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Showing papers on "Supersonic speed published in 1995"



Journal ArticleDOI
TL;DR: In this article, the mixing augmentation methods employed efficiently in sub- sonic flows failed to work at elevated Mach numbers, and some were inefficient because they were utilized outside their effective range.
Abstract: Recent interest in supersonic combustion (scramjets) and noise reduction for the high speed civil transport (HSCT) plane prompted renewed research in supersonic mixing processes and means to control them. The scramjet propulsion concept requires rapid mixing between fuel and air in order to minimize the size of the combustor and affect the performance of the entire vehicle system. Also, accelerated mixing of exhaust plumes with coflowing air has been shown to lead to jet noise reduction. Other examples of technological applications requiring control of mixing in compressible flows include thrust augmenting ejectors, thrust vector control, metal deposition, and gas dynamic lasers. The technological challenge of mixing enhancement in compressible flows stems from the inherently low growth rates of supersonic shear layers. Many mixing augmentation methods employed efficiently in sub­ sonic flows failed to work at elevated Mach numbers, and some were inefficient because they were utilized outside their effective range. Never-

316 citations


Journal ArticleDOI
06 Oct 1995-Science
TL;DR: In this paper, the authors measured the exhaust plume of a Concorde aircraft cruising at supersonic speeds in the stratosphere and found that the number of submicrometer particles in the plume implies efficient conversion of fuel sulfur to sulfuric acid in the engine or at emission.
Abstract: Emission indices of reactive gases and particles were determined from measurements in the exhaust plume of a Concorde aircraft cruising at supersonic speeds in the stratosphere. Values for NO x (sum of NO and NO 2 ) agree well with ground-based estimates. Measurements of NO x and HO x indicate a limited role for nitric acid in the plume. The large number of submicrometer particles measured implies efficient conversion of fuel sulfur to sulfuric acid in the engine or at emission. A new fleet of supersonic aircraft with similar particle emissions would significantly increase stratospheric aerosol surface areas and may increase ozone loss above that expected for NO x emissions alone.

184 citations



Journal ArticleDOI
TL;DR: In this article, the effects of the spike length, Mach number, and angle of attack on the supersonic flow were examined using three-dimensional thin-layer compressible Navier-Stokes equations.
Abstract: In supersonic flow, a spike attached to the nose reduces the drag of a blunt body. In this paper, supersonic flows around a spiked blunt body are numerically simulated to examine the effects of the spike length, Mach number, and angle of attack. Three-dimensional thin-layer compressible Navier-Stokes equations are solved using a highresolution upwind scheme with LU-ADI time-integration algorithm. The computed results show that the drag of the spiked blunt body is significantly influenced by the spike length, Mach number, and angle of attack. Scales of the separated region are not significantly influenced by the freestream Mach number. For the spiked blunt body at angle of attack, the flowfield becomes complex with spiral flows. The computed results are in reasonable agreement with experimental data.

101 citations


Journal ArticleDOI
TL;DR: In this paper, the authors present results from a recent gas + N-body simulation of a cluster merger, suggesting that mergers can result in long-lived, supersonic bulk flows, as well as shocks, within a few hundred kiloparsecs of the core of the dominant cluster.
Abstract: The intracluster medium (ICM) within merging clusters of galaxies is likely to be in a violent or turbulent dynamical state which may have a significant effect on the evolution of cluster radio sources. We present results from a recent gas + N-body simulation of a cluster merger, suggesting that mergers can result in long-lived, supersonic bulk flows, as well as shocks, within a few hundred kiloparsecs of the core of the dominant cluster. These results have motivated our new two-dimensional and three-dimensional simulations of jet propagation in such environments. The first set of simulations models the ISM/ICM transition as a contact discontinuity with a strong velocity shear. A supersonic (M(sub j) = 6) jet crossing this discontinuity into an ICM with a transverse, supersonic wind bends continuously, becomes 'naked' on the upwind side, and forms a distended cocoon on the downwind side. In the case of a mildly supersonic jet (M(sub j) = 3), however, a shock is driven into the ISM and ISM material is pulled along with the jet into the ICM. Instabilities excited at the ISM/ICM interface result in the jet repeatedly pinching off and reestablishing itself in a series of 'disconnection events.' The second set of simulations deals with a jet encountering a shock in the merging cluster environment. A series of relatively high-resolution two-dimensional calculations is used to confirm earlier analysis predicting that the jet will not disrupt when the jet Mach number is greater than the shock Mach number. A jet which survives the encounter with the shock will decrease in radius and disrupt shortly thereafter as a result of the growth of Kelvin-Helmholtz instabilities. We also find, in disagreement with predictions, that the jet flaring angle decreases with increasing jet density. Finally, a three-dimensional simulation of a jet crossing an oblique shock gives rise to a morphology which resembles a wide-angle tailed radio source with the jet flaring at the shock and disrupting to form a long, turbulent tail which is dragged downstream by the preshock wind.

