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Flow separation due to shock wave/boundary layer interaction is dominated in blade passage with supersonic relative incoming flow, which always accompanies aerodynamic performance penalties. A loss reduction method for smearing the passage shock foot via Shock Control Bump (SCB) located on transonic compressor rotor blade suction side is implemented to shrink the region of boundary layer separation. The curved windward section of SCB with constant adverse pressure gradient is constructed ahead of passage shock-impingement point at design rotor speed of Rotor 37 to get the improved model. Numerical investigations on both two models have been conducted employing Reynolds-Averaged Navier-Stokes (RANS) method to reveal flow physics of SCB. Comparisons and analyses on simulation results have also been carried out, showing that passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot for moderate adverse pressure gradient as well as suppression of boundary layer separations and secondary flow of low-momentum fluid within boundary layer. It is found that adiabatic efficiency and total pressure ratio of improved blade exceeds those of baseline at 95%-100% design rotor speed, and then slightly worsens with decrease of rotatory speed till both equal below 60% rated speed. The investigated conclusion implies a potential promise for future practical applications of SCB in both transonic and supersonic compressors.

The supersonic mode in hypersonic flows is studied numerically. The radiation characteristics of this mode in a Ma=6 flat-plate boundary layer with cooled walls are focused on. The linear stability...

Boundary layer development in a supersonic inlet is studied, focusing on the flow over the compression ramp and the ramp–isolator junction. Two designs of the junction, a backward-facing step (BFS)...

This paper reports an experimental study on the supersonic jet surface flow structure visualization and shear stress field measurement issuing from a rectangular nozzle with extended shelf. The evolution of the near-field surface flow structures with an increased Nozzle Pressure Ratio (NPR) is successfully captured by the surface oil flow, infrared detection technology, and the Shear-Sensitive Liquid Crystal Coating (SSLCC) technique. Results reveal that under smaller NPR, the wall flow structure is similar to that of a jet without the extended shelf i.e., clean jets, and this is caused by insufficient effect on the boundary layer. However, at higher amplitudes of NPR, there exists a significant effect of the boundary layer, as a near triangular separation forms on the trailing edge of the Mach stem due to the adverse pressure gradient, which is visualized for the very first time in this paper. Furthermore, the vector field of shear stress is measured quantitatively by SSLCC technique. Results shows that the magnitude of shear stress heightened with NPR increasing, and the directions of shear stress changes across the shock wave and expansion fans. In addition, surface streamlines measured by SSLCC is significantly consistent with the streamlines visualized using the oil flow technique.

The current experimental investigation explores the viscous/inviscid interaction of a supersonic turbulent boundary layer and a sharp fin-generated shock wave. The study examines the effect of Reyn...

We address supersonic boundary layer flow to slow acoustic waves using homotopy renormalization method. The model involves an ordinary equation system, which allows us to investigate the problem analytically and find asymptotic solutions in explicit form. At first, we rewrite the original problem in the form of a system of inhomogeneous variable coefficients homotopy equations and then handle them by the traditional perturbation theory, using renormalization group methods to eliminate the secular terms. We prove analytically and numerically the high accuracy of our solutions, and study in some details the effects of Mach number and wall temperature. Finally, we discuss how the explicit expression of boundary thickness makes our solutions suitable for practical applications. To the best of our knowledge, this is the first time that the HTR method is applied to supersonic boundary flow, and the analytic solutions are obtained. These solutions are easy to apply, and they could also help provide deeper insight into the supersonic boundary layer model.

Experiments show that a sudden expansion of supersonic flow may strongly stabilize or even relaminarize the boundary layer. Only few theoretical studies addressed this effect considering evolution ...

Abstract— The linear and nonlinear stages of disturbance development in the supersonic boundary layer over a 10° expansion corner is investigated numerically within the framework of Navier—Stokes equations for Mach number 3. The effect of sudden flow expansion on the disturbance evolution is analyzed. The flow stabilization effect observable in the aerodynamic experiment is also discussed.

A direct numerical simulation database (Zuo et al., J. Fluid Mech. vol. 877, 2019, 167–195) is analyzed to investigate the size of the separation bubble and the relation with the pressure rise in the conical shock wave/turbulent boundary layer interaction. The characteristics of the separation bubbles in the supersonic turbulent boundary layer induced by conical shock waves are discussed. The inviscid reflected conical shock theory is analyzed, to conclude that regular reflection transition to Mach reflection will occur somewhere along the spanwise direction in the interaction zone. Then, the typical shape of a turbulent separation bubble in the conical shock wave/turbulent boundary layer interaction is described. The scaling between the size of the turbulent separation bubble and moderate conical shock intensities is established, giving a potential alternative method to rapidly predict the size of turbulent separation bubbles, especially in the supersonic viscous corrections under condition of boundary layer separation.

