layer by blowing and suction within a narrow disturbance strip at the wall In response to the timewise periodic forcing, two types of disturbance waves are generated, a “first-mode” wave and a “multiple-viscous-solution” The “multiple-viscous-solution” was described by Mack (1969, 1984) but was not seen before in a direct numerical simulation The results of the simulations are compared with results of linear stability theory, and the agreement is very good In simulations for larger amplitudes, fundamental resonance is observed, where both types of 3-D waves are nonlinearly amplified and synchronize their phase velocities with the 2-D disturbance waves Subharmonic resonance is found for 3-D waves with large wave numbers, where the phase velocities of the linear 2-D and 3-D waves are nearly the same

Experimental study of the possibility of reducing supersonic drag by employing plasma technology

Abstract: The possibility of reducing the aerodynamic drag of a body by injecting plasma into the oncoming supersonic flow is confirmed experimentally.

Temperature and density mapping of supersonic jet expansions using linear Raman spectroscopy.

Abstract: Supersonic expansions associated to molecular gas jets are of considerable interest for various branches of physics, chemistry, and engineering. The properties of such rarefield environments differ markedly from those of a gas at thermodynamical equilibrium, and a complete diagnostic of their local properties is desired. The spatial distribution of molecular species, and its absolute density, the aggregation states (monomers, dimers, clusters, droplets, liquid, solid, aerosol, etc.), the translational speeds, the distribution of translational, rotational, and vibrational temperatures, and, in some cases, the anisotropic orientation of molecules along privileged directions are the main variables to be characterized. Spectroscopic methods can provide answers to these questions, but none of them can be regarded as a complete, versatile, diagnostic tool for the supersonic jet. Here we stress the use of linear Raman spectroscopy, specially improved and adapted for this purpose, as a highly competitive diagnosis method capable of providing a wealth of data about the jet. As shown below, the mapping of several basic physical quantities of a supersonic expansion by means of Raman spectroscopy is nowadays an affordable methodology, with specific advantages over other spectroscopic methods, namely, (a) the universality (all molecules can in principle be studied), (b) a high

Mixing of swirling jets in a supersonic duct flow

Abstract: Hydrogen fuel injected into a scramjet combustor must mix rapidly if complete combustion is to occur within a reasonable stream wise distance. An experiment has been conducted to determine whether the addition of swirl will improve the mixing of a supersonic jet of fuel simulant (helium or air) injected at 30 deg to the wall into a confined Mach 2 airflow. The swirling jets were created by injecting the fuel simulant tangentially into a cylindrical chamber and accelerating it through a convergent-div er gent nozzle. The flow was visualized by imaging Rayleigh scattering from a laser light sheet, and the plume penetration and cross-sectional area were obtained. The plumes from the swirling and nonswirling jets had comparable penetration and area, but the swirling jets contained substantially less mass flow, suggesting better mixing efficiency. Interaction of streamwise vorticity within the plumes of the swirling jets with their images in the duct wall caused the plumes to be inclined laterally to the freestream.

Multigrid acceleration of second‐order eno schemes from low subsonic to high supersonic flows

Abstract: SUMMARY In the present work the multigrid strategy is applied to second-order EN0 schemes for the computation of steady compressible flows. The performances of the algorithm are analysed in many flow situations, ranging from low subsonic to high supersonic flows, for both internal and external problems. Three different Riemann solvers were considered in the study of computational efficiency and solution accuracy.

Experimental studies of supersonic film cooling with shock wave interaction

Abstract: The supersonic film cooling was tested in the Mach 2.35 wind tunnel to investigate the effect of the external shock wave on the film cooling. The coolant was injected with sonic speed. The weak shock wave with the pressure ratio of 1.21 did not reduce the film cooling effectiveness. The stronger shock wave with the pressure ratio of 1.44 decreased the effectiveness of the film cooling in the restricted region. The decrease of the effectiveness was mainly the result of the increase of the adiabatic wall temperature by the decrease of the local Mach number. The increase of the heat transfer coefficient must be considered as well as that of the adiabatic wall temperature. In the region of the interaction, energy and mass were not transferred, but the momentum was transferred from the primary flow to the coolant.

Experimental investigation of supersonic gaseous injection into a supersonic freestream

Abstract: An experimental study of the mean and turbulent flowfield associated with low-angled supersonic gaseous injection into a supersonic freestream was performed. Air was injected at Mach 1.8, with an effective back pressure ratio of 3.0, through an orifice at an angle of 25 deg into a Mach 2.9 air freestream (Re/m - 15 x 10 6). Cross-film anemometry and conventional mean flow probe surveys were acquired across the plume at two downstream stations (xld = 20 and 40). Schlieren photography was used for qualitative flow visualization. Turbulence measurements included contours of the turbulent kinetic energy and the full compressible Reynolds shear stresses in both the x-y and x-z planes. Mean flow data included Mach number, three-dimensional velocity components, and vorticity. The measurements indicated that the mean and turbulent flow structure of the injection plume were strongly influenced by the presence of a counter-rotating vortex pair (\ux |max « 15,000 /s). The turbulent kinetic energy was found to have two peaks colocated with the vortices. The turbulent shear stress distributions across the plume were found to be highly three dimensional and complicated by both the additional strain rates associated with the vorticity and turbulent convection. The present results also implied that the compressibility terms in the Reynolds shear stress accounted for about 67.0-75.0% of the total shear stress level, i.e., up'v'lpu'v' and up'w'/pu'w' were in the range of 2.0-3.0.

