Showing papers on "Mach wave published in 2013"

Ship Wakes: Kelvin or Mach Angle?

Abstract: From the analysis of a set of airborne images of ship wakes, we show that the wake angles decrease as U(-1) at large velocities, in a way similar to the Mach cone for supersonic airplanes. This previously unnoticed Mach-like regime is in contradiction with the celebrated Kelvin prediction of a constant angle of 19.47° independent of the ship's speed. We propose here a model, confirmed by numerical simulations, in which the finite size of the disturbance explains this transition between the Kelvin and Mach regimes at a Froude number Fr=U/√[gL]~/=0.5, where L is the hull ship length.

A scaling analysis for turbulent shock-wave/boundary-layer interactions

Abstract: A model based on mass conservation properties is developed for shock-wave/boundarylayer interactions (SWBLIs), aimed at reconciling the observed great diversity in flow organization documented in the literature, induced by variations in interaction geometry and aerodynamic conditions. It is the basis for a scaling approach for the interaction length that is valid independent of the geometry of the flow (considering compression corners and incident-reflecting shock interactions). As part of the analysis, a scaling argument is proposed for the imposed pressure jump that depends principally on the free-stream Mach number and the flow deflection angle. Its interpretation as a separation criterion leads to a successful classification of the separation states for turbulent SWBLIs (attached, incipient or separated). In addition, the dependence of the interaction length on the Reynolds number and the Mach numbers is accounted for. A large compilation of available data provides support for the validity of the model. Some general properties on the state of the flow are derived, independent of the geometry of the flow and for a wide range of Mach numbers and Reynolds numbers.

Numerical investigation of shock wave reflections near the head ends of rotating detonation engines

Abstract: The influence of various chamber geometries on shock wave reflections near the head end of rotating detonation engines was investigated. A hydrogen/air one-step chemical reaction model was used. The results demonstrated that the variation in flow field along the radial direction was not obvious when the chamber width was small, but became progressively more obvious as the chamber width increased. The thrust increased linearly, and the detonation height and the fuel-based gross specific impulse were almost constant as the chamber width increased. Near the head end, shock waves reflected repeatedly between the inner and outer walls. Both regular and Mach reflections were found near the head end. The length of the Mach stem increased as the chamber length increased. When the chamber width, chamber length and injection parameters were the same, the larger inner radius resulted in more shock wave reflections between the inner and outer walls. The greater the ratio of the chamber width to the inner radius, the weaker the shock wave reflection near the head end. The detonation height on the outer wall and the thrust, both increased correspondingly, while the specific impulse was almost constant as the inner radius of the chamber increased. The numerical shock wave reflection phenomena coincided qualitatively with the experimental results.

On the interaction of shock waves and sound waves in transonic buffet flow

Abstract: To support Lee's buffet mechanism model [B. H. K. Lee, “Self-sustained shock oscillations on airfoils at transonic speeds,” Prog. Aerosp. Sci. 37, 147–196 (2001)10.1016/S0376-0421(01)00003-3], the sound wave propagation in the flow field outside the separation of a transonic buffet flow at a Mach number M∞ = 0.73 and an angle of attack α = 3.5° over a DRA 2303 supercritical airfoil is determined using high-speed particle-image velocimetry. Furthermore, the shock wave is influenced by an artificial sound source which evidently changes the shock oscillation properties. The dominant buffet mechanism is shown to be a feedback loop between the shock position and the noise generation at the trailing edge of the airfoil. The sound wave propagation speed is detected by correlating the surface pressure signals and the velocity fluctuations in the flow field. The quantitative results for the natural and the artificial sound source convincingly coincide and are in good agreement with a reformulated version of Lee's ...