In this paper, the effects of streamwise Nanosecond Dielectric Barrier Discharge (NS-DBD) actuators on Shock Wave/Boundary Layer Interaction (SWBLI) are investigated in a Mach 2.5 supersonic flow. In this regard, the numerical investigation of NS-DBD plasma actuator effects on unsteady supersonic flow passing a 14° shock wave generator is performed using simulation of Navier-Stokes equations for 3D-flow, unsteady, compressible, and k ‐ ω SST turbulent model. In order to evaluate plasma discharge capabilities, the effects of plasma discharge length on the flow behavior are studied by investigating the flow friction factor, the region of separation bubble formation, velocity, and temperature distribution fields in the SWBLI region. The numerical results showed that plasma discharge increased the temperature of the discharge region and boundary layer temperature in the vicinity of flow separation and consequently reduced the Mach number in the plasma discharge region. Plasma excitation to the separation bubbles shifted the separation region to the upstream around 6 mm, increased SWBLI height, and increased the angle of the separation shock wave. Besides, the investigations on the variations of pressure recovery coefficient illustrated that plasma discharge to the separation bubbles had no impressive effect and decreased pressure recovery coefficient. The numerical results showed that although the NS-DBD plasma actuator was not effective in reducing the separation area in SWBLI, they were capable of shifting the separation shock position upstream. This feature can be used to modify the structure of the shock wave in supersonic intakes in off-design conditions.

The interaction length induced by Shock Wave/Turbulent Boundary-Layer Interactions (SWTBLIs) in the hypersonic flow was investigated using a scaling analysis, in which the interaction length normalized by the displacement thickness of boundary layer was correlated with a corrected non-dimensional separation criterion across the interaction after accounting for the wall temperature effects. A large number of hypersonic SWTBLIs were compiled to examine the scaling analysis over a wide range of Mach numbers, Reynolds numbers, and wall temperatures. The results indicate that the hypersonic SWTBLIs with low Reynolds numbers collapse on the supersonic SWTBLIs, while the hypersonic cases with high Reynolds numbers show a more rapid growth of the interaction length than that with low Reynolds numbers. Thus, two scaling relationships are identified according to different Reynolds numbers for the hypersonic SWTBLIs. The scaling analysis provides valuable guidelines for engineering prediction of the interaction length, and thus, enriches the knowledge of hypersonic SWTBLIs.

In order to understand the physical phenomenon of the reflected shock/turbulent boundary layer interaction, the Large Eddy Simulation (LES) is conducted to investigate shock wave and turbulent boundary layer interaction in a 12 ° compression ramp with inlet high Mach number of 2.9. Rescaling/recycling method is used as inflow turbulence generation technique and validated on a supersonic flat plate turbulent boundary layer. The flow field of recycling plane in the plate computation domain is obtained to give the inlet boundary condition for the LES computation. This paper focuses on the reflected shock/turbulent boundary layer interaction region, where the fine flow structure and instantaneous flow field are analyzed in detail. It is found that the unsteady motion of the shock wave leads to the increase of wall pressure fluctuation.

Shock waves can significantly affect the film cooling for supersonic flow and shock waves may have different influence when impinging in different regions. The present study numerically compared the results of shock wave impinging in three different regions and analyzed the effect of impinging region. The shock wave generators were located at x/s = 5, 25, 45 with 4°, 7° and 10° shock wave incidence. The mainstream Mach number was 3.2 and the coolant Mach number was 1.2 or 1.5. The numerical results illustrated that the shock wave impinged in the further upstream region led to a larger high-pressure region and a larger vortex in the boundary layer. Moreover, placing the shock wave generator upstream resulted in the lower mass fraction of coolant in the downstream region. The velocity in the upstream part of the cooling layer was lower than the midstream and downstream part, which resulted in the less ability to resist the shock wave impingement. Therefore, the upstream impingement deteriorated the cooling performance to a greater extent. The study also manifested that the stronger shock wave had a larger effect on supersonic film cooling. Increasing the coolant inlet Mach number can increase the blowing ratio and reduce the mixing, which was of benefit to improve cooling effect.

The interaction of an impinging oblique shock wave with an angle of 30° and a supersonic turbulent boundary layer at Ma∞=2.9 and Reθ = 2400 over a wavy-wall is investigated through direct numerical simulation and compared with the interaction on a flat-plate under the same flow conditions. A sinusoidal wave with amplitude to wavelength ratio of 0.26 moves in the streamwise direction and is uniformly distributed across the spanwise direction. The influences of the wavy-wall on the interaction, including the characterization of the flow field, the skin-friction and pressure and the budget of turbulence kinetic energy, are systematically studied. The region of separation grows slightly and decomposes into four bubbles. Local peaks of skin-friction are observed at the rear part of the interaction region. The low-frequency shock motion can be seen in the wall pressure spectra. Analyses of the turbulence kinetic energy budget indicate that both diffusion and transport significantly increase near the crests, balanced by an amplified dissipation in the near-wall region. Proper orthogonal decomposition analyses show that the most energetic structures are associated with the separated shock and the shear layer over the bubbles. Only the bubbles in the first two troughs are dominated by a low-frequency enlargement or shrinkage.

Separated flows arising due to shock-wave/turbulent boundary-layer interactions can cause problematic low-frequency unsteadiness with potentially severe structural response. High-fidelity large-edd...

The paper aimed at studying the peculiarities of boundary layer on a rotating body. Rotation of axisymmetric bodies is often used for their stabilization at a high velocity motion in the atmosphere. Presented are calculation of boundary layer flow on a rotating cone at zero angle of attack. Longitudinal and full enthalpy profiles in the laminar boundary layer obtained as a result of calculations were used to determine upstream disturbances propagation velocity. New integral relation was determined to find out disturbances propagation velocity for the regime of weak hypersonic viscous-inviscid interaction. Effects of surface rotation velocity along with the temperature factor influence were investigated.