95 citations


Journal ArticleDOI
TL;DR: The development of Doppler global velocimetry from its inception to its use as a flow diagnostics tool is described in this paper, which traces the evolution from an elementary one-component laboratory prototype, to a full three-component configuration operating in a wind tunnel at focal distances exceeding 15 m.
Abstract: The development of Doppler global velocimetry is described from its inception to its use as a flow diagnostics tool. Its evolution is traced from an elementary one-component laboratory prototype, to a full three-component configuration operating in a wind tunnel at focal distances exceeding 15 m. As part of the developmental process, several wind tunnel flow field investigations were conducted. These included supersonic flow measurements about an oblique shock, subsonic and supersonic measurements of the vortex flow above a delta wing, and three-component measurements of a high-speed jet.

95 citations


Journal ArticleDOI
TL;DR: The classification of the hydrodynamical growth mechanisms for the spherical bubbles of the low-temperature phase in cosmological phase transitions is completed by showing that the bubbles can grow as supersonic deflagrations.
Abstract: The classification of the hydrodynamical growth mechanisms for the spherical bubbles of the low-temperature phase in cosmological phase transitions is completed by showing that the bubbles can grow as supersonic deflagrations. Such deflagrations consist of a Jouguet deflagration, followed by a rarefaction wave. Depending on the amount of supercooling, the maximal velocity of supersonic deflagrations varies between the sound and the light velocities. The solutions faster than supersonic deflagrations are weak detonations.

94 citations


Journal ArticleDOI
TL;DR: In this paper, the axial force on the model was measured using a stress wave force balance, which is a recently developed technique for measuring forces in shock tunnels, and the results were found to compare satisfactorily with predictions based on established theoretical models, used with simplifying approximations.
Abstract: Measurements have been made of the propulsive effect of supersonic combustion ramjets incorporated into a simple axisymmetric model in a free piston shock tunnel. The nominal Mach number was 6, and the stagnation enthalpy varied from 2.8 to 8.5 MJ kg-1. A mixture of 13% silane and 87% hydrogen was used as fuel, and experiments were conducted at equivalence ratios up to approximately 0.8. The measurements involved the axial force on the model, and were made using a stress wave force balance, which is a recently developed technique for measuring forces in shock tunnels. A net thrust was experienced up to a stagnation enthalpy of 3.7 MJ kg-1, but as the stagnation enthalpy increased, an increasing net drag was recorded. Pitot and static pressure measurements showed that the combustion was supersonic. The results were found to compare satisfactorily with predictions based on established theoretical models, used with some simplifying approximations. The rapid reduction of net thrust with increasing stagnation enthalpy was seen to arise from increasing precombustion temperature, showing the need to control this variable if thrust performance was to be maintained over a range of stagnation enthalpies. Both the inviscid and viscous drag were seen to be relatively insensitive to stagnation enthalpy, with the combustion chambers making a particularly significant contribution to drag. The maximum fuel specific impulse achieved in the experiments was only 175 s, but the theory indicates that there is considerable scope for improvement on this through aerodynamic design.

88 citations


Journal ArticleDOI
TL;DR: A comprehensive review of experimental base pressure and base heating data related to supersonic and hypersonic flight vehicles is presented in this article, where a series of internally consistent, empirical predictions are developed for planar and axisymmetric geometries (wedges, cones, and cylinders).
Abstract: A comprehensive review of experimental base pressure and base heating data related to supersonic and hypersonic flight vehicles is presented. Particular attention is paid to free-flight data as well as wind-tunnel data for models without rear sting support. Using theoretically based correlation parameters, a series of internally consistent, empirical predictions are developed for planar and axisymmetric geometries (wedges, cones, and cylinders). These equations encompass the speed range from low supersonic to hypersonic flow and laminar and turbulent forebody boundary layers. A wide range of cone and wedge angles and cone bluntness ratios is included in the data base used to develop the correlations. The present investigation also includes an analysis of the effect of the angle of attack and the specific-heat ratio of the gas. Angle-of-attack effects are considered on sharp and blunted cones and cylindrical afterbodies.

