Supersonic Freestream Research Articles

Control of flow separation on a contour bump by jets in a Mach 1.9 free-stream: An experimental study

Flow separation control over a three-dimensional contour bump using jet in a Mach 1.9 supersonic free-stream has been experimentally investigated using a transonic/supersonic wind tunnel. Jet total pressure in the range of 0–4bar was blowing at the valley of the contour bump. Schlieren photography, surface oil flow visualisation and particle image velocimetry measurements were employed for flow visualisation and diagnostics. Experimental results show that blowing jet at the valley of the contour bump can hinder the formation and distort the spanwise vortices. The blowing jet can also reduce the extent of flow separation appears downstream of the bump crest. It was observed that this approach of flow control is more effective when high jet total pressure is employed. It is believed that a pressure gradient is generated as a result of the interaction between the flow downstream of the bump crest and the jet induced shock leads to the downwards flow motion around the bump valley.

Shock wave interactions between slender bodies

A complex interference flowfield consisting of multiple shocks and expansion waves is produced when high-speed slender bodies are placed in close proximity. The disturbances originating from a generator body impinge onto the adjacent receiver body, modifying the local flow conditions over the receiver. This paper aims to uncover the basic gas dynamics produced by two closely spaced slender bodies in a supersonic freestream. Experiments and numerical simulations were used to interpret the flowfield, where good agreement between the predictions and measurements was observed. The numerical data were then used to characterise the attenuation associated with shock wave diffraction, which was found to be interdependent with the bow shock contact perimeter over the receiver bodies. Shock-induced boundary layer separation was observed over the conical and hemispherical receiver bodies. These strong viscous-shock interactions result in double-reflected, as well as double-diffracted shock wave geometries in the interference region, and the diffracting waves progress over the conical and hemispherical receivers’ surfaces in “lambda” type configurations. This gives evidence that viscous effects can have a substantial influence on the local bow shock structure surrounding high-speed slender bodies in close proximity.

Supersonic cavity flows over concave and convex walls

Supersonic cavity flows are characterized by compression and expansion waves, shear layer, and oscillations inside the cavity. For decades, investigations into cavity flows have been conducted, mostly with flows at zero pressure gradient entering the cavity in straight walls. Since cavity flows on curved walls exert centrifugal force, the features of these flows are likely to differ from those of straight wall flows. The aim of the present work is to study the flow physics of a cavity that is cut out on a curved wall. Steady and unsteady numerical simulations were carried out for supersonic flow through curved channels over the cavity with L/H = 1. A straight channel flow was also analyzed which serves as the base model. The velocity gradient along the width of the channel was observed to increase with increasing the channel curvature for both concave and convex channels. The pressure on the cavity floor increases with the increase in channel curvature for concave channels and decreases for convex channels. Moreover, unsteady flow characteristics are more dependent on channel curvature under supersonic free stream conditions.

Shock wave bifurcation in convergent–divergent channels of rectangular cross section

This work addresses two- and three-dimensional turbulent flow in simple channels, modeling the air intakes of rectangular cross section. Flow regimes with a supersonic free stream and supersonic velocities at the throat or immediately downstream of the throat are considered. Bifurcations of the shock wave arising ahead of the cowl are studied numerically. Solutions of the Reynolds-averaged Navier–Stokes equations are obtained with a finite-volume solver of second-order accuracy on fine computational meshes. The solutions reveal jumps of the shock leg position with variations of the free-stream Mach number. The dependence of the shock position on the cowl slope and streamwise location of the throat is examined.

Numerical investigations on shock oscillations ahead of a hemispherical shell in supersonic flow