The influence of thermal nonequilibrium on the laminar–turbulent transition is studied with the use of the $$\mathrm{e}^N$$ -method for two widespread flow regimes in a supersonic boundary layer at the Mach number $$\textrm{M}=4.5$$ . The set of the actual frequencies of spatial disturbances is determined on the basis of the neutral curves for temporal disturbances. Families of the curves of $$N$$ -factors are calculated for selected frequencies. Then, based on the envelopes of these curves, the transition Reynolds number $$\mbox{Re}_{\delta T}$$ for a given transition factor $$N_T$$ is determined. The calculations show that the transition region in the case with $$N_T=8$$ and vibrational excitation level below the dissociation limit is located 12–14% downstream as compared to the transition region in a perfect gas.

The interaction between the shock wave and boundary layer and its induced separation is an undesirable phenomenon in the design of supersonic systems. The control of this phenomenon by active and passive methods is highly important in the design of aerodynamic systems. In the present work, different series of micro vortex generators were investigated as a passive control method (with the ability to increase the flow momentum), affecting the structure of a compressible flow at the Mach number of 2.05 in a low-aspect-ratio duct. In this regard, the finite volume method was used in the OpenFoam package. The obtained results were first validated against experimental data, and then four different vortex generators with a height of 10 μ m were investigated. The results revealed that the use of vortex generators caused a reduction in the shock wave strength and energy loss across the flow. Furthermore, applying a vortex micro-generator with the outer body of national advisory committee for aeronautics 4412 airfoil led to a significant decrease in the strength of shock waves. Also, a vortex generator could shift the shock wave occurrence location to the downstream.

A computational investigation of the role of thermal nonuniformity on the development of instability modes in the shear-layer of a supersonic M=1.5, Re=850,000 jet is performed in this study. Cold ...

In some cases, the nozzle of a rocket engine could deform during manufacturing or testing, flare up and distort during operation in such a way that periodically repeating structures of depressions and protrusions form on the wall surface. Complex configurations of roughness topologies have been experimentally observed to form due to this fluid-structure interaction and have been known to distort the boundary layer and increase local drag and heat transfer. A numerical experiment was carried out using ANSYS FLUENT to assess the effects of geometry (streamwise wavelengths ? and length-to-depth ratios (L/D)) of discrete roughness topologies in the form of steps and cavities on flow field distortions and resulting total drag. It was found that cavity L/D ratios of 19.8 – 23.1 and 12.8 – 15.5 yielded the highest drag coefficients. The flow fields over these periodic arrays exhibited closed and a combination of closed cavity type features, and it can be concluded that the flow phenomena associated with separation caused by forward facing steps in closed cavity type topologies contributed the most to the total drag. L/D ratios of 6.3 – 8.4 exhibited the lowest drag coefficient. The lower drag coefficient is due to the modification of the turbulent boundary layer by the trapped coherent vortical structure within the cavities. Pressure forces become more dominant as the surface wavelength is decreased. For all cases, there was a pronounced increase in the boundary layer thickness after each successive cavity.

In this paper, lattice Boltzmann simulation for unsteady shock wave/boundary layer interaction in a shock tube is presented. This kind of problem, which contains the typical characteristics of supersonic viscous flows with shock wave and boundary layer, has rarely been well validated in the lattice Boltzmann community. This leads to few application of lattice Boltzmann method in supersonic flows with boundary layer. Some problems related to numerical dissipation and stability are exposed when simulating shock wave/boundary layer interaction using the conventional compressible lattice Boltzmann approach. The implementations of lattice Boltzmann for this kind of flow are investigated. As a novel problem for the lattice Boltzmann method, three main issues need to be addressed carefully, i.e., lattice Boltzmann model, solving scheme, and boundary treatment. For the first issue, a double-distribution-function compressible lattice Boltzmann model, in which the lattice velocity is studied for the stability in the simulation, is employed since we have demonstrated its feasibility in supersonic viscous flow previously. For the solving scheme, various space discretization methods are tested to enhance numerical performance. Meanwhile, the boundary treatment for the ghost layer is studied. The simulation result by the improved lattice Boltzmann implementation agrees well with the results through Navier–Stokes equations and gas-kinetic scheme reported in previous literatures. This work probes the numerical implementations of lattice Boltzmann method for unsteady shock wave/boundary layer interaction, providing important theoretical support for lattice Boltzmann application in supersonic viscous flows with shock wave and boundary layer.

The bleed hole diameter, depth, and boundary layer thickness are key design parameters of a supersonic bleed system. The evolution trend of single-hole bleed flow coefficient with the ratio of boundary layer thickness to bleed hole diameter and the ratio of bleed hole depth to diameter is investigated by numerical simulations under choking and non-choking conditions. The results show that the subsonic leading edge of the circular hole and the subsonic part of the boundary layer are the main factors causing lateral flow of the bleed hole. The effect of diameter on bleed mass flow rate is due to the viscous effect which reduces the effective diameter. The larger the ratio of displacement thickness to bleed hole diameter, the more obvious the viscous effect is. The depth affects bleed flow rate by changing the opening and closing states of the separation zone. When a certain depth is reached, the development of the boundary layer reduces the effective captured stream tube and thus reduces the bleed mass flow rate. The main objective of the study is to obtain the physical mechanism of the bleed hole size parameters affecting the bleed mass flow rate, and to provide theoretical guidance for the selection of the size of bleed holes in the design of a porous arrays bleed system in hypersonic inlets.