76 citations


Journal ArticleDOI
TL;DR: In this paper, a series of flow visualizations has been performed on two flat-plate zero-pressure-gradient supersonic boundary layers and a number of new visualization techniques were applied.
Abstract: A series of flow visualizations has been performed on two flat-plate zero-pressure-gradient supersonic boundary layers. The two different boundary layers had moderate Mach numbers of 2.8 and 2.5 and Reθ's of 82, 000 and 25, 000 respectively. A number of new visualization techniques were applied. One was a variation of conventional schlieren employing “selective cut-off” at the knife edge plane. Motion pictures of the flow were generated with this technique. Droplet seeding was also used to mark the flow, and high speed movies were made to show structure evolution. Still pictures were also taken to show details within the large-scale motions. Finally, Rayleigh scattering was used to construct planar images of the flow. Together, these techniques provide detailed information regarding the character and kinematics of the large-scale motions appearing in boundary layers in supersonic flow. Using these data, in concert with existing hot-wire data, some suggestions are made regarding the characteristics of the “average” large-scale motion.

Journal ArticleDOI
TL;DR: In this paper, a swept-wing leading-edge model at Mach 3.5 is investigated and the experimental and computational results compare favorably in most cases, and suggest that transition is probably dominated by traveling, rather than stationary, crossflow disturbances for the present model.
Abstract: Transition on a swept-wing leading-edge model at Mach 3.5 is investigated. Surface pressure and temperature measurements are obtained in the NASA Langley Research Center Supersonic Low-Disturbance Tunnel. For one case, temperature-sensitive paint and a sublimating chemical are used to visualize surface flow features such as transition location. The experimental data are compared with 1) mean-flow results computed as solutions to the thin-layer Navier-Stokes equations and 2) N-factors obtained using the envelope e N method. The experimental and computational results compare favorably in most cases. In particular, N ≃ 13 correlates best with the observed transition location over a range of freestream unit Reynolds numbers and angles of attack. Computed traveling crossflow disturbances with frequencies of 40-60 kHz have the largest N factors, and the surface flow visualizations reveal smooth transition fronts with only faint evidence of stationary crossflow vortices. These results suggest that transition is probably dominated by traveling, rather than stationary, crossflow disturbances for the present model.

Journal ArticleDOI
TL;DR: In this article, the supersonic acoustic intensity vector is defined for measurements on plate and cylinder-like structures, which is composed only of wave components which radiate to the far field (supersonic), with the nonradiating (subsonic) components eliminated.
Abstract: A new acoustic quantity called the supersonic acoustic intensity vector is defined in this paper for application to measurements on plate and cylinderlike structures. As the name implies, the supersonic intensity is composed only of wave components which radiate to the far field (supersonic), with the nonradiating (subsonic) components eliminated. The normal component of this vector, measured in the extreme near field or on the surface of the structure, provides an accurate tool for locating regions (‘‘hot spots’’) on the structure which radiate to the farfield. Furthermore, the supersonic intensity provides an accurate quantification of these source regions, providing a ranking of the strength of the identified source regions as a function of frequency. This identification and ranking provides a powerful new tool in the understanding and control of radiated noise.

Patent
25 Jan 1995
TL;DR: In this paper, a supersonic airplane with four or two SUVs and one or more boost engines is described, where the SUVs are operated at a lower thrust setting within acceptable noise limits, and the subsonic engines are operated to provide boost thrust to enable the airplane to operate a take-off and climb.
Abstract: A supersonic airplane having four or two supersonic engines and one or more boost engines. During take-off and initial climb the supersonic engines are operated at a lower thrust setting within acceptable noise limits, and the subsonic engine(s) is operated to provide boost thrust to enable the airplane to operate a take-off and climb. During cruise, the subsonic engines are in a nonoperating mode, and the supersonic engines alone provide the thrust for supersonic operation. In one embodiment, one subsonic engine is deployed on one side of the fuselage during the operating mode. In a second embodiment, there are two subsonic engines deployed on opposite sides of the fuselage. In another embodiment, a single subsonic engine is installed inside the fuselage.