Shock-wave-enhancement of the mixing and the stability limits of supersonic hydrogen-air jet flames

Abstract: A supersonic, non-premixed, jetlike flame was stabilized along the axis of a Mach-2.5 wind tunnel, and wedges were mounted on the sidewall in order to interact oblique shock waves with the flame. Schlieren photographs show how the interaction occurs, and measurements quantify how the flame length and the flame blowout limits are affected by the shocks. An optimum shock-interaction location was investigated by adjusting the wedge position. It was found that shock waves enhance the fuel-air mixing such that flame lengths decreased by 20% when an optimum shock location and shock strength were chosen. Enhanced mixing resulted, in part, because the shocks turn the flow and induce radial inflows of air into the fuel jet. A Mach disk sometimes occurs, which appears to split the reaction zone into two parts and severely distorts the flame shape. Substantial improvements in the flame stability (i.e., changes in the blowout limits) were achieved by properly interacting the shock waves with the flame-holding recirculation zone. The reason for the significant improvement in flame stability is believed to be the adverse pressure gradient caused by the shock, which can elongate the recirculation zone. Excessive shock strength (or poor shock placement) caused thermal choking to occur, and the flame base moved upstream of the fuel tube exit, leading to dangerously high wall heat transfer rates. Optimization of the mixing and stability limits requires a careful matching of the shock-flame interaction location and the shock strength. The experimental results show that the best mixing and stability correspond to 10° wedges placed at an upstream position (4 d F ) such that the primary shocks create radial inflow near the flame base and interact with the recirculation zone. This upstream wedge position also allowed the second (recompression) set of shocks to provide radial inflow near the flame tip.

Experimental investigation of side-jet steering for supersonic and hypersonic missiles

Abstract: An experimental investigation of the interaction between a side jet and the external flow is presented. The experiments were carried out in three different wind tunnels at supersonic and hypersonic Mach numbers ranging from 2 to 10. The model used in all the experiments consisted of a cylindrical body and various ogival nose sections, and had no lifting surfaces. The single sonic jet nozzle was located on the lee side of the cylindrical center body section with respect to a positive angle of attack. Several nozzle shapes and jet injection angles were examined. Forces and moments were measured directly in all the experiments, and surface pressure surveys were taken in some supersonic experiments. The results reveal a small jet force amplification at supersonic and hypersonic Mach numbers around zero angle of attack. Significant amplification because of positive angle of attack was noted at hypersonic speeds. The injection nozzle shape and vector angle can have a noticeable effect on the magnitude of the jet force amplification, especially at hypersonic speeds. In addition, the interaction gives rise to a significant pitching moment that may be utilized to produce an additional aerodynamic control force.

Hydrodynamical Models of Outflow Collimation in Young Stellar Objects

Abstract: In this paper we explore the physics of time-dependent hydrodynamic collimation of jets from young stellar objects (YSOs). Using parameters appropriate to YSOs, we have carried out high-resolution hydrodynamic simulations modeling the interaction of a central wind with an environment characterized by a toroidal density distribution which has a moderate opening angle of θρ 90°. The results show that for all but low values of the equator-to-pole density contrast the wind/environment interaction produces strongly collimated supersonic jets. The jet is composed of shocked wind gas. Using analytical models of wind-blown bubble evolution, we show that the scenario studied here should be applicable to YSOs and can, in principle, initiate collimation on the correct scales (R 100 AU). Comparison of our simulations with analytical models demonstrates that the evolution seen in the simulations is a mix of wind-blown bubble and jet dynamics. The simulations reveal a number of time-dependent nonlinear features not anticipated in previous analytical studies. These include: a prolate wind shock; a chimney of cold swept-up ambient material dragged into the bubble cavity; a plug of dense material between the jet and bow shocks. We find that the collimation of the jet occurs through both de Laval nozzles and focusing of the wind via the prolate wind shock. Using an analytical model for shock focusing we demonstrate that a prolate wind shock can, by itself, produce highly collimated supersonic jets.Animations from these simulations are available over the internet at http://www.msi.umn.edu/Projects/twj/jetcol.html.

Supersonic mechanisms for charge and energy transfers in anharmonic molecular chains

Abstract: The dynamical theory of a quantum quasiparticle moving in a deformable anharmonic chain is extended into the supersonic region. Besides a supersonic self-trapping mode, which is a direct extension of the well-known subsonic Davydov-Scott mode, two additional dynamically stable transfer mechanisms have been discovered in this region: (i) the capture and transfer of the self-trapping state by a supersonic acoustic (lattice) soliton and (ii) the pairing of two lattice solitons via their interaction with a quasiparticle. \textcopyright{} 1996 The American Physical Society.