Showing papers on "Mach wave published in 1989"

Prediction and control of transition in supersonic and hypersonic boundary layers

Abstract: Computations for sharp cones, using the e N method with N=10, show that the first oblique Tollmien-Schlichting mode is responsible for transition at adiabatic wall conditions for freestream Mach numbers up to about 7. For cold walls, the two-dimensional second mode dominates the transition process at lower hypersonic Mach numbers due to the well-known destabilizing effect of cooling on the second mode

Inviscid spatial stability of a compressible mixing layer

Abstract: Presented are the results of a study of the inviscid spatial stability of a parallel compressible mixing layer. The parameters of this study are the Mach number of the moving stream, the ratio of the temperature of the stationary stream to that of the moving stream, the frequency and the direction of propagation of the disturbance wave. Stability characteristics of the flow as a function of these parameters are given. It is shown that if the Mach number exceeds a critical value there are always two groups of unstable waves. One of these groups is fast with phase speeds greater than 1/2, and the other is slow with speeds less than 1/2. Phase speeds for the neutral and unstable modes are given, as well as growth rates for the unstable modes. It is shown that three-dimensional modes have the same general behavior as the two-dimensional modes but with higher growth rates over some range of propagation direction. Finally, a group of very low frequency unstable modes was found for sufficiently large Mach numbers. These modes have very low phase speeds but large growth rates.

Propagation mechanism of quasi-detonations

Abstract: The propagation mechanism of quasi-detonations in very rough tubes is studied using high speed schlieren photography. Stoichiometric mixtures of H2, C2H4 and C3H8 in oxygen at an initial pressure range 10≤po≤160 torr are investigated in a 61.8×61.8 mm by 1.5 m long channel with two-dimensional obstacles with a height of 25.4 mm and for various obstacle spacings. The results indicate that shock reflections (transition from regular to Mach reflections) from the walls lead to re-initiation. The obstacle spacing is found to represent an effective reaction zone length (or cell length) of the quasi-detonation. At the critical condition of transition from the choking to the quasi-detonation regime, this effective reaction zone length is found to be about twice the normal cell length of the mixture in accordance with Shchelkin's stability criterion for a perturbed wave. The minimum open dimension of the channel is found to be of the order of a cell size λ for transition to the quasi-detonation regime in agreement with the previous results of Peraldi17 and Gu8 for rough tubes. Photographic observations of the propagation mechanism in the choking regime reveal the absence of ignition via shock reflection. The placement of wire screens to damp the shock reflections at the channel walls suppresses the transition to quasi-detonations indicating the essential role of shock reflections. It is not clear whether the adiabatic heating or the turbulent vortex mixing associated with the shear layer wall jet by the Mach stem near the wall is the responsible mechanism for re-initiation.

Experimental Shock-Wave Interference Heating on a Cylinder at Mach 6 and 8

Abstract: This paper presents the details of an experimental study of shock-wave interference heating on a cylindrical leading edge representative of the cowl of a rectangular hypersonic engine inlet. The study, which was conducted at Mach numbers of 6.3, 6.5, and 8.0, has provided 1) detailed pressure and heat-transfer-rate distributions for a two-dimensional shock-wave interference on a cylinder and 2) insight into the effects of temperature-d ependent specific heats on the phenomena. The peak pressure and heat-transfer rates were 2-25 times the undisturbed flow stagnation-point levels. The peak levels and their gradients increased with Mach number. Variation in specific heats and, hence, in the ratio of specific heats with temperature is manifested in slightly lower loads and amplification factors than for corresponding perfect-gas conditions.

A numerical investigation of the compressible mixing layer

Abstract: The effect of Mach number on the plane muong layer has been investigated by means of linear stability theory and two- and three-dimensional direct numerical simulations of the compressible Navier-Stokes equations. The objective was to identify the effects of compressibility on a building-block fluid flow, with applications to supersonic mixing and combustion. Results from linear stability theory show that the amplification rate is reduced as Mach number is increased. Above a convective Mach number of 0.6 it is found that three-dimensional waves are more amplified than two-dimensional waves and a simple relation is found to give the orientation of the most amplified waves. It is also shown that the linear stability theory can be used to predict the mixing layer growth rate as a function of velocity ratio, density ratio and Mach number. Two-dimensional simulations show a strong reduction in growth rate of the two-dimensional motion as Mach number is increased, with more elongated structures forming at high Mach numbers. Shock waves are observed in two-dimensional simulations above a convective Mach number of 0.7. The supersonic modes of instability, which are the only two-dimensional unstable modes at high Mach numbers, are shown to be radiating and vortical, but have very low growth rates. Three-dimensional simulations with random initial conditions confirm the linear stability result that oblique waves become the most amplified waves at high Mach numbers, with no evidence for any other modes of instability. Simulations beginning with a two-dimensional wave and a pair of equal and opposite oblique waves show a change in the evolved large-scale structure as Mach number is increased. Above a convective Mach number of 0.6 the oblique modes have most of the energy in the developed structure, and above a convective Mach number of 1 the two-dimensional instability wave has little effect on flow structure. Similar organized structure was found in a simulation with random initial conditions. No shock waves were found in the three-dimensional simulations, even at convective Mach numbers above 1.

