Numerical and experimental investigations on the mach 2 pseudo-shock wave in a square duct

This paper describes numerical and experimental investigations for the multiple shock wave/turbulent boundary layer interaction in a Mach 2 supersonic square duct. The numerical simulation is carried out with the Harten-Yee second-order accuracy TVD scheme and the Baldwin-Lomax turbulence model. The flow conditions are a free-stream Mach number ofM _∞≈=2.0 and a Reynolds number ofRe _∞;=2.5×10⁷ and the flow confinements are δ_∞/h=0.15 (case A) and δ_∞/h=0.25 (case B), respectively. The computational results for both cases show good agreement with the experimental results. Based on these agreements, the flow quantities, which are very difficult to obtain experimentally, are analyzed by numerical simulation. Moreover, the effect of flow confinement on the pseudo-shock wave characteristics is also presented.     

Shock associated noise of supersonic jets from convergent-divergent nozzles

Results of experimental and theoretical studies of the characteristics of shock associated noise from imperfectly expanded supersonic jets over an extensive range of underexpanded and overexpanded operating conditions are described. This kind of broadband noise is believed to be generated by the weak but coherent interaction between the downstream propagating large scale turbulent flow structures in the mixing layer of the jet and the nearly periodic shock cell system. Theoretical reasoning based on this mechanism leads to the scaling formula that the intensity of shock associated noise varies as (M_j² ? M_d²)² where M_j and M_d are the fully expanded jet operating Mach number and nozzle design Mach number, respectively. This formula holds for underexpanded as well as overexpanded jet Mach numbers. In addition, a peak frequency formula is also derived from the same model. The noise intensity, directivity and spectra of supersonic jets from a convergent-divergent nozzle of design Mach number 1·67 were measured in an anechoic facility over the Mach number range of 1·1 to 2·0. The effect of jet temperature was investigated by operating the jet at three temperature conditions. These sets of data provide sufficient information for fully assessing the relative importance and characteristics of shock associated noise of supersonic jets from convergent-divergent nozzles. Comparisons between theoretical results and measurements show very favorable agreement.     

Experimental study of the evolution of periodic controlled disturbances in a hypersonic viscous shock layer on a flat plate

Characteristics of the fields of mean density and density fluctuations measured with introduction of periodic disturbances into a hypersonic viscous boundary layer on a flat plate are presented. The experiments are performed for a flow Mach number M_∞ = 21, Reynolds number per meter Re_1∞ = 6·10⁵ m⁻¹, and temperature factor of the surface T _w /T ₀ = 0.26. The disturbances are introduced into the shock layer by an oblique gasdynamic whistle. The work was financially supported by the Russian Foundation for Basic Research (Grants Nos. 04-01-00474 and 05-08-33436).     

Design of supersonic three-dimensional inlets using two-dimensional isentropic compression flow

The design of supersonic three-dimensional inlets using the V-shaped body forming a two-dimensional flow including an initial oblique shock wave and a subsequent isentropic compression wave is considered. Such a flow appears attractive for inlets design due to a possibility of obtaining high compression levels of external flow over the inlet ramp with small losses of the total pressure. Numerical computations of the flows around the designed configurations were carried out in design and off-design regimes using Euler code. The flow structure was identified, the aerodynamic characteristics of the inlets were determined. The investigation covers the range of supersonic speeds corresponding to the freestream Mach numbers M_∞= 1.8−2.5.     

Numerical flow visualization of a Single Expansion Ramp Nozzle with hypersonic external flow

Numerical simulation of scramjet asymmetric nozzle flow is carried out to visualize and investigate the effects of interaction between engine exhaust and hypersonic external flow. The Single Expansion Ramp Nozzle (SERN) configuration studied here consists of flat ramp and a cowl with different combinations of ramp angle and cowl geometry. UsingPARAS 3D, simulations are performed for a free stream Mach number of 6.5 that constitutes the external flow around the vehicle. Appropriate specific heats ratio has been simulated for the jet and free stream flow. External shock wave due to jet plume interaction with free stream flow, the internal barrel shock wave and the shear layer emanating from the cowl trailing edge and sidewalls are well captured. Wall static pressure distribution on the nozzle ramp for different nozzle expansion angles has been computed for both with and without side fence. Axial thrust and normal force have been evaluated by integrating the wall static pressure. Effect of cowl length variation and side fence on the SERN performance has also been studied and found to be quite significant. Based on this study, an optimum ramp angle at which the SERN generates maximum axial thrust is obtained. SERN angle of 20° was found to be optimum when the flight axis coincides with nozzle axis.     