The effect of wall expansion on the structural and statistical characteristics of wall-shear stress (WSS) fluctuations was investigated by direct numerical simulations of a supersonic turbulent bou...

In transonic flow, buffet is a phenomenon of flow instability caused by shock wave/boundary layer interaction and flow separation. The phenomenon is common in transonic flow, and it has serious impact on the structural strength and fatigue life of aircraft. In this paper, three typical airfoils: the supercritical OAT15A, the high-speed symmetrical NACA64A010, and the thin, transonic/supersonic NACA64A204 are selected as the research objects. The flow fields of these airfoils under pre-buffet and buffet onset conditions are simulated by Unsteady Reynolds Averaged Navier-Stokes (URANS) method, and the mode analysis of numerical results is carried out by Dynamic Mode Decomposition (DMD). Qualitative and quantitative analysis of the shock wave motion, shock wave intensity, shock foot bubble and trailing edge separation, and pressure coefficient fluctuation were performed to attain deep insight of transonic buffet flow features of different airfoils near buffet onset conditions. The results of DMD analysis show that the energy proportion of the steady mode of these airfoils decreases dramatically when approaching the buffet onset angle of attack, while the growth rate of the primary mode increases inversely. It was found that at the onset of buffet, there exist different degrees of merging behavior between shock foot bubble and trailing edge separation during one buffet cycle, and the instability of shock wave and separation induced shear layer are closely related to the merging behavior.

Shock-wave/boundary-layer interactions (SBLI) are an important feature of high-speed gas dynamics. In many numerical studies of SBLI span-periodicity is assumed to reduce computational complexity. However, span-periodicity is not a valid assumption for aeronautical applications such as supersonic engine intakes where lateral confinement leads to highly three-dimensional behaviour. In this work transitional oblique SBLI are simulated for a rectangular duct with a $$\theta _{sg} = 5^{\circ }$$ θ sg = 5 ∘ shock generator ramp at Mach 2. The baseline configuration is a duct with an aspect ratio of 0.5. Time-dependent disturbances are added to the base laminar flow via wall localised blowing/suction strips to obtain intermittent transition upstream of the SBLI. Two forcing configurations are evaluated to assess the response of the SBLI to different tripping locations. The transition is observed to develop first in the low-momentum corners of the duct and spread out in a wedge shape. The central separation bubble is seen to react dynamically to oncoming turbulent spots, shifting laterally across the span. While instantaneous corner separations do occur, the time-averaged corner flow remains attached. Comparisons to a one-to-one aspect ratio duct show that the SBLI is heavily dependent on aspect ratio; the wider duct exhibited significantly larger regions of flow-reversal due to a strengthened interaction.

A subscale flight experiment configuration propelled by a Mach 8 supersonic combustion ramjet (scramjet) was designed within the framework of the European Commission co-funded Long Term Advanced Propulsion Concepts and Technologies II project. The focus of this design exercise was to verify by ground testing the ability of the proposed scramjet engine to produce adequate thrust for hypersonic level flight. Experiments performed in DLR’s High Enthalpy Shock Tunnel Göttingen confirmed precedent CFD predictions of the total thrust and demonstrated the operability of the vehicle. Yet, significant discrepancies between the CFD analyses, which were performed to design the vehicle, and subsequent detailed measurements of the pressure distribution in the combustor were observed and could not be resolved so far. This paper focuses on a further analysis of these residual discrepancies. It was found that the CFD predictions of the combustor pressure distribution are sensitive to the configuration of the intake boundary layer. Particularly, different assumptions for the location of the laminar to turbulent boundary layer transition strongly influence cross-flow structures which develop on the intake and which are able to trigger different combustion modes in the combustor. While the effect on total vehicle performance remains limited, a significant impact on the structure and magnitude of the surface pressure distribution was observed. i.e., the large combustor peak pressures, which occur in the experiment, can be explained by the occurrence of a strong shock train in the vicinity of the combustor wall.

Theoretical investigation of the supersonic flat-plate boundary-layer properties under conditions of the surface material sublimation has been performed for Mach number M = 2. Naphthalene (C10H8) was chosen as the substance for the sublimation coating. Performed computations show that with increasing surface temperature due to stagnation temperature increase, the mass flow rate of naphthalene evaporation increases. Calculations performed on the basis of linear stability theory show that such an increase of evaporation leads to a noticeable decrease of the local growth rates of unstable perturbations in the boundary layer. It is found that stabilization of the boundary layer by the surface coating sublimation occurs with increasing temperature of the sublimation coating, reaching a maximum near the triple point temperature of the sublimation material. The carried out experiments confirmed the stabilizing effect of surface sublimation.