An experimental investigation of the sonic criterion for transition from regular to Mach reflection of weak shock waves

Abstract: The signal speed, namely the local sound speed plus the flow velocity, behind the reflected shocks produced by the interaction of weak shock waves (M i < 1.4) with rigid inclined surfaces has been measured for several shock strengths close to the point of transition from regular to Mach reflection. The signal speed was measured using piezo-electric transducers, and with a multiple schlieren system to photograph acoustic signals created by a spark discharge behind a small aperture in the reflecting surfaces. Both methods yielded results with equal values within experimental error. The theoretical signal speeds behind regularly reflected shocks were calculated using a non-stationary model, and these agreed with the measured results at large angles of incidence. As the angle of incidence was reduced, for the same incident shock Mach number, so as to approach the point of transition from regular to Mach reflection, the measured values of the signal speed deviated significantly from the theoretical predictions. It was found, within experimental uncertainty, that transition from regular to Mach reflection occurred at the experimentally observed sonic point, namely, when the signal speed was equal to the speed of the reflection point along the reflecting surface. This sonic condition did not coincide with the theoretical value.

Documentation of Two- and Three-Dimensional Hypersonic Shock Wave/Turbulent Boundary Layer Interaction Flows

Abstract: Experimental data for a series of two- and three-dimensional shock wave/turbulent boundary layer interaction flows at Mach 7 are presented. Test bodies, composed of simple geometric shapes, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface-pressure and heat-transfer distributions as well as limited mean-flow-field surveys in both the undisturbed and the interaction regimes. The data are presented in a convenient form for use in validating existing or future computational models of these generic hypersonic flows.

A Newton/upwind method and numerical study of shock wave/boundary layer interactions

Abstract: The objective of the paper is two-fold. First, an upwind/central differencing method for solving the steady Navier-Stokes equations is described. The symmetric line relation method is used to solve the resulting algebraic system to achieve high computational efficiency. The grid spacings used in the calculations are determined from the triple-deck theory, in terms of Mach and Reynolds numbers and other flow parameters. Thus the accuracy of the numerical solutions is improved by comparing them with experimental, analytical, and other computational results. Secondly, the shock wave/boundary layer interactions are studied numerically, with special attention given to the flow separation. The concept of free interaction is confirmed. Although the separated region varies with Mach and Reynolds numbers, it is found that the transverse velocity component behind the incident shock, which has not been identified heretofore, is also an important parameter. A small change of this quantity is sufficient to eliminate the flow separation entirely.

Comparison with experiment for tvd calculations of blast waves from a shock tube

Abstract: Harten's second-order-accurate total-variation-diminishing (TVD) scheme is applied to calculation of flow from the open end of a shock tube. Comparison of numerical results with available experimental data for overpressure at selected points around the shock tube exit shows good agreement. Numerically indicated positions of the moving shock front and Mach stem also compare well with flow shadowgraph data. Where the problem geometry is sufficiently simple and rectangular gridding can be used, Harten's method affords a good choice for blast wave calculations.

Stability of compressible boundary layers

Abstract: The stability of compressible 2-D and 3-D boundary layers is reviewed. The stability of 2-D compressible flows differs from that of incompressible flows in two important features: There is more than one mode of instability contributing to the growth of disturbances in supersonic laminar boundary layers and the most unstable first mode wave is 3-D. Whereas viscosity has a destabilizing effect on incompressible flows, it is stabilizing for high supersonic Mach numbers. Whereas cooling stabilizes first mode waves, it destabilizes second mode waves. However, second order waves can be stabilized by suction and favorable pressure gradients. The influence of the nonparallelism on the spatial growth rate of disturbances is evaluated. The growth rate depends on the flow variable as well as the distance from the body. Floquet theory is used to investigate the subharmonic secondary instability.