Detonation re-initiation mechanism following the Mach reflection of a quenched detonation

Abstract: This experimental study addresses the re-initiation mechanism of detonation waves following the Mach reflection of a shock–flame complex. The detonation diffraction around a cylinder is used to reproducibly generate the shock–flame complex of interest. The experiments are performed in methane–oxygen. We use a novel experimental technique of coupling a two-in-line-spark flash system with a double-frame camera in order to obtain microsecond time resolution permitting accurate schlieren velocimetry. The first series of experiments compares the non-reactive sequence of shock reflections with the reflection over a rough wall under identical conditions. It was found that the hot reaction products generated along the rough wall are entrained by the wall jet into a large vortex structure behind the Mach stem. The second series of experiments performed in more sensitive mixtures addressed the sequence of events leading to the detonation establishment along the Mach and transverse waves. Following ignition and jet entrainment, a detonation first appears along the Mach stem while the transverse wave remains non-reactive. The structure of the unburned tongue however indicates local instabilities and hot spot formation, leading to the rapid reaction of this gas. Numerical simulations are also reported, confirming the sequence of ignition events obtained experimentally.

Incident shock Mach number effects on Richtmyer-Meshkov mixing in a heavy gas layer

Abstract: Experiments were performed at the horizontal shock tube facility at Los Alamos National Laboratory to study the effect of incident shock Mach number (M) on the development of Richtmyer-Meshkov instability after a shock wave impulsively accelerates a varicose-perturbed, heavy-gas curtain. Three cases of incident shock strength were experimentally investigated: M = 1.21, 1.36, and 1.50. We discuss the state of the mixing and the mechanisms that drive the mixing at both large and small scales by examining the time evolution of 2D density fields derived from quantitative planar laser-induced fluorescence measurements. Several differences in qualitative flow features are identified as a result of Mach number variation, and differences in vortex interaction, observed using particle image velocimetry, play a critical role in the development of the flow field. Several quantities, including mixing layer width, mixing layer area, interface length, instantaneous mixing rate, the density self-correlation parameter, probability density functions of the density field, and mixing progress variables are examined as a function of time. These quantities are also examined versus time scaled with the convection velocity of the mixing layer. A higher incident Mach number yields greater mixing uniformity at a given downstream location, while a lower Mach number produces a greater amount of total mixing between the two gases, suggesting possible implications for optimization in applications with confined geometries.

The attenuation of sound by turbulence in internal flows

Abstract: The attenuation of sound waves due to interaction with low Mach number turbulent boundary layers in internal flows (channel or pipe flow) is examined. Dynamic equations for the turbulent Reynolds stress on the sound wave are derived, and the analytical solution to the equation provides a frequency dependent eddy viscosity model. This model is used to predict the attenuation of sound propagating in fully developed turbulent pipe flow. The predictions are shown to compare well with the experimental data. The proposed dynamic equation shows that the turbulence behaves like a viscoelastic fluid in the interaction process, and that the ratio of turbulent relaxation time near the wall and the sound wave period is the parameter that controls the characteristics of the attenuation induced by the turbulent flow.

Investigation of the Flow Structures in Supersonic Free and Impinging Jet Flows

Abstract: A numerical study of compressible jet flows is carried out using Reynolds averaged Navier-Stokes (RANS) turbulence models such as k-E and k-ω-SST. An experimental investigation is performed concurrently using high-speed optical methods such as Schlieren photography and shadowgraphy. Numerical and experimental studies are carried out for the compressible impinging at various impinging angles and nozzle-to-wall distances. The results from both investigations converge remarkably well and agree with experimental data from the open literature. From the flow visualizations of the velocity fields, the RANS simulations accurately model the shock structures within the core jet region. The first shock cell is found to be constraint due to the interaction with the bow-shock structure for nozzle-to-wall distance less than 1.5 nozzle diameter. The results from the current study show that the RANS models utilized are suitable to simulate compressible free jets and impinging jet flows with varying impinging angles. © 2013 by ASME.

A Critical Shock Mach Number for Particle Acceleration in the Absence of Pre-existing Cosmic Rays: M = √5

Abstract: It is shown that, under some generic assumptions, shocks cannot accelerate particles unless the overall shock Mach number exceeds a critical value M > √5. The reason is that for M ≤ √5 the work done to compress the flow in a particle precursor requires more enthalpy flux than the system can sustain. This lower limit applies to situations without significant magnetic field pressure. In case that the magnetic field pressure dominates the pressure in the unshocked medium, i.e., for low plasma beta, the resistivity of the magnetic field makes it even more difficult to fulfill the energetic requirements for the formation of shock with an accelerated particle precursor and associated compression of the upstream plasma. We illustrate the effects of magnetic fields for the extreme situation of a purely perpendicular magnetic field configuration with plasma beta β = 0, which gives a minimum Mach number of M = 5/2. The situation becomes more complex, if we incorporate the effects of pre-existing cosmic rays, indicating that the additional degree of freedom allows for less strict Mach number limits on acceleration. We discuss the implications of this result for low Mach number shock acceleration as found in solar system shocks, and shocks in clusters of galaxies.