Proceedings ArticleDOI
01 Jan 1995
TL;DR: In this article, a flux splitting and limiting technique which yields one-point stationary shock capturing is presented, which is applied to the full NavierStokes and Reynolds Averaged Navier-Stokes equations.
Abstract: A new flux splitting and limiting technique which yields one-point stationary shock capturing is presented. The technique is applied to the full NavierStokes and Reynolds Averaged Navier-Stokes equations. Calculations of laminar boundary layers at subsonic and supersonic speeds are presented together with calculations of transonic flows around airfoils. The results exhibit very good agreement with theoretical solutions and existing experimental data. It is found that. the proposed scheme improves the resolution of viscous flows while maintaining excellent one-point shock capturing characteristics. d

Journal ArticleDOI
TL;DR: In this article, the authors used a strut divided streamwise into two parts, and hydrogen gas was injected into the interval between the two parts of the strut to stabilize the combustion region.
Abstract: Flame stabilization experiments were conducted in a supersonic airflow of Mach number 1.5, using a strut divided streamwise into two parts. Hydrogen gas was injected into the interval between the two parts of the strut. The flame stabilization was definitely affected by whether the combustion region could be established in this space, and the flame stabilization characteristics changed drastically according to the distance between the two parts of the strut. A shadowgraph and schlieren photographs showed that no shock waves or expansion waves existed in the intervening space, and that waves did not directly control the flame-stabilization mechanism that was altered by the distance between two parts. In order to explain the present characteristic flamestabilization from the standpoint of the competition between mass transfer and reaction, the velocity fields were measured by laser Doppler velocimeter, and the residence times in the intervening space were estimated. Through these observations and measurements, the flame-stabilization mechanism was clarified and the usefulness of this type of strut was demonstrated.


Journal ArticleDOI
TL;DR: In this article, the effects of layering are included by use of an exact factorization of the displacement and stress fields in terms of generalized transmission and reflection coefficients in a multilayered viscoelastic half-space.

Journal ArticleDOI
R. F. Tate1, B.S. Hunt1, C.A. Helms1, K.A. Kruesdell1, G.D. Hager1 
TL;DR: In this paper, the spatial distribution of small signal gain has been investigated on the RADICL device, a supersonic chemical oxygen-iodine laser (COIL), operating on the F=3/spl rarr/F=4 hyperfine levels of the spin-orbit transition in atomic iodine was used as a small signal probe.
Abstract: The spatial distribution of small signal gain has been investigated on the RADICL device, a supersonic chemical oxygen-iodine laser (COIL). A frequency-stabilized, narrow linewidth diode laser system operating on the F=3/spl rarr/F=4 hyperfine levels of the (/sup 2/P/sub 1/2/) to (/sup 2/P/sub 3/2/) spin-orbit transition in atomic iodine was used as a small signal probe. A peak gain of 1.2%/cm was measured along the horizontal centerline of the single-slit, supersonic nozzle, which is about two times greater than measurements made on ReCOIL by Hager et al. (1988) and compares favorably with measurements made on the RotoCOIL device by Keating et al. (1990). Gain distribution was investigated under three I/sub 2/ flow conditions. Scans across the supersonic expansion indicate a gradient in gain distribution due to higher gas temperatures along the walls and mixing phenomena. >

PatentDOI
Kimihiro Kishi1
TL;DR: In this article, a sound absorbing member made of a porous plate (4) and a honeycomb structure (2) was provided between a nozzle plate and a liner to reduce noise and to cool a nozzle in a supersonic jet propelling engine.
Abstract: To reduce a noise and to cool a nozzle in a supersonic jet propelling engine, a sound absorbing member made of a porous plate (4) and a honeycomb structure (2) or a porous plate (4) and a torus core (3) is provided between a nozzle plate (5) and a liner (1), a cooling air (7) is caused to flow along an inner surface of the nozzle plate (5) and to impinge against the liner (1) through holes (10) of the porous plate (4). Furthermore, a number of holes (11) which are slanted on a rear side are formed in the liner (1) so that the cooling air (9) are injected through the holes (11).

Journal ArticleDOI
TL;DR: In this article, a fourth-order shock-capturing scheme was developed for steady-state and time-accurate simulation of chemically reacting flows with finite rate chemical kinetics, where a diagonal algorithm and a local implicit technique were presented to remove the stiffness of chemical reaction and to achieve high computation efficiency.
Abstract: A fourth-order shock-capturing scheme has been developed for steady-state and time-accurate simulation of chemically reacting flows with finite rate chemical kinetics. A diagonal algorithm and a local implicit technique are presented to remove the stiffness of chemical reaction and to achieve high computation efficiency. A fully implicit code is obtained by combining the present algorithm with the lower-upper scheme. The validity of this code is demonstrated by calculating the unsteady shocks in an inviscid supersonic duct and by comparing the results with the previous calculation. Then the code is used to calculate the ignition of a premixed hydrogen/air reacting flow in a ramped duct and nonpremixed hydrogen/air streams as well as nonpremixed methane/air streams in a supersonic mixing layer. The efficiency and robustness of the present scheme are shown through these numerical simulations.