Improvements and applications of a streamwise upwind algorithm

Abstract: An improved streamwise upwind algorithm has been used to study conical flow fields. In the present method, additional terms have been introduced in the cross-flow direction to prevent solution decoupling in supersonic flows, and the local Mach number is taken into account in order to evaluate the rotated differencing. It is found that the formula captures oblique shock waves in the same manner as Roe's (1986) formula, has good convergence properties, and accurately computes shear flows.

Thermodynamics and wave processes in high Mach number propulsive ducts

Abstract: Analysis of the flow in a propulsive duct indicates that, at high Mach numbers, the thermodynamic energy of the fluid is delivered directly into compression and expansion waves. The importance of the resulting wave phenomena is explored by considering the net thrust delivered by a two dimensional convergent-divergent duct with a simplified, planar heat addition zone. It is found that net thrust is reduced when operating at other than the design Mach number, an effect which is most severe for Mach numbers exceeding the design value. An idealized model is developed, involving a self-induced heat injection cycle, and it is seen that the waves produced by this cycle can be the dominant agency in producing thrust.

A transition detection study at Mach 1.5, 2.0 and 2.5 using a micro-thin hot-film system

Abstract: A boundary-layer transition detection study was conducted in the Langley unitary plan wind tunnel with an array of micro-thin hot films on a flat plate at Mach numbers from 1.0 to 4.5 million per foot. Transition locations were obtained online from the variation of normalized RMS (root mean square) voltages from an array of hot-film sensors for both natural transition and a grit-induced wedge of turbulence. The effects of Mach number and Reynolds number on the location and lengths of the transition region for the two types of transition are presented from the online data. Also shown are the unit Reynolds number effects on transition Reynolds number, voltage-versus-time traces, spectra, and the data-acquisition system. >

A reconsideration of the transition criterion from mach to regular reflection over cylindrical concave surfaces

Abstract: Following the idea forwarded by Ben-Dor and Takayama (1985) a new propagation path was assumed for the corner generated signals. In addition to the new chosen propagation path, one of the simplified assumptions used by Ben-Dor and Takayama, namely thatu+a remains constant behind the incident shock wave, was further simplified, i.e., bothu anda were assumed to be constant behind the incident shock wave. This new path and the further simplified assumption led to a new transition criterion from Mach to regular reflection over a cylindrical wedge which unlike the two criteria developed in Ref. by Ben-Dor and Takayama (1985) has the ability to predict the transition wedge angle quite accurately in the entire range of the incident shock wave Mach numbers which were investigated in Ref. by Ben-Dor and Takayama(1985) and in this study.

Sound power spectrum and wave drag of a propeller in flight

Abstract: Theory is presented for the sound power and sound power spectrum of a single rotation propeller in forward flight. Calculations are based on the linear wave equation with sources distributed over helicoidal surfaces to represent effects of blade thickness and steady loading. Sound power is distributed continuously over frequecy, as would be expected from Doppler effects, rather than in discrete harmonics. The theory is applied to study effects of sweep and Mach number in propfans. An acoustic efficiency is defined as the ratio of radiated sound power to shaft input power. This value is the linear estimate of the effect of wave drag due to the supersonic blade section speeds. It is shown that the acoustic efficiency is somewhat less than 1 percent for a well designed propfan.

Direct numerical simulations of a plane compressible wake: Stability, vorticity dynamics, and topology

Abstract: Recent interest in supersonic combustion and problems of transatmospheric flight has prompted renewed research efforts in laminar-turbulent free shear flow transition. In the present work, linear stability theory and direct numerical simulations are used to study the effect of Mach number on the linear, nonlinear, and three-dimensional aspects of transition in a plane compressible wake. Direct numerical simulations are also used to study the sensitivity of a compressible wake to (1) phase effects and (2) two- and three-dimensional subharmonics. A linear stability analysis shows that the influence of increasing Mach number is stabilizing, resulting in reduced growth rates for both antisymmetric and symmetric modes of the wake. This reduction is due to baroclinic and dilatational effects as revealed from the linear eigenfunctions. For both low and high Mach numbers, the least stable wave is a two-dimensional antisymmetric mode aligned with the stream-wise direction. Three-dimensional simulations were performed to study the effect of phase angle between a fundamental and a pair of oblique waves on the development of the large-scale structures in a wake. Finally, the topology of the computed velocity, vorticity, and pressure gradient fields is determined using a generalized three-dimensional critical point theory. 78 refs., 83 figs.