Analysis of possible improvement of acceleration of a high-velocity air-breathing flying vehicle

Results of parametric calculations of the total aeropropulsive characteristics and characteristics of acceleration of a small-scale high-velocity flying vehicle with an air-breathing engine are presented. Integral parameters of acceleration from the flight Mach number M_∞ = 4 to M_∞ = 7 are determined, namely, the time required fuel stock, and range. A schematic configuration of the vehicle is considered, which allows studying the basic parameters, such as the forebody shape, the angles of surfaces of compression of the stream captured by the inlet, angles of external aerodynamic surfaces of the airframe, relative planform area of the wing panels, and relative area of the nozzle cross section. A comparative estimate of the effect of these parameters shows that it is possible to improve the characteristics of acceleration of vehicles of the type considered.     

Comparative analysis of aerodynamics of waveriders designed on the basis of conical and planar flows

A comparative analysis of the aerogasdynamics of waveriders designed from two different flow types, namely the axisymmetric conical flows and the flows behind two-dimensional planar shocks, was carried out. Integral aerodynamic characteristics of a configuration with different forms of transverse contours of the lower surface were considered, and the “conical” and equivalent “planar” waveriders optimal in terms of the life-todrag ratio were compared. The gas-dynamic structure features in off-design flow regimes were investigated on the basis of numerical solution of the Euler equations for the Mach numbers, which were both higher and lower than the design values. Obtained data refer to the freestream Mach number range M_∞ = 4–10.     

Evolution of a vortex system in the vicinity of an external dihedral corner

Results of experimental investigations of the evolution of a vortex system formed in a supersonic flow past a streamwise-aligned external dihedral right angle owing to a difference in pressures on the upper and side faces of the corner are analyzed. The experiments are performed in a T-313 wind tunnel based at ITAM SB RAS at Mach numbers M_∞ = 2.27, 3, and 4, and angles of attack α = −4° ÷ +20°. It is shown that the size of the vortex system influence zones is almost independent of the free-stream Mach number in the examined range of the angles of attack, and the relative values of flow rarefaction on the model surface under the primary vortex core smoothly tend to their minimum values.     

Transition of the boundary layer on a flat plate at supersonic and hypersonic velocities

The transition of the boundary layer from the laminar to the turbulent state on a smooth flat plate at a zero angle of attack is studied in the range of Mach numbers M_∞ = 2–6. It is demonstrated that the results measured at the end of the transition region can be approximated by a simple dependence suitable for applications, which does not require additional measurements, is valid in the range of Mach numbers M_∞ = 2–10, and, with an error lower than 20 %, can be used to estimate the location of the transition region on a flat plate in geometrically similar wind tunnels.     

Measurement and simulation of edge plasma flow in the Tore Supra tokamak

A simple fluid model is used to interpret two edge flow phenomena that are observed in the scrape-off layer of the Tore Supra tokamak. (1) The Mach number of the parallel flow midway between the two sides of the toroidal limiter is always a significant fraction of the ion sound speed, typicallyM _?=?0.3÷?0.6. directed towards the high field side. In Tore Supra theE×B drift is relatively unimportant, so the esixtence of alarge parallel flow implies a significant poloidal aysmmetry of the SOL source. (2) Massive injection of neutral gas leads to a recycling source that is poloidally localized near the nozzle. The parallel Mach number changes during injection. The sign of the change is positive for injection on the high field side, and negative for injection on the low field side. The ionization source that is needed to produce the change is about 5÷10% of the total amount of injected gas, consistent with atomic physics calculations.     

Experimental Study of Non-Equilibrium Ionization in a Supersonic CO Flow

Experimental results are presented on ionization processes in the vibrationally non-equilibrium flow of CO preliminarily heated by a shock wave and adiabatically cooled in the supersonic nozzle with a flow Mach number of M = 3.9. It is shown that for a wide range of flow stagnation temperatures, the value of the recorded semi-self-maintained current considerably exceeds the one due to thermal gas ionization in the plenum chamber.     