Journal ArticleDOI
TL;DR: In this article, an experimental study of supersonic wing tip vortices has been conducted at Mach 2.49 using small-scale four-hole and five-hole conical probes.
Abstract: An experimental study of supersonic wing tip vortices has been conducted at Mach 2.49 using small-scale four-hole and five-hole conical probes, The study was performed 2.25 chords downstream of a semispan rectangular wing at angles of attack of 5.7 and 10.4 deg. The main objective of the experiments was to determine the Mach number, flow angularity and total pressure distribution in the core region of supersonic wing tip vortices. A secondary aim was to demonstrate the feasibility of calibrating a conical probe using a computational solution to predict the flow characteristics. Results of the present investigation showed that the numerically generated calibration data can be used for pointed nose four-hole conical probes but were not sufficiently accurate for conventional five-hole probes due to nose bluntness effects. A combination of four-hole conical probe measurements with independent pitot pressure measurements indicated a significant Mach number and total pressure deficit in the core regions of supersonic wing tip vortices, combined with an asymmetric ''Burger-like'' swirl distribution.

01 Jul 1995
TL;DR: In this article, the authors describe the implementation of optimization techniques based on control theory for wing and wing-body design of supersonic configurations, which represents an extension of earlier research in which control theory is used to devise a design procedure that significantly reduces the computational cost by employing an adjoint equation.
Abstract: This paper describes the implementation of optimization techniques based on control theory for wing and wing-body design of supersonic configurations. The work represents an extension of our earlier research in which control theory is used to devise a design procedure that significantly reduces the computational cost by employing an adjoint equation. In previous studies it was shown that control theory could be used to~eviseransonic design methods for airfoils and wings in which the shape and the surrounding body-fitted mesh are both generated analytically, and the control is the mapping function. The method has also been implemented for both transonic potential flows and transonic flows governed by the Euler equations using an alternative formulation which employs numerically generated grids, so that it can treat more general configurations. Here results are presented for three-dimensional design cases subject to supersonic flows governed by the Euler equation.

Journal ArticleDOI
TL;DR: In this article, the authors defined the scale of turbulence in the planetary boundary layer as a function of wind speed and surface roughness and the lapse rate and mixing layer thickness of the boundary layer turbulence.
Abstract: Nomenclature a = lift curve slope, per rad #„, bn — Fourier series coefficients for e.g. acceleration c = wing mean geometric chord, ft FK — flight profile alleviation factor FKm = flight profile factor due to airplane weight FKZ = flight profile factor due to altitude g = acceleration due to gravity, 32.2 ft/s g;j, hn = Fourier series coefficients for true gust velocity H = gust gradient distance, ft K, KK = gust load alleviation factors L = scale of turbulence, ft M = Mach number q = dynamic pressure, psf /?, = maximum landing weight/maximum takeoff weight R2 = maximum zero fuel weight/maximum takeoff weight S = wing area, ft s = distance along flight path, ft T = local atmospheric temperature, °R 7,,, /„ = real and imaginary parts of the e.g. acceleration transfer function T() = ambient atmospheric temperature, °R U = gust velocity, fps, true airspeed (7dc = derived equivalent gust velocity, fps, equivalent airspeed (7ds = design gust velocity, fps, equivalent airspeed £/dt = derived true gust velocity, fps, true airspeed £/rct = design reference gust velocity, fps, equivalent airspeed Utr = power spectral scale factor V = aircraft velocity, fps, true airspeed VB = rough air penetration speed, Kt, equivalent airspeed Vc = design cruise speed, Kt, equivalent airspeed VD = design dive speed, Kt, equivalent airspeed W = aircraft weight, Ib Zmo = maximum operating altitude, ft y = ratio of specific heats Aft = incremental load factor, g fjif, = airplane mass parameter p = air density, slugs/ft Introduction S UBSONIC aircraft respond to atmospheric turbulent air motions or eddies of, approximately, 30-2000 ft in extent. Smaller eddies will generally be averaged out over the surface of the aircraft, larger eddies typically will not cause sharp or excessive aircraft accelerations or structural loads on the airplane. Aircraft in supersonic flight respond to ever longer wavelengths as the flight speed is increased. From the aircraft design standpoint, atmospheric turbulence may be separated into two categories: 1) turbulence, which contributes to aircraft structural fatigue and passenger inconvenience and discomfort, is generally associated with the less intense, smaller scales (small eddy size or higher spatial frequencies as characterized by the turbulence power spectrum); and 2) turbulence that can cause aircraft upset, passenger injury, and possibly structural damage or failure is associated with the more intense larger scales of the turbulence spectrum. Whereas the first category may be in the inertial subrange of the turbulence power spectrum where local homogeneity and stationarity may be assumed, the second category is definitely associated with the larger energy-containing scales of turbulence that are definitely nonstationary and inhomogeneous. In fact, the second type of atmospheric turbulence may not be turbulence at all, but may be a part of, or derive directly from, the ordered convective or geostrophic motions of the atmosphere. Frictionally induced turbulence in the planetary boundary layer is dependent on wind speed near the ground and the surface roughness. Convective turbulence in the planetary boundary layer is dependent on the lapse rate (the rate of temperature change with altitude) and the depth of the mixing layer. Thus, the wind speed and surface roughness and the lapse rate and mixing layer thickness largely determine the intensity and scale of the boundary-layer turbulence. This low altitude boundary-layer turbulence will affect landing and takeoff operations for the larger commercial aircraft that normally cruise at high altitude, and it will be the primary cause of disturbance for small aircraft that operate at low altitudes. Usually, it can be characterized as random, locally stationary, and Gaussian, and the turbulence scale at the surface is of the order of 1000 ft and increases with altitude. It is only natural that pilots avoid flying into areas of obviously rough air such as thunderstorms or even cumulus clouds;