Caveats for Using Shock Tube in Blast-Induced Traumatic Brain Injury Research

Abstract: Blast-induced traumatic brain injury (TBI) is currently an important and very “hot” research topic because it has been acknowledged to be a significant source of morbidity and disability during the wars in Iraq and Afghanistan, among blast victims. A total of 545 academic articles about blast TBI research have been published since 1946, of which 82% (447 articles) have been published since 2003, and 57% (312 articles) were published from 2010 to 2013. A number of experimental models are currently implemented to investigate the mechanisms of blast-induced TBI in rodents and larger animals such as rabbits and swine. As the fundamental shock wave generator, shock tubes (either compressed air-driven or detonation-driven) are generally employed in these experimental models. The compressed air-driven shock tube is a horizontally mounted, circular steel tube, in which a gas at low pressure (the driven gas) and a gas at high pressure (the driver gas) are separated using diaphragms (such as polyester Mylar membrane). After the diaphragm suddenly ruptures at predetermined pressure thresholds (e.g., 126–147 kPa), shock waves are generated and propagate through the low pressure section (the driven section) toward the mouth of the shock tube. The detonation-driven shock tube is a cylindrical metal tube that is closed at one end. The blast, causing the shock waves, is generated by detonation of an explosive charge in the closed end of the tube. Both compressed air-driven and detonation-driven shock tubes can produce blast shock waves to induce blast injuries in animals. However, because of their designs and structures, both shock tubes are not able to generate the Friedlander wave (an ideal form of a primary blast wave) that occurs when a powerful explosive detonates in a free field, without nearby surfaces that can interact with the wave. A series of complex shock waves are then generated following the lead shock wave (the original shock front), including reflected shock waves, a Mach stem, an unsteady turbulent jet, and rarefaction waves. These waves can cause sudden compression or rarefaction effects upon any object encountered in their motion path, and transfer kinetic energy to the object. Therefore, if an experimental animal is placed inside the shock tube, these complex pressure waves will cause more severe and complex injuries that are rarely observed in blast victims, thus leading to false-positive results in the studies of blast TBI mechanism.

Studies on the influence of steady microactuators on shock-wave/boundary layer interaction

Abstract: An array of high-momentum microjets are used upstream of a compression corner to control the shock-wave/boundary-layer interaction on a 24 deg unswept compression ramp in a Mach 2 flow. Measurements include schlieren flow visualization and unsteady pressure measurements using fast-response pressure sensors of the interaction region. Results show that the array of microjets issuing in the supersonic crossflow create oblique shocks, which effectively reduce the incoming Mach number at the compression corner. This leads to a modified separation shock of significantly reduced strength. The location of the modified shock is moved upstream by as much as 4δ0 from its mean undisturbed location. The mean pressure distribution on the surface is altered with microjet control leading to a more gradual compression of the incoming flow relative to the separation shock without control. The wall-pressure fluctuations in the interaction region are reduced by approximately 50%, and the flow near the compression corner appe...

The non-stationary hysteresis phenomenon in shock wave reflections

Abstract: The non-stationary transition from Mach to regular reflection followed by a reversetransition from regular to Mach reflection is investigated experimentally. A newexperimental setup in which an incident shock wave reflects from a cylindricalconcave surface followed by a cylindrical convex surface of the same radius isintroduced. Unlike other studies that indicate problems in identifying the triple point,an in-house image processing program, which enables automatic detection of the triplepoint, is developed and presented. The experiments are performed in air having aspecific heats ratio 1.4 at three different incident-shock-wave Mach numbers: 1.2, 1.3and 1.4. The data are extracted from high-resolution schlieren images obtained bymeans of a fully automatically operated shock-tube system. Each experiment producesa single image. However, the high accuracy and repeatability of the control systemtogether with the fast opening valve enables us to monitor the dynamic evolution ofthe shock reflections. Consequently, high-resolution results both in space and time areobtained. The credibility of the present analysis is demonstrated by comparing the firsttransition from Mach to regular reflection (MR !RR) with previous single cylindricalconcave surface experiments. It is found that the second transition, back to Machreflection (RR !MR), occurs earlier than one would expect when the shock reflectsfrom a single cylindrical convex surface. Furthermore, the hysteresis is observed atincident-shock-wave Mach numbers smaller than those at which the dual-solutiondomain starts, which is the minimal value for obtaining hysteresis in steady andpseudo-steady flows. The existence of a non-stationary hysteresis phenomenon, whichis different from the steady-state hysteresis phenomenon, is discovered.Key words: compressible flows, shock waves