This study numerically investigated the influence of flow structures on the mixing characteristics of a sonic transverse gaseous ethylene jet in a supersonic crossflow with a Mach number of 2.95. RANS (Reynolds Average Navier-Stokes) simulations were employed to reveal the transport mechanism of jet fluid. The influencing factors, the jet-to-crossflow momentum flux ratio (J) and the nondimensional thickness of the turbulent boundary layer in the mixing characteristics of the sonic transverse gaseous jet are investigated. In the near-wall field, there is a V-shaped region in the front of the separation bubble that contains more jet gas under a lower momentum flux ratio (J) condition, which is caused by the collision of different inflow streamlines. As the turbulent boundary layer grows, there is more space for the jet gas to expand so that the penetration of the jet is much higher.

The interaction between self-sustaining dual jets and supersonic laminar boundary layer is experimentally studied using the nanoparticle-based planar laser scattering in a Ma 2.95 supersonic low-turbulence wind tunnel for the first time. The typical fine flow structures, such as separation shock, “W” bow shock, barrel shock, horseshoe vortex, large-scale vortex structures and counter-rotating vortex pairs, are clearly identified by streamwise and spanwise flowfield images. The curves of jet penetration depth are fitted at the momentum ratio of 0.55 between self-sustaining dual jets and supersonic crossflow. The results play a significant role in further and comprehensive exploration of the active flow control for a supersonic flow.

In supersonic wind tunnels, the airflow at the exit of a convergent-divergent nozzle is affected by the connection between the nozzle and test section, because the connection is a source of disturbance for supersonic flow and the source of disturbance generated by this disturbance propagates downstream. In order to avoid the disturbance, the test can only be carried out in the rhombus area. However, for the supersonic nozzle, the rhombus region is small, limiting the size and attitude angle of the test model. An integrated supersonic nozzle is a nozzle and a test section as a whole, which is designed to weaken or eliminate the disturbance. The inviscid contour of the supersonic nozzle is based on the method of characteristics. A new curve is formed by the smooth connection between the inviscid contour and test section, and the boundary layer is corrected for the overall curve. Integrated supersonic nozzles with Mach number 1.5 and 2 are designed, which are based on this method. The flow field is validated by numerical and experimental results. The results of the study highlight the importance of the connection about the nozzle outlet and test section. They clearly show that the wave system does not exist at the exit of the supersonic nozzle, and the flow field is uniform throughout the test section. Keywords: Boundary layer, Disturbance, Flow field, Integrated supersonic nozzle, Supersonic wind tunnels, The rhombus region

A direct numerical simulation study of self-similar compressible flat-plate turbulent boundary layers (TBLs) with pressure gradients (PGs) has been performed for inflow Mach numbers of 0.5 and 2.0. All cases are computed with smooth PGs for both favourable and adverse PG distributions (FPG, APG) and thus are akin to experiments using a reflected-wave set-up. The equilibrium character allows for a systematic comparison between sub- and supersonic cases, enabling the isolation of pure PG effects from Mach-number effects and thus an investigation of the validity of common compressibility transformations for compressible PG TBLs. It turned out that the kinematic Rotta–Clauser parameter $\unicode[STIX]{x1D6FD}_{K}$ calculated using the incompressible form of the boundary-layer displacement thickness as length scale is the appropriate similarity parameter to compare both sub- and supersonic cases. Whereas the subsonic APG cases show trends known from incompressible flow, the interpretation of the supersonic PG cases is intricate. Both sub- and supersonic regions exist in the boundary layer, which counteract in their spatial evolution. The boundary-layer thickness $\unicode[STIX]{x1D6FF}_{99}$ and the skin-friction coefficient $c_{f}$, for instance, are therefore in a comparable range for all compressible APG cases. The evaluation of local non-dimensionalized total and turbulent shear stresses shows an almost identical behaviour for both sub- and supersonic cases characterized by similar $\unicode[STIX]{x1D6FD}_{K}$, which indicates the (approximate) validity of Morkovin’s scaling/hypothesis also for compressible PG TBLs. Likewise, the local non-dimensionalized distributions of the mean-flow pressure and the pressure fluctuations are virtually invariant to the local Mach number for same $\unicode[STIX]{x1D6FD}_{K}$-cases. In the inner layer, the van Driest transformation collapses compressible mean-flow data of the streamwise velocity component well into their nearly incompressible counterparts with the same $\unicode[STIX]{x1D6FD}_{K}$. However, noticeable differences can be observed in the wake region of the velocity profiles, depending on the strength of the PG. For both sub- and supersonic cases the recovery factor was found to be significantly decreased by APGs and increased by FPGs, but also to remain virtually constant in regions of approximated equilibrium.