A clear understanding of the mechanism responsible for large amplitude shock pulsations ahead of a hemispherical cavity in supersonic flow is presented for the first time in this article. This has applications in supersonic parachute decelerators during the atmospheric descent stage of aerospace vehicles. A cell-centered finite volume code FaSTAR is used to solve the full Navier–Stokes equations on a hemispherical shell facing a Mach 4.0 supersonic free stream. The numerical method is validated against earlier experimental results. First, Flow Configuration A appears consisting of an axisymmetric shock that undergoes low-amplitude oscillations. This flow transitions to Flow Configuration B that has an asymmetric shock structure and undergoes large-amplitude non-stationary shock pulsations. The shock stand-off distance in Flow Configuration B is 1.65 times that in Flow Configuration A. The generation of vortices from the curved shock, amplification of vortices of one kind due to the dynamics of the cavity flow, and further interaction of these amplified vortices with the shock in a loop causes the large-amplitude shock pulsations. The oscillation frequencies as determined from cavity pressure and shock stand-off distance signals extracted from the unsteady results are 1.26 kHz during Flow Configuration A, and 859 and 863 Hz during the non-stationary pulsations of Flow Configuration B. The Helmholtz resonator model predicts quite accurately the frequency of Flow Configuration A (1.27 kHz), and to a good extent the frequency in Flow Configuration B (916.7 Hz).

液体燃料ロケットタンク破壊時の推進剤ジェット・主流の干渉メカニズムに関する数値解析

Numerical simulations of the jet injected into supersonic cross-flow are carried out. In aborting process in liquid rocket flight, the fuel is transversely injected into the supersonic freestream. Developed jet/freestream interaction structure may trigger unexpected auto-ignition or explosion events; thus the understanding of the flow-field of jet-in-supersonic cross-flow is essential to meet the reliability of rocket. The two-dimensional numerical simulations are carried out in a wide-range of jet and freestream conditions: the jet-to-freestream momentum flux ratio and the pressure ratio. Present result shows recirculation regions are formed both at the upstream and downstream of injector with the characteristic vortex structures. With increasing the jet-to-freestream momentum flux ratio, the vortex structures in the recirculation region are largely developed. It is founded that the highest temperature in the separation region depends on the total temperature in the separation region and furthermore, it is the lower temperature than the total temperature of freestream.

Numerical modeling of the conditions for realization of flow regimes in supersonic axisymmetric conical inlets of internal compression

The results of the numerical investigation of flow regimes in the axisymmetric inlets of internal compression at a supersonic flow around them are presented in the work. The main attention is paid to the determination of the ranges of the duct geometric convergence, in which a supersonic inflow in the inlet realizes. The investigation has been carried out at high supersonic freestream velocities corresponding to the Mach numbers М = 2−8 by the example of the frontal conical (funnel-shaped) inlets with the angles of the internal cone wall inclination δw = 7.5−15° under the variation of the relative area of the throat cross section. The flow structure alteration was studied and the critical relative areas of the inlet throat were determined, at which either there is no starting of the inlet in the process of flow steadying at the initial subsonic flow in it or a flow breakdown occurs in the process of flow steadying at an initial supersonic inflow. Numerical computations of the axisymmetric flow were done on the basis of the solution of the Navier–Stokes equations and the k-ω SST turbulence model.

Numerical Investigation of Jet Interaction for Missile with Continuous Type Side Jet Thruster

A continuous type side jet controller which has four nozzles with thrust control devices was considered. It is deployed to a missile for high maneuverability and fast controllability in the terminal guidance phase. However, it causes more complex aerodynamic jet interactions between the side jet and the supersonic free stream than does the conventional impulse type side jet with a small single thruster. In this paper, a numerical investigation of the jet interference effects for the missile equipped with a continuous type side jet thruster is presented. A three-dimensional flow field was simulated by using a commercial unstructured-based CFD solver. The numerical simulation method was validated through comparison with wind tunnel test results for the single jet. The method of defining jet direction for this type of side jet control to minimize simulation cases was also introduced. Flow fields investigation and jet interaction effects for various flow conditions, jet pressure ratios and defined jet direction conditions were performed. From the numerical simulation for the continuous type side jet, extensive aerodynamic interference data were obtained to construct an aerodynamic coefficients database for precise missile control.

Propulsive jet influence on generic launcher base flow

Afterbody flow phenomena represent a significant source of uncertainties in the design of a launcher. Therefore, there is a demand for measuring such flows in wind tunnels. For propulsive jet simulation a new jet facility was integrated into a hypersonic/supersonic wind tunnel. The jet simulation resembles the generic model of a staged rocket launcher. The design and the qualification of the facility are reported. This includes measurements of pressure, temperature and Mach number distribution. Pressure and Schlieren measurements are conducted in the wake of the generic launcher. The unsteady pressure characteristics at the generic rocket base and fairing are analyzed for supersonic and hypersonic freestream. The influence of the under-expanded jet is reported and the jet temperatures are varied. On the base fluctuations at a Strouhal number around 0.25 dominates supersonic freestream flows. Additionally, a fluctuation level increase on the base is observed for Strouhal numbers above 0.75 in hypersonic flow regime, which is attributed to the interactions of wake flow and jet.