High-Speed Schlieren Imaging of Disturbances in a Transitional Hypersonic Boundary Layer.

Abstract: A high-speed schlieren system was developed for the Sandia Hypersonic Wind Tunnel. Schlieren images were captured at 290 kHz and used to study the growth and breakdown of second-mode instabilities into turbulent spots on a 7 ◦ cone. At Mach 5, wave packets would intermittently occur and break down into isolated turbulent spots surrounded by an otherwise smooth, laminar boundary layer. At Mach 8, the boundary layer was dominated by second-mode instabilities which would break down into larger regions of turbulence. Second-mode waves surrounded these turbulent patches as opposed to the smooth laminar flow seen at Mach 5. Detailed pressure and thermocouple measurements were also made along the cone at Mach 5, 8 and 14, in a separate tunnel entry. These measurements give an average picture of the transition behavior that complements the intermittent behavior captured by the schlieren system. At Mach 14, the boundary-layer remained laminar so the transition process could not be studied. However, the first measurements of second-mode waves were made in HWT-14.

Aero-Optical Mitigation of Shocks Around Turrets at Transonic Speeds Using Passive Flow Control

Abstract: Experimental studies of the aero-optical effects around a partially-protruding cylindrical turret for a range of incoming transonic Mach numbers are presented and discussed. Spatially-temporally resolved wavefronts were collected using a high-speed Shack- Hartmann sensor and flow visualization was performed with a Schlieren system. Different flow regimes with a local shock, either a steady or an unsteady one, were described for the baseline case and the shock dynamics was found to be sensitive to a local flow speed. In addition, several passive flow control devices, consisted of a single spanwise row of vertically-placed small-diameter pins or porous screens, were tested in order to mitigate detrimental unsteady-shock-related aero-optical effects. It was found that passive flow control devices with large blockage values slowed the flow near the cylinder surface down to subsonic speeds by introducing total pressure losses in the wall region upstream of the cylinder, thus eliminating the shock formation over a wide range of transonic Mach numbers and significantly improving aero-optical environment at some elevation angles.

Numerical Study of Pressure Fluctuations due to a Mach 6 Turbulent Boundary Layer

Abstract: Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by a Mach 6 turbulent boundary layer with nominal freestream Mach number of 6 and Reynolds number of Re(sub t) approx. =. 464. The emphasis is on comparing the primarily vortical pressure signal at the wall with the acoustic freestream signal under higher Mach number conditions. Moreover, the Mach-number dependence of pressure signals is demonstrated by comparing the current results with those of a supersonic boundary layer at Mach 2.5 and Re(sub t) approx. = 510. It is found that the freestream pressure intensity exhibits a strong Mach number dependence, irrespective of whether it is normalized by the mean wall shear stress or by the mean pressure, with the normalized fluctuation amplitude being significantly larger for the Mach 6 case. Spectral analysis shows that both the wall and freestream pressure fluctuations of the Mach 6 boundary layer have enhanced energy content at high frequencies, with the peak of the premultiplied frequency spectrum of freestream pressure fluctuations being at a frequency of omega(delta)/U(sub infinity) approx. = 3.1, which is more than twice the corresponding frequency in the Mach 2.5 case. The space-time correlations indicate that the pressure-carrying eddies for the higher Mach number case are of smaller size, less elongated in the spanwise direction, and convect with higher convection speeds relative to the Mach 2.5 case. The demonstrated Mach-number dependence of the pressure field, including radiation intensity, directionality, and convection speed, is consistent with the trend exhibited in experimental data and can be qualitatively explained by the notion of "eddy Mach wave" radiation.