Double-diaphragm shock tube at the Ioffe Physicotechnical Institute

Results of numerical calculations of the flow in a double-diaphragm shock tube with a tailored contact surface are reported. The calculations were carried out using a model of an ideal shock tube allowing for the real properties of the driver gas at high pressures and equilibrium thermodynamics of the processes behind the shock waves at Mach numbers M _s1 of the shock wave in the working gas varying in the range 5–25. Flow regimes with a tailored contact surface were obtained for Mach numbers M _s1=6.3, 11, and 15 using the double-diaphragm shock tube at the Ioffe Physicotechnical Institute. Under these conditions, the parameters of the working gas were kept constant for more than 1 ms. The calculated data were compared with the experimental results and it was shown that the calculated data may be used to determine the section lengths in a double-diaphragm shock tube and to estimate its operating time. The calculated values of the initial pressure in the sections of the tube were substantially lower than those achieved experimentally. Measurements were made of the static pressure along the axis of a conical nozzle during the expansion of hydrogen (initial temperature T ₀=293 K) and shock-heated nitrogen (T ₀=4000 K). It is found that the expansion of hydrogen is accompanied by deactivation of the rotational degrees of freedom, and that partial freezing of the vibrational degrees of freedom takes place in the nitrogen stream. Zh. Tekh. Fiz. 67, 88–95 (November 1997)     

Supersonic Jets with Periodic Structure due to Arc plasma Expansion into a Low pressure ambient

A supersonic plasma jet was produced by a d.c. arc plasma generator operated at normal pressure and connected to a low-pressure (p∞ = 0.2-50 kPa) chamber via cylindrical nozzle with diameter of 2.5 mm. The argon gas flow rate was G = 0.025 to 0.35 g.s^?1. In some experiments current I_E ≦ 30 A passed coaxially through the initial part of the jet. Photographic records of the jet and pressure measurements are in agreement with theoretical predictions by a simple one-dimensional, gasdynamical model capable of self-consistent calculations throughout the plasma source/jet system. Periodic jet structure is observed over a wide range of experimental conditions, incl. in highly under-expanded flow. The jet expansion angle and Mach disc position vary with p∞, G and I_E, but are nearly constant at different arcing currents.     

Development of three-dimensional turbulent separation in the neighborhood of incident crossing shock waves

The results are presented for experimental investigation of the peculiarities of the development of three-dimensional turbulent separated flows on a flat surface for the Mach number M_∞ = 4 and the Reynolds number Re₁ ∼ 55·10⁶ m⁻¹ under the conditions of the flow around two identical cylindrical bodies of revolution of diameter D = 50 mm and the body aspect ratio L _b/D = 5 with conical forebodies with semi-apex angles β_c = 30, 20, 15, and 10° located above plate in parallel to one another and to the flow. The typical stages of the three-dimensional separation development are considered under the reducing distance between the axes of the bodies within the range of Z = Δz/D = 1.06–3.0 at their fixed distance from the surface (Y = Δy/D = 0.96). The topology of limiting streamlines and the peculiarities of pressure fields on the surface as well as the gasdynamic structure of separated flows arising at the interaction of crossing bow shocks propagating from the bodies and at the interaction of secondary disturbances with the boundary layer are analysed.     

The influence of surface porosity on the stability and transition of supersonic boundary layer on a flat plate

In the present study, we examined, both experimentally and theoretically, the influence of surface permeability on the stability and laminar-turbulent transition of supersonic boundary layer at free-stream Mach number M_∞ = 2. A satisfactory agreement was obtained between the data calculated by the linear theory of stability and the data obtained in experiments with natural disturbances performed on models with different porous inserts.     

Gasdynamic design of a two-dimensional supersonic inlet with the increased flow rate factor

A two-dimensional inlet of external compression with the increased flow rate factor at high supersonic velocities is constructed by the method of gasdynamic design. Its feature is that a flow with the initial oblique shock wave and the subsequent centered isentropic compression wave is formed over the external compression ramp of the inlet. These waves interact with one another so that a resulting stronger oblique shock wave and a velocity discontinuity arise in front of the entrance to the inlet internal duct. An example of an inlet configuration with the design flow regime corresponding to the Mach number M_d = 7 is considered. The characteristics of this inlet were obtained in the range of the free-stream Mach numbers M = 4–7 with the use of a Navier—Stokes code for turbulent flow. They are compared with characteristics of an equivalent conventional shocked inlet. As computations have shown, the inlet with the isentropic compression wave has much higher values of flow rate factor φ at Mach numbers M < M_d. So, for example, at M = 4 the value φ ≈ 0.72 for it is by 33 % higher in comparison with φ ≈ 0.54 for the equivalent shocked inlet.     