Fluctuating pressures measured beneath a high-temperature, turbulent boundary layer on a flat plate at Mach number of 5

Abstract: Fluctuating pressures were measured beneath a Mach 5, turbulent boundary layer on a flat plate with an array of piezoresistive sensors. The data were obtained with a digital signal acquisition system during a test run of 4 seconds. Data sampling rate was such that frequency analysis up to 62.5 kHz could be performed. To assess in situ frequency response of the sensors, a specially designed waveguide calibration system was employed to measure transfer functions of all sensors and related instrumentation. Pressure time histories were approximated well by a Gaussian prohibiting distribution. Pressure spectra were very repeatable over the array span of 76 mm. Total rms pressures ranged from 0.0017 to 0.0046 of the freestream dynamic pressure. Streamwise, space-time correlations exhibited expected decaying behavior of a turbulence generated pressure field. Average convection speed was 0.87 of freestream velocity. The trendless behavior with sensor separation indicated possible systematic errors.

High explosive assembly for projecting high velocity long rods

Abstract: A high explosive assembly and a method are disclosed for projecting a longod at high velocity with enhanced penetrating energy. The high explosive assembly has an elongated core of a first high explosive having a first Chapman-Jouguet detonation velocity, an elongated liner positioned substantially along the longitudinal axis of the core, and an elongated jacket of a second high explosive encasing the core and having a second Chapman-Jouguet detonation velocity greater than the core Chapman-Jouguet detonation velocity. The jacket high explosive, upon detonation, continuously initiates detonation of the core high explosive by an imposed oblique detonation front which converges toward the center of the detonating core with time, until a trailing mach stem emerges therefrom as detonation progresses. The mach stem grows with time as the detonation continues until a steady state mach stem disk results, and detonation proceeds further as a highly overdriven detonation of the core to expel the liner as a long rod at high velocity.

Full-potential analysis of a supersonic delta wing/body

Abstract: Experimental spanwise pressure distributions for a 60 deg delta wing/body of approximate fineness ratio 7.6 have been obtained and compared to predictions using full-potential theory. Analysis was performed at Mach 1.6 for angles of attack in the range of 0.8-10 deg, and for Mach numbers ranging from 1.4 to 1.8 at lift coefficient 0.3. The intent of the study was to examine an attached flow approach for maneuver wing design in the presence of a fuselage. For the Mach number and angle-of-attack conditions considered, the full-potential theory accurately modeled the pressure distributions, provided the flow remained attached. By combining the full-potential theory results with empirical shock-induced separation criteria, it was found that the onset of shock-induced separation can be predicted.

Experimental study of free-shear layer transition above a cavity at Mach 3.5

Abstract: The transition behavior of a free-shear layer above a cavity with high and low levels of freestream acoustic disturbances has been studied at Mach 3.5. Optical techniques, mean pitot pressure measurements, and hot-wire measurements were employed to detect transition locations. Transition Reynolds numbers of between 363,000 and 530,000 were found, in agreement with previous results. It is suggested that upstream convected disturbances may be at least partially responsible for the insensitivity of transition Reynolds numbers to the freestream acoustic disturbance field.

Nonaxisymmetric Body, Supersonic, Inviscid Dynamic Derivative Prediction

Abstract: : A supersonic, aerodynamic computational model, which is the basis of the NANC code, has been extended to compute dynamic derivatives. The extension is to the inviscid contribution of constant angular rates and axial accelerations. The body geometry limitations are the same as for the steady- state model. Here, a pointed body or equivalent pointed body is assumed for low Mach numbers; at higher Mach numbers, the effect of axial acceleration is neglected and the body may be blunt. The body may be noncircular with planar discontinuities, including inlets, with fins (up to six per fin set), which lie on a cylindrical coordinate ray. For the low Mach number range, the original second-order potential model has been extended for angular rate derivative prediction. For the acceleration rate derivatives, a 'hybrid' first- and second- order model has been developed. For the high Mach number range, an equivalent angle-of-attack vector is defined and combined with local solution models. Computational comparisons are made with experimental data, primarily for pitch and roll damping derivatives. Keywords: Aerodynamic loading; Supersonic characteristics; Mathematical prediction; Projectile models; Dynamic derivatives; Supersonic/aerodynamic computational model; First and second order model; Roll/pitch damping; Fin magnus.

A parametric study of three-dimensional separation at a wing/body junction for supersonic free-stream conditions

Abstract: The problem of three-dimensional separation at a wing/body junction has been investigated numerically using a three-dimensional Navier-Stokes code which employs the MacCormack's time split finite volume technique. An algebraic grid generation technique is used for generating the grid at a wing/body junction. Specific computational results on velocity and pressure distribution in the separated flow region are compared with the experimental results. A parametric study of flow parameters such as Mach number and Reynolds number have been carried out to understand their effect in interaction flow field. The parametric study indicates a strong dependency of the number of vortices at the junction on Mach number and Reynolds number.