Effects of Inflow Mach Number and Step Height on Supersonic Flows over a Backward-Facing Step

Abstract: The backward-facing step is practically implicated in many devices, encountering the massive separation flows. In the present study, simulations of supersonic flow over a backward-facing step have been carried out employing both RANS and LES. The simulated results are validated against the experimental data. The results of RANS and LES show a good comparison with the experimental results. Different inflow Mach numbers and expansion ratios are also investigated. The reattachment length decreases with the increase of inflow Mach number. The duct height has a great effect on the flow patterns. The present conclusions are helpful to understand the physics in supersonic separation flows and also provide theory basis for engineering applications.

Vortex dynamics in 3D shock-bubble interaction

Abstract: The dynamics of shock-bubble interaction involve an interplay of vortex stretching, dilation, and baroclinic vorticity generation. Here, we quantify the interplay of these contributions through high resolution 3D simulations for several Mach and Atwood numbers. We present a volume rendering of density and vorticity magnitude fields of shock-bubble interaction at M = 3 and air/helium density ratio η = 7.25 to elucidate the evolution of the flow structures. We distinguish the vorticity growth rates due to baroclinicity, stretching, and dilatation at low and high Mach numbers as well as the late time evolution of the circulation. The results demonstrate that a number of analytical models need to be revised in order to predict the late time circulation of shock-bubble interactions at high Mach numbers. To this effect, we propose a simple model for the dependence of the circulation to Mach number and ambient to bubble density ratio for air/helium shock-bubble interactions.

Plane and Three-Dimensional Flow at High Subsonic Speeds

Abstract: For two- and three-dimensional flow in a compressible medium, a simple relation is given by which, to a first approximation, the quantitative influence of compressibility upon the velocities and pressures can be understood in a clear manner. In the application of this relation the distinct behaviors of two-dimensional and axially symmetric three-dimensional flow with increasing Mach number are brought out. For slender elliptic cylinders and ellipsoids of revolution, calculations are made of the critical Mach number; that is, the Mach number at which local sonic velocity is achieved on the body. As a further example, the lifting wing of finite span is considered, and it is shown that the increase of wing lift with Mach number at a given angle of attack is greatly dependent upon the aspect ratio b(exp 2)/F.

Cfd Analysis of Nozzle Jet Plume Effects on Sonic Boom Signature

Abstract: A computational fluid dynamics study is conducted to examine nozzle exhaust jet plume effects on the Sonic boom signature of a supersonic aircraft. A simplified axisymmetric nozzle geometry, representative of the nozzle on the NASA Dryden NF-15B Lift and Nozzle Change Effects on Tail Shock research airplane, is considered. The computational fluid dynamics code is validated using available wind-tunnel sonic boom experimental data. The effects of grid size, spatial order of accuracy. grid type, and flow viscosity on the accuracy of the predicted sonic boom pressure signature are quantified. Grid lines parallel to the Mach wave direction are found to give the best results. Second-order accurate upwind methods are required as a minimum for accurate sonic boom simulations. The highly underexpanded nozzle flow is found to provide significantly more reduction in the tail shock strength in the sonic boom N-wave pressure signature than perfectly expanded and overexpanded nozzle flows. A tail shock train in the sonic boom signature is observed for the highly underexpanded nozzle flow. Axisymmetric computational fluid dynamics simulations show the flow physics inside the F-15 nozzle to be nonisentropic and complex.

Parametric study of non-relativistic electrostatic shocks and the structure of their transition layer

Abstract: Nonrelativistic electrostatic unmagnetized shocks are frequently observed in laboratory plasmas and they are likely to exist in astrophysical plasmas. Their maximum speed, expressed in units of the ion acoustic speed far upstream of the shock, depends only on the electron-to-ion temperature ratio if binary collisions are absent. The formation and evolution of such shocks is examined here for a wide range of shock speeds with particle-in-cell simulations. The initial temperatures of the electrons and the 400 times heavier ions are equal. Shocks form on electron time scales at Mach numbers between 1.7 and 2.2. Shocks with Mach numbers up to 2.5 form after tens of inverse ion plasma frequencies. The density of the shock-reflected ion beam increases and the number of ions crossing the shock thus decreases with an increasing Mach number, causing a slower expansion of the downstream region in its rest frame. The interval occupied by this ion beam is on a positive potential relative to the far upstream. This potential pre-heats the electrons ahead of the shock even in the absence of beam instabilities and decouples the electron temperature in the foreshock ahead of the shock from the one in the far upstream plasma. The effective Mach number of the shock is reduced by this electron heating. This effect can potentially stabilize nonrelativistic electrostatic shocks moving as fast as supernova remnant shocks.