Nonequilibrium Supersonic Freestream Studied Using Coherent Anti-Stokes Raman Spectroscopy

Measurements were conducted at the University of Virginia Supersonic Combustion Facility of the flow in a constant-area duct downstream of a Mach 2 nozzle. The airflow was heated to approximately 1200 K in the facility heater upstream of the nozzle. Dual-pump coherent anti-Stokes Raman spectroscopy was used to measure the rotational and vibrational temperatures of and at two planes in the duct. The expectation was that the vibrational temperature would be in equilibrium, because most scramjet facilities are vitiated air facilities and are in vibrational equilibrium. However, with a flow of clean air, the vibrational temperature of along a streamline remains approximately constant between the measurement plane and the facility heater, the vibrational temperature of in the duct is about 1000 K, and the rotational temperature is consistent with the isentropic flow. The measurements of vibrational temperature enabled cross-stream nonuniformities in the temperature exiting the facility heater to be documented. The measurements are in agreement with computational fluid dynamics models employing separate lumped vibrational and translational/rotational temperatures. Measurements and computations are also reported for a few percent steam addition to the air. The effect of the steam is to bring the flow to thermal equilibrium, also in agreement with the computational fluid dynamics.

Multiobjective Design Optimization of a Cantilevered Ramp Injector Using the Surrogate-Assisted Evolutionary Algorithm

AbstractThe mixing efficiency of the fuel and the supersonic freestream has a great impact on the operation of the scramjet engine because of the very short residence time. In the current study, the cantilevered ramp injector was used to promote mixing of the fuel and the supersonic flow, and the injector’s configuration was optimized by the surrogate-assisted evolutionary algorithm—namely, the nondominated sorting genetic algorithm coupled with the Kriging surrogate model. In the optimization process, the swept angle, the ramp angle, the compression angle, the length of the step, and the diameter of the injector were considered as the design variables, and the mixing efficiency, the total pressure recovery efficiency, and the drag force were taken as the objective functions. Finally, the Pareto front for the multiobjective design optimization results was obtained, and the relationship between the objective functions was explored. The results show that optimized drag force increases with increased mixing ...

Aerodynamic force by Lamb vector integrals in compressible flow

A new exact expression of the aerodynamic force acting on a body in steady high Reynolds number (laminar and turbulent) compressible flow is proposed. The aerodynamic force is obtained by integration of the Lamb vector field given by the cross product of vorticity times velocity. The result is obtained extending a theory developed for the incompressible case. A decomposition in lift and drag contribution is obtained in the two-dimensional case. The theory links the force generation to local flow properties, in particular to the Lamb vector field and to the kinetic energy. The theoretical results are confirmed analyzing numerical solutions obtained by a standard Reynolds Averaged Navier-Stokes solver. Results are discussed for the case of a two-dimensional airfoil in subsonic, transonic, and supersonic free stream conditions.

Partially-Averaged Navier–Stokes Simulations of High-Speed Mixing Environment

Accurate simulation of the fuel-air mixing environment is crucial for high-fidelity scramjet calculations. We compute the velocity fields of jet into supersonic freestream flow and cavity flow typical of scramjet flame-holding applications at different scale resolutions using the partially-averaged Navier–Stokes (PANS) method. We present a sequence of variable resolution computations to demonstrate the potential of PANS method for high-speed mixing environment calculations.