Numerical simulations and experiments showed that bump inlet had a remarkable effect on boundary layer diversion of supersonic flow. However, the design and analysis of bump in hypersonic flow was still few. In this paper, the mechanism of a supersonic bump inlet is introduced to the design of hypersonic forebody. A hypersonic inlet with an integrated bump/forebody is obtained by the Method Of Characteristics (MOC) based on a chin inlet. Numerical simulations show that the modified inlet achieves diversion of low-speed flow. Besides, the integrated bump/forebody is also beneficial to inlet start. During the starting process, the shape of the separation zone is rebuilt by the modified forebody surface which makes spillage much easier. This new design leads to a reduction of the self-start Mach number by 0.95. Keywords: Hypersonic inlets, Inlet start, Integration design, Method Of Characteristics (MOC), Separation zone

A Bayesian method to remove correlated noise from multi-channel measurements is introduced. It is based on Bayesian factor analysis coupled with prior but uncertain knowledge of the correlation structure of the noise. This technique is well suited to denoise cross-spectral matrices measured in the frame of aeroacoustic experiments when background noise measurements are available, because it allows separating the engine noise contribution from the turbulent boundary layer and uniform noise components that are all sensed by in-flow microphones. In-flight data measured on flush-mounted microphones on an aircraft fuselage are denoised using this method. It is shown that it has a significant benefit for studying the broadband shock-associated noise generated by the engines in realistic flight conditions.

International audience; In order to be able to judge the effectiveness of transition induction in WP-2, reference flow cases were planned in WP-1. There are two obvious reference cases—a fully laminar interaction and a fully turbulent interaction. Here it should be explained that the terms “laminar” and “turbulent” interaction refer to the boundary layer state at the beginning of interaction only. There are two basic configurations of shock wave boundary layer interaction and these are a part of the TFAST project. One is the normal shock wave, which typically appears at the transonic wing and on the turbine cascade. The characteristic incipient separation Mach number range is about M = 1.2 in the case of a laminar boundary layer and about M = 1.32 in the case of turbulent boundary layer. The second typical flow case is the oblique shock wave reflection. The most characteristic case in European research is connected to the 6th FP IP HISAC project concerning a supersonic business jet. The design speed of this airplane is M = 1.6. Therefore the TFAST consortium decided to use this Mach number as the basic case. Pressure disturbance at this Mach number is not very high and can be compared to the disturbance of the normal shock at the incipient separation Mach number mentioned earlier. As mentioned earlier, shock reflection at M = 1.6 may be related to incipient separation. Therefore two additional test cases were planned with different Mach numbers. ITAM conducted an M = 1.5 test case, and TUD an M = 1.7 test case. These partners have also previously made very specialized and successful contributions to the UFAST project.

Upon its development and initial characterization, the supersonic variant of the nanoscale thermal anemometry probe (S-NSTAP) was deployed in a supersonic wind tunnel facility, where both freestream and boundary layer measurements were obtained at $$M_\infty =2$$ . The low operating stagnation pressures generated reliable data, where the effects of Reynolds number, Mach number and overheat ratio on the sensor’s heat transfer were investigated in detail. The performance of the S-NSTAP was also compared to that of a conventional cylindrical hot-wire and the S-NSTAP was shown to exhibit unparalleled temporal resolution ( $$\sim $$ 300 kHz). The mass flux sensitivity coefficient of both hot-wires was further computed and appeared to vary between probes, yielding a coefficient twice as large for the conventional probe than for the S-NSTAP. The experimental data obtained from both hot-wires were also compared, via spectral analysis and turbulence statistics, to the results of a numerically modelled turbulent boundary layer.

The present paper investigates the receptivity of inviscid first and second modes in a supersonic boundary layer to time-periodic wall disturbances in the form of local blowing/suction, streamwise velocity perturbation and temperature perturbation, all introduced via a small forcing slot on the flat plate. The receptivity is studied using direct numerical simulations (DNS), finite- and high-Reynolds-number approaches, which complement each other. The finite-Reynolds-number formulation predicts the receptivity as accurately as DNS, but does not give much insight to the detailed excitation process, nor can it explain the significantly weaker receptivity efficiency of the streamwise velocity and temperature perturbations relative to the blowing/suction. In order to shed light on these issues, an asymptotic analysis was performed in the limit of large Reynolds number. It shows that the receptivity to all three forms of wall perturbations is reduced to the same mathematical form: the Rayleigh equation subject to an equivalent suction/blowing velocity, which can be expressed explicitly in terms of the physical wall perturbations. Estimates of the magnitude of the excited eigenmode can be made a priori for each case. Furthermore, the receptivity efficiencies for the streamwise velocity and temperature perturbations are quantitatively related to that for the blowing/suction by simple ratios, which are of $$O(R^{-1/2})$$ O ( R - 1 / 2 ) and have simple expressions, where R is the Reynolds number based on the boundary-layer thickness at the centre of the forcing slot. The simple leading-order asymptotic theory predicts the instability and receptivity characteristics accurately for sufficiently large Reynolds numbers (about $$10^4$$ 10 4 ), but appreciable error exists for moderate Reynolds numbers. An improved asymptotic theory is developed by using the appropriate impedance condition that accounts for the $$O(R^{-1/2})$$ O ( R - 1 / 2 ) transverse velocity induced by the viscous motion in the Stokes layer adjacent to the wall. The improved theory predicts both the instability and receptivity at moderate Reynolds numbers ($$R=O(10^3)$$ R = O ( 10 3 ) ) with satisfactory accuracy. In particular, it captures well the finite-Reynolds-number effects, including the Reynolds-number dependence of the receptivity and the strong excitation occurring near the so-called synchronisation point.