Global and Koopman modes analysis of sound generation in mixing layers

Abstract: It is now well established that linear and nonlinear instability waves play a significant role in the noise generation process for a wide variety of shear flows such as jets or mixing layers. In that context, the problem of acoustic radiation generated by spatially growing instability waves of two-dimensional subsonic and supersonic mixing layers are revisited in a global point of view, i.e., without any assumption about the base flow, in both a linear and a nonlinear framework by using global and Koopman mode decompositions. In that respect, a timestepping technique based on disturbance equations is employed to extract the most dynamically relevant coherent structures for both linear and nonlinear regimes. The present analysis proposes thus a general strategy for analysing the near-field coherent structures which are responsible for the acoustic noise in these configurations. In particular, we illustrate the failure of linear global modes to describe the noise generation mechanism associated with the vortex pairing for the subsonic regime whereas they appropriately explain the Mach wave radiation of instability waves in the supersonic regime. By contrast, the Dynamic Mode Decomposition (DMD) analysis captures both the near-field dynamics and the far-field acoustics with a few number of modes for both configurations. In addition, the combination of DMD and linear global modes analyses provides new insight about the influence on the radiated noise of nonlinear interactions and saturation of instability waves as well as their interaction with the mean flow.

Experimental Investigation of Effects of Mach Number on the Flow Instability in a Supersonic Inlet

Abstract: An extensive wind tunnel tests were conducted on an axisymmetric supersonic inlet at Mach numbers from 1.8 to 2.2 and at different values of mass flow rates. Frequencies of the buzz were achieved from the pressure data as well as the high speed shadowgraph pictures. For each Mach number, two main frequencies for the buzz were obtained. The inlet at its design condition was stable, but when the mass flow rate was reduced, at first the shock wave started to oscillate with a small amplitude which is matched the Ferri criterion (little buzz). In this situation, both high- and low-frequency oscillations occurred; however, the high frequency one was dominant, but the oscillations seemed to be irregular. By further decreasing the inlet mass flow, the shock oscillations got a single low frequency and high amplitude, which in agreement with the Dailey criterion (big buzz). At a free stream Mach number of 2.2, the frequencies of these two kinds of instabilities are about 554 and 137 Hz, respectively. At this Mach number for moderate mass flow, an oscillation with high frequency/large amplitude and constant period was observed. The oscillation has the combined characteristics of the little buzz and the big buzz, which is named “added buzz” here. It is shown that at fairly constant mass flow, the frequency of the buzz is independent of the Mach number. It is further found that the buzz instability affects the external flow as well as the internal one with the same frequency.

Weak shock wave reflection from concave surfaces

Abstract: The reflection of very weak shock waves from concave curved surfaces has not been well documented in the past, and recent studies have shown the possible existence of a variation in the accepted reflection configuration evolution as a shock wave encounters an increasing gradient on the reflecting surface. The current study set out to investigate this anomaly using high-resolution photography. Shock tube tests were done on various concave circular and parabolic geometries, all with zero initial ramp angle. Although the results have limitations due to the achievable image resolution, the results indicate that for very weak Mach numbers, M S < 1.1, there may be a region in which the reflection configuration resembles that of a regular reflection, unlike for the stronger shock wave case. This region exists after the triple point of the Mach reflection meets the reflecting surface and prior to the formation of the additional shock structures that represent a transitioned regular reflection. The Mach and transitioned regular reflections at 1.03 < M s < 1.05 also exhibit no signs of a visible shear layer, or a clear discontinuity at the triple point, and are thus also apparently different in the weak shock regime than what has been described for stronger shocks, similar to what has been shown for weak shocks reflecting off a plane wedge.