Proceedings ArticleDOI
09 Jan 1995

Journal ArticleDOI
TL;DR: In this article, it was shown that above a second and larger critical ramp angle, the boundary-layer flow develops an instability associated with the occurrence of inflection points in the streamwise velocity profiles within the recirculation region and develops as a wave packet which remains stationary near the corner.
Abstract: Separation of a supersonic boundary layer (or equivalently a hypersonic boundary layer in a region of weak global interaction) near a compression ramp is considered for moderate wall temperatures. For small ramp angles, the flow in the vicinity of the ramp is described by the classical supersonic triple-deck structure governing a local viscous-inviscid interaction. The boundary layer is known to exhibit recirculating flow near the corner once the ramp angle exceeds a certain critical value. Here it is shown that above a second and larger critical ramp angle, the boundary-layer flow develops an instability. The instability appears to be associated with the occurrence of inflection points in the streamwise velocity profiles within the recirculation region and develops as a wave packet which remains stationary near the corner and grows in amplitude with time.

Patent
08 May 1995
TL;DR: A gas-liquid cleaning spray system employs one or more converging-diverging nozzles to accelerate a gas liquid mixture to a supersonic velocity for cleaning various types of articles, such as mechanical, electrical and fluid components as mentioned in this paper.
Abstract: A gas-liquid cleaning spray system employs one or more converging-diverging nozzles to accelerate a gas-liquid mixture to a supersonic velocity for cleaning various types of articles, such as mechanical, electrical and fluid components. The gas, such as air or nitrogen, is supplied at high pressure to a nozzle body where it is mixed with cleaning liquid, such as water or liquid detergent, which is supplied to the nozzle body at a relatively low flow rate. Acceleration of the gas-liquid mixture to a supersonic velocity eliminates the need for a high pressure, high flow rate and high volume liquid supply. After the components are contacted with the gas-liquid mixture, the cleaning liquid can be recaptured and analyzed for cleanliness verification of the components.


Journal ArticleDOI
TL;DR: In this paper, tangential injection is an efficient method for generating swirling jets, and the swirling jets mix much more rapidly with the stagnant air than comparable straight jets, when overexpanded, turbulence is created in the jet core as a result of vortex breakdown.
Abstract: The addition of swirl to scramjet fuel jets has been proposed as a method of enhancing fuel mixing, but little of a fundamental nature is known about supersonic swirling flows. Several jets with different amounts of swirl were created by tangential injection and acceleration through a convergent-divergent nozzle. The flowfields near the nozzle exit were investigated using pitot, cone, and total-temperature probes, and Rayleigh scattering from a laser light sheet. The results show that tangential injection is an efficient method for generating swirling jets, that the swirling jets mix much more rapidly with the stagnant air than comparable straight jets, and that, when overexpanded, turbulence is created in the jet core as a result of vortex breakdown. Mixing layer growth rates are shown to correlate with Richardson number