The development of supersonic combustion engines (scramjet-type) presents several challenges. Numerical simulations of scramjet engine combustion have become an attractive alternative to experimental investigations. However, supersonic combustion simulations still have several challenges: such as adequate modelling of shock/boundary layer and turbulence/chemistry interactions. The present work presents two large-eddy simulation probability density function (LES-PDF) novel formulations developed for high-speed applications. One is a conservative joint-scalar approach, where the joint-probability for reactive scalars and energy is solved, while the other is a joint velocity-scalar. The LES-PDF transport equations are solved using the Eulerian stochastic fields technique implemented in a density-based compressible solver. The performance of the models is verified through the simulations of two and three-dimensional supersonic reacting mixing layers and compared against DNS data from the literature. The results show that the joint-scalar formulation is accurate and robust, while the joint velocity-scalar closures require further development.

An experimental study was carried out to examine the unsteady pressure field on the base of a cylindrical afterbody at freestream Mach numbers ranging from high subsonic to low supersonic speeds. T...

In one of our earlier papers, we considered supersonic, laminar boundary-layer flows past bodies made of different materials and examined the possibility of controlling the temperature regime of a body immersed in flow. In the present work, for the conditions corresponding to aerodynamic experiments and to long flight times, we analyzed such values of deceleration parameters when regions of laminar, transitional, and turbulent regimes of flow are realized simultaneously in a boundary layer. The possibility of decreasing the level of maximal temperatures depending on the determining criteria of the problem has been evaluated for bodies made of different materials. The expediency of using simplified approaches is shown, including those for nonstationary conditions based on stationary solutions and on well-studied limiting cases of isothermal and adiabatic surface temperatures and allowing the determination of the temperature regimes in the entire range of thermophysical characteristics.

Shaped-hole film cooling is an essential cooling technology in modern gas turbine engines. The performance of a given film cooling design, and thus the life of the cooled part, is dependent on the scenario parameters that define the environment in which it operates, including Reynolds and Mach numbers, upstream boundary layer and turbulence characteristics, blowing ratio, density ratio, pressure gradient, surface curvature, etc. Of these, one parameter that has received relatively little attention in the open literature is Mach number. This work evaluates the role of Mach number in shaped-hole film cooling performance by comparing results of two large eddy simulations of a plenum-fed 7-7-7 hole cooling a flat plate at freestream Mach numbers of 0.25 and 0.5, with all other relevant scenario parameters fixed. This comparison shows important differences between the two cases. To begin, the performance, as measured by laterally-averaged adiabatic effectiveness, is dramatically different, with the high Mach case performing up to 50% worse. This decrease in performance at high Mach is due to the coolant jet separating from the surface and a large scale oscillation that leads to significant asymmetry in the time-averaged flow, features which are absent in the low Mach case. These changes appear to be caused by changes to the flow in the cooling hole, which is also asymmetric in the high Mach case. This asymmetry leads to entrainment of hot gas into the hole on one side. Further, in the high Mach case, the flow in the hole is mildly supersonic and has sequence of weak shocks, which are, of course, absent in the low Mach scenario.

The paper presents the experimental results and the results of direct numerical simulation of the development and interaction of two wave trains from two point sources of controlled disturbances in a supersonic boundary layer on a flat plate with an incident flow Mach number of 2.5. Sources were located parallel to the leading edge of the model. For the introduction of controlled disturbances into the boundary layer, the normal component of the mass flow rate was varied in the calculations. In the experiment, periodic glow discharges at a frequency of 20 kHz were used. In both cases, the disturbances sources worked synchronously. Mass flow rate pulsations were measured and recorded in sections, parallel to the leading edge of the model, near the maximum of disturbances along the boundary layer. In the experiment, a constant-temperature hot-wire anemometer was used. After performing a discrete Fourier transform, the spatial distributions of disturbances, the beta-spectra were determined, and the wave characteristics of the development of disturbances downstream were estimated. In addition, direct numerical simulation of the downstream development of disturbances from a single source was performed. The work presents a comparison of experimental and theoretical calculated data. The paper discusses the effects inherent in the interaction of unstable traveling controlled disturbances from two sources operating synchronously.

In this paper, we investigate the receptivity of a supersonic boundary layer to impinging acoustic waves. Unlike previous studies of acoustic receptivity, where the sound waves have phase speeds comparable with or larger than the free-stream velocity $U_{\infty }$, the acoustic waves here have much slower ($O(R^{-1/8}U_{\infty })$) phase velocity, and their characteristic wavelength and frequency are of $O(R^{-3/8}L)$ and $O(R^{1/4}U_{\infty }/L)$ respectively, compatible with the triple-deck structure, where $L$ is the distance to the leading edge and $R$ the Reynolds number based on $L$ and $U_{\infty }$. A significant feature of a sound wave on the triple-deck scale is that an $O(\unicode[STIX]{x1D700}_{s})$ perturbation in the free stream generates much stronger ($O(\unicode[STIX]{x1D700}_{s}R^{1/8})$) velocity fluctuations in the boundary layer. Two receptivity mechanisms are considered. The first is new, involving the interaction of two such acoustic waves and operating in a boundary layer over a smooth wall. The second involves the interaction between an acoustic wave and the steady perturbation induced by a wavy wall. The sound–sound, or sound–roughness, interactions generate a forcing in resonance with a neutral Tollmien–Schlichting (T–S) wave. The latter is thus excited near the lower branch of the neutral curve, and subsequently undergoes exponential amplification. The excitation through sound–sound interaction may offer a possible explanation for the appearance of instability modes downstream of their neutral locations as was observed in a supersonic boundary layer over a smooth wall. The triple-deck formalism is adopted to describe impingement and reflection of the acoustic waves, and ensuing receptivity, allowing the coupling coefficient to be calculated. The two receptivity processes with the acoustic waves on the triple-deck scale are much more effective compared with those involving usual sound waves, with the coupling coefficient being greater by a factor of $O(R^{1/4})$ and $O(R^{1/8})$ in the sound–sound and sound–roughness interactions, respectively. A parametric study for both the reflection and coupling coefficients is conducted for representative T–S waves, to assess the influence of the streamwise and spanwise wavenumbers, and the phase speed (or frequency) of the acoustic wave.