Beam forming of the underwater sound field from impact pile driving

Abstract: Observations of underwater noise from impact pile driving were made with a vertical line array. Previous studies [Reinhall and Dahl, J. Acoust. Soc. Am. 130, 1209–1216 (2011)] show that the dominant underwater noise from impact driving is from the Mach wave associated with the radial expansion of the pile that propagates down the pile at supersonic speed after impact. Here precise estimates of the vertical arrival angles associated with the down- and up-going Mach wave are made via beam forming, and the energy budget of the arrival structure is quantified.

Relativistic generalization of formation and ion-reflection conditions in electrostatic shocks.

Abstract: The theoretical model by Sorasio et al. [Phys. Rev. Lett. 96, 045005 (2006)] for the steady state Mach number of electrostatic shocks formed in the interaction of two plasma slabs of arbitrary density and temperature is generalized for relativistic electron and nonrelativistic ion temperatures. We find that the relativistic correction leads to lower Mach numbers and as a consequence ions are reflected with lower energies. The steady state bulk velocity of the downstream population is introduced as an additional parameter to describe the transition between the minimum and maximum Mach numbers that is dependent on the initial density and temperature ratios. In order to transform the solitonlike solution in the upstream region into a shock, a population of reflected ions is considered and differences from a zero-ion temperature model are discussed.

Research on Supersonic Inlet Bleed

Abstract: Phase I data results of the Fundamental Inlet Bleed Experiments project at NASA Glenn Research Center (GRC) are presented which include flow coefficient results for two single-hole boundary-layer bleed configurations. The bleed configurations tested are round holes at inclination angles of 90deg and 20deg both having length-to-diameter ratios of 2.0. Results were obtained at freestream Mach numbers of 1.33, 1.62, 1.98, 2.46, and 2.92 and unit Reynolds numbers of 0.984, 1.89, and 2.46 10(exp 7)/m. Approach boundary-layer data are presented for each flow condition and the flow coefficient results are compared to existing multi-hole data obtained under similar conditions. For the 90deg hole, the single and multi-hole distributions agree fairly well with the exception that under supercritical operation, the multi-hole data chokes at higher flow coefficient levels. This behavior is also observed for the 20deg hole but to a lesser extent. The 20deg hole also shows a markedly different characteristic at subcritical operation. Also presented are preliminary results of a Computational Fluid Dynamics (CFD) analysis of both configurations at the Mach 1.33 and a unit Reynolds number of 2.46 10(exp 7)/m. Comparison of the results shows the agreement to be very good.

Effect of Angle of Incidence on Stability Derivatives of A Wing

Abstract: In the Present paper effect of angle of incidence on pitching derivatives of a delta wing with curved leading edges of a attached shock case is been studied. A Strip theory is used in which strips at different span wise location are independent. This combines with similitude to give a piston theory. From the results it is found that stiffness and damping derivatives in pitch increasing linearly up to angle of attack 250 and then non-linearity creeps in. The Present theory is valid only for attached shock case. Effects of wave reflection and viscosity have not been taken into account. Results have been obtained for hypersonic flow of perfect gases over a wide range of angle of attack and Mach number.

Linear stability analysis on supersonic streamwise vortices

Abstract: In this study, the spatial growth rates of supersonic streamwise vortices were investigated by inviscid linear stability analysis. The freestream Mach numbers were 2.5, 5.0, and 7.5. In previous measurements taken to define the streamwise vortices, the stagnation temperature profile of supersonic flows is approximately uniform. This study found that the growth rate of vortices at the uniform stagnation temperature is smaller than that of isentropic vortices. The instability properties of the streamwise vortices can be explained by the ratio of the circulation to the axial velocity deficit, and also by the Mach number. Moreover, it is found that the compressibility effect, by which the instability reduces as the Mach number increases, is caused by the negative energy arising from the entropy gradient of supersonic vortices that accompanies the axial velocity deficit-like wake. From an energy perspective, the effect may reasonably be correlated with the large density perturbations in supersonic flows. This study also proposes a general convective Mach number for supersonic streamwise vortices. The normalized growth rates are shown to be a function of convective Mach number within the investigated range of ratio parameters.