Multiobjective design optimization of the performance for the cavity flameholder in supersonic flows

The wall-mounted cavity makes a great difference to the flameholding mechanism in the supersonic flow, and it has been widely employed in the design process of the scramjet flowpath. However, the cavity would bring additional drag force to the engine, and the temperature within the cavity is very high. Therefore, both the drag force and the temperature should be minimized in the design process of the cavity flameholder. In the current study, the non-dominated sorting genetic algorithm (NSGA II) coupled with the Kriging surrogate model has been employed to optimize the nonreacting flow field around the cavity flameholder in the supersonic freestream with a Mach number of 3.0, and the test cases have been selected by the orthogonal table. At the same time, the numerical results have been compared with the experimental data obtained by Gruber et al., and four grid scales have been utilized to perform the grid independency analysis as well. The obtained results show that the wall static pressure profiles predicted by the numerical approaches show very reasonable agreement with the experimental data, and the static pressure along the floor face of the cavity is underpredicted. This may be induced by the inaccuracy of the turbulence model. The Pareto front for the multiobjective design optimization results is obtained, and there must be a compromise between the drag force and the area-weighted average temperature. The optimized drag force increases with the increase of the length-to-depth ratio and with the decrease of the swept angle, and the optimized area-weighted average temperature increases with the decrease of the length-to-depth ratio and with the increase of the swept angle. The drag force of the acoustic cavity is the smallest in the range considered in the current study. However, it would generate the strongest pressure oscillation.

Computational Aerodynamic Predictions of Supersonic Retropropulsion Flowfields

Supersonic retropropulsion, or the initiation of a retropropulsion phase at supersonic freestream conditions, is an enabling decelerator technology for high-mass planetary entries at Mars. Supersonic retropropulsion relies on retrothrust to decelerate the vehicle. A thrust-driven interaction of underexpanded jet flow with the shock layer of a blunt body alters the aerodynamic characteristics of the vehicle. Little literature exists on analytical and computational modeling approaches for supersonic aerodynamic-propulsive interactions at moderate thrust levels and flight-relevant conditions. This investigation is an exploratory application of a steady, turbulent computational approach to supersonic retropropulsion flowfields using modern and historical test cases. The results obtained from this approach agree reasonably well with experimental data for the locations of the bow shock, stagnation point, and Mach disk for a retropropulsion configuration with a single nozzle at the nose. The surface pressure distributions agree less favorably, showing a pressure rise toward the shoulder that was not observed in the original experiment. The flowfield structures for a peripheral retropropulsion configuration agree qualitatively with the expected structures, but proper inboard jet flow expansion and interactions between individual jet flows are likely not being captured; accordingly, pressure is preserved inboard of the nozzles to higher thrust than was observed experimentally.

Investigation of the turbulent wake flow of generic launcher configurations via a zonal RANS/LES method

A numerical analysis of the turbulent wake flow of a generic space launcher at supersonic freestream conditions (Ma∞ = 6.0 and ReD = 1.7 × 106) is performed using a zonal RANS/LES method. To investigate the influence of various components of a rocket model on the base flow, three supported wind tunnel configurations with the same main body geometry and different aft-body extensions consisting of a blunt base, a nozzle dummy, and a full flowing underexpanded TIC nozzle (Mae = 2.52, pe/p∞ = 100) are considered. Flow topologies for three cases are described in detail including an estimate of the impact of the wind tunnel model support on the flow field. To validate the applied numerical method, the computed flow fields are compared to experimental data from high-speed schlieren measurements provided by DLR Cologne. The influence of the used aft-body extensions on the steady-state and dynamic base flow characteristics is evaluated by a detailed analysis and comparison of the pressure distribution and its spectra along the base and nozzle walls for three investigated configurations. The numerical findings are compared to experimental wall pressure oscillation measurements provided by DLR Cologne. The major results are the non-negligible influence of the model support on the wake even on the strut averted side, the base drag reduction effect of the aft-expanding jet plume consisting of an increase of the base pressure level from \(p/p_\infty\approx0.2-0.25\) (blunt base and nozzle dummy configurations) up to \(p/p_\infty\approx0.7\) leading to a decrease of the base pressure drag coefficient from CDp base = 0.032 to 0.012 correspondingly, and the identified dominant low-frequency modes of the base pressure oscillations at SrD ≈ 0.05, SrD ≈ 0.1, and SrD ≈ 0.2 also detected in the experiments.