PSP visualization studies on a convergent nozzle with an

An in-house Pressure Sensitive Paint (PSP) formulation has been developed at the Aero-Physics Laboratory at the University of Manchester. The PSP uses Bathophenanthroline Ruthenium as the luminophore molecule and is incorporated in a sol-gel matrix. Excitation occurs at400–500 nm and emission at550–650 nm. The Stern-Volmer plot of the PSP reveals small temperature dependence, which has always been an intrinsic drawback of PSPs. As a baseline experiment the PSP has been applied to examine the side-wall pressure field of the flow through a convergent nozzle with an ejector, at fully expanded Mach numbers in the rangeM _j =0.52–1.36. Simultaneous static pressure measurements were also conducted to ascertain the accuracy of the PSP results. The paint has demonstrated satisfactory capabilities in not only measuring static pressures but also in visualizing key physical elements of the flow, such as the location of the expansion and oblique shock waves present in such flows.     

Investigation of temperature fields in supersonic flow behind a backward-facing step

The results of numerical modelling and experimental investigations of high-enthalpy turbulent flows in the neighborhood of 90-degree backward-facing steps at the Mach numbers M_∞ = 2–4 are presented. The experiments were conducted in the hot-shot wind tunnel IT-302M of ITAM SB RAS. The computations were carried out on the basis of the full Favres-averaged Navier — Stokes equations augmented by the Wilcox turbulence model. The temperature factor influence on the flow structure in the separated zone and temperature distributions was investigated numerically for different Mach numbers. The wall temperature is shown to affect significantly the quantity and sizes of recirculation vortices as well as the temperature distribution in the zone of flow separation and reattachment. The computational results are compared with experimental data on the pressure distribution on the model surface and the wave structure of the flow.     

激波作用下SF6气泡界面演化和射流发展的数值模拟

 数值研究了平面激波冲击氮气环境中SF₆气泡界面的Richtmyer-Meshkov不稳定性,重点关注其中的激波聚焦及射流的产生和发展过程。在入射激波马赫数为1.23的情况下,给出了压力、密度、数值纹影和涡量等物理量的演化图像,定量分析了流场中压力最大值、密度最大值、射流速度、环量和斜压力矩随时间的变化关系。计算结果表明,平面激波冲击SF₆气泡过程有很强的聚能效应,在气泡内部靠近下游极点处发生激波近似理想聚焦和点爆炸现象,直接导致出现二次波系以及向下游运动的细长射流结构。相比入射激波,二次波系产生斜压力矩和涡量的能力要弱得多。     

Shock-structures in the acoustic field of a Mach 3 jet with crackle

The acoustic waveforms produced by an unheated supersonic and shock free jet operating at a gas dynamic Mach number of 3 and an acoustic Mach number of 1.79 are examined over a large spatial domain in the (x,r)-plane. Under these operating conditions, acoustic waveforms within the Mach cone comprise sawtooth-like structures which cause a crackling sound to occur. The crackling structures produced by our laboratory-scale nozzle are studied in a range-restricted environment, and so, they are not the consequence of cumulative nonlinear waveform distortions, but are rather generated solely by local mechanisms in, or in close vicinity to, the jet plume. Our current work focuses on characterizing the temporal and spectral properties of these shock-structures. A detection algorithm is introduced which isolates the shock-structures in the temporal waveforms based on a pressure rise time and shock strength that satisfy user defined thresholds. The average shapes of the shock-structures are shown to vary along polar angles centered on the post-potential core region of the jet. Spectral characteristics of the crackling structures are then determined using conventional wavelet-based time–frequency analyses. Differences between the global wavelet spectrum and the local wavelet spectrum computed from instances when shocks are detected in the waveform show how shock-structures are more pronounced at shallow angles to the jet axis. The findings from this energy-based metric differ from those obtained using the skewness of the pressure and the pressure derivative.