Progress in research on mixing techniques for transverse injection flow fields in supersonic crossflows

The transverse injection flow field has an important impact on the flowpath design of scramjet engines. At present a combination of the transverse injection scheme and any other flame holder has been widely employed in hypersonic propulsion systems to promote the mixing process between the fuel and the supersonic freestream; combustion efficiency has been improved thereby, as well as engine thrust. Research on mixing techniques for the transverse injection flow field is summarized from four aspects, namely the jet-to-crossflow pressure ratio, the geometric configuration of the injection port, the number of injection ports, and the injection angle. In conclusion, urgent investigations of mixing techniques of the transverse injection flow field are proposed, especially data mining in the quantitative analytical results for transverse injection flow field, based on results from multi-objective design optimization theory.

Analytical Modeling of Supersonic Retropropulsion Plume Structures

The propulsive-aerodynamic interactions of a vehicle undergoing supersonic retropropulsion are heavily dependent on the configuration of the nozzles and their integration with the vehicle. Knowledge of the interaction effects is gained from wind-tunnel results as well as computational fluid dynamics simulations. Because of the complex flowfield structures created by a vehicle employing supersonic retropropulsion, there are significant time and computational resource requirements to generate solutions for a particular configuration under specific flow conditions. Analytical flowfield models allow for preliminary information regarding flow structure to be available with minimal computational resources. By using these analytical models in the design process, the efficiency at which these studies can be performed can be improved by reducing the design space of nonfavorable designs and illustrating scenarios for higher-fidelity analysis. This paper proposes a method for analytically describing the plume boundaries present for two supersonic retropropulsion configurations: a single-nozzle and a three-nozzle configuration. The method is based on previous work on plume boundaries with modifications to account for the effects of a supersonic freestream. Comparisons with experimental data and computational simulations verify the model results for both configurations at Mach 2 and angle of attack of zero.

On an Acoustic Oscillation Energy for Shear Flows

The pressure perturbation and its derivative for an acoustic wave in a uniformly moving medium are canonical variables whose Hamiltonean is the acoustic energy. In the present paper is introduced a generalized acoustic Hamiltonean or oscillation energy that: (i) is conserved in a uniform flow, when it reduces to the acoustic energy; (ii) is not conserved in an homentropic unidirectional shear flow, in which case it has minimum at the critical layer where the Doppler shifted frequency vanishes. These properties are illustrated in the case of an unidirectional shear flow with an hyperbolic tangent velocity profile, representing a boundary layer over a flat rigid or impedance wall. The exact solution in the boundary layer can be matched to plane waves in the free stream, and reflection and transmission coefficients are defined for a incident wave of unit amplitude. If there is no critical layer in the boundary layer, a single powers series solutions covers the whole flow region from the free stream up to the wall; it is shown that for a rigid wall the modulus of the reflection coefficient is unity, but the argument is not zero, i.e. all sound is reflected with a phase shift. Also it is shown that, in this case, of a boundary layer without critical layer over a rigid wall, the phase of the transmission coefficient is one-half the phase the phase of the reflection coefficient, i.e. sound takes the same time to cross the boundary layer both ways, bouncing from the wall in between. In more general cases, viz. impedance wall and/or boundary layer with a critical layer, the modulus and phase of the reflection and transmission coefficients are calculated as functions of the angle of incidence, free stream Mach number and boundary layer thickness measured on a wavelength scale. When a critical layer is present in the boundary layer, the two pairs of solutions around it and the free stream must be matched, in order to obtain the acoustic field across the whole boundary layer, and calculate the reflection and transmission coefficients. The special case of critical layer at the free stream is also discussed. The cases discussed include boundary layers with and without a critical layer, subsonic and supersonic free streams with one zone of silence and one or two propagation zone(s), and rigid, inductive, reactive and mixed impedance walls. The generalized acoustic Hamiltonean or oscillation energy is also used to obtain a simple approximation for the modulus of the transmission factor, which is reasonably accurate not far from normal incidence.

Acoustic and velocity field over 3D cavities

Effects of wall mounted cavity on a Mach 1.7 freestream flow over it are investigated experimentally and numerically. Three different three dimensional (3D) cavity configurations have been used in the study. The results are compared with those of a two dimensional (2D) cavity. Flow field over the cavity is observed to depend intensely on the cavity width and on the allied aerodynamic flow structure in the vicinity of the cavity. Pressure oscillations generated by presence of wall mounted cavity in supersonic freestream was also observed to affect the fluid motion over cavities.