The paper is devoted to an experimental and theoretical study of effect of moderate angle-of-attack variation on disturbances evolution and laminar-turbulent transition in a supersonic boundary layer on swept wing at Mach 2. Monotonous growth of the transition Reynolds numbers with angle of attack increasing from −2° to 2.7° is confirmed. For the same conditions, calculations based on linear stability theory are performed. The experimental and computational results show a favourable comparison.

In order to investigate the applicability of the skin-friction reduction technique using hydrogen injecting into turbulent boundary layer, three-dimensional numerical simulation was carried out for a constant-cross-confined-space with rearward facing steps. The flow characteristics near wall surface and development of wall shear stress were analyzed and compared under different coming flow and injection conditions. The simulation results show that the hydrogen injection can achieve around 13.5% skin-friction drag reduction under the coming flow Mach number of 2.3Ma or 2.8Ma. At 2.8Ma, the optimal reduction profit is 13.5% which is obtained when the equivalent ratio is 0.06. The gases mixings are gradually enhanced along the flow path. At the positions of shock wave-boundary-layer interactions, the mixings are first strengthened and then suppressed, and meanwhile, the wall shear stress and density changes with similar law that first decreases and then rebounds at the positions. The declines of skin-friction drag decrease along the flow direction, the best reduction area can profit nearly 60%.

One of the current problems of the detached eddy simulation (DES) and delayed DES (DDES) methods is the inaccuracy in prediction of the separated shear layer flows, which mainly reflects in the delayed development of the shear layer. To correctly simulate the separated shear flow, various modified definitions of the DDES length scale are investigated and the improved definition of the subgrid eddy viscosity is introduced to deal with the delayed development problem for the first time. In this paper, the shear layer without the upstream boundary layer is studied in order to assess the subgrid model of the DDES in the simulation of shear layers. Large eddy simulation (LES) and experiment results are introduced to test the performance of different definitions of the DDES parameters. According to the results, both the definitions of the length scale and subgrid eddy viscosity coefficient should be improved to get the best prediction. DDES with the LES length scale and vorticity-correlated length scale as well as the modified definition of the subgrid eddy viscosity coefficient predict a more accurate development of the shear layer than DDES with the original definitions, which provides a possible solution to the delay development problem of the shear layer. Furthermore, the LES mode of DDES performs considerably differently from the Smagorinsky-LES, which shows the former subgrid model is essentially different from that of Smagorinsky-LES. Theoretical analyses are made in order to find out the quantitative difference of the subgrid eddy viscosity between the LES mode of DDES and Smagorinsky-LES.

Numerical simulations of three-dimensional flow of a gas mixture with chemical reactions over a flat thermally destructing material surface are presented. To create an effective numerical model, two ways of determining the heat removal are considered, and a comparison with the analytical solution obtained within the frame of the boundary layer approximation is represented. As an example, the numerical simulations of three-dimensional flow in combustion chamber of hybrid rocket are performed. The temperature maps and molar fraction of fuel within the chamber for different times are shown.

Results of direct numerical simulations of the roughness-induced development of instability and transition to turbulence in a supersonic boundary layer on a blunted cone for the free-stream Mach number M∞ = 5.95 are presented. The flow parameters and model geometry are consistent with the conditions of the experiments performed in the study. The following roughness types are considered: random distributed roughness, isolated roughness elements of different shapes, and a group of regularly arranged roughness elements. The processes of the instability development and transition for different roughness types are compared, and possible mechanisms of the roughness influence on the stability of boundary layers on blunt bodies are discussed.

In the present study, a flame flashback phenomenon inside an ethylene-fuelled scramjet combustor equipped with a cavity flameholder is investigated under flight Mach 5.5 condition. Experimental results exhibit the quasi-periodic combustion oscillation between the fuel injectors and the leading edge of the cavity under a specified condition. As an indispensable key sub-process of combustion oscillation, the flame flashback from the boundary layer downstream of the cavity is responsible for unsteady combustion process in scramjet combustors. Numerical simulation has been carried out in the specified condition. Attributed by (i) thicker boundary layer, (ii) closer thermal disturbance and (iii) improved local mixing degree, the flame flashback phenomenon can be induced by thermal throat which is generated by interaction between separated boundary layer and intense combustion. Quantitative analysis also indicates that the flame flashing is more sensitive to temperature fluctuation downstream of the cavity. In addition, a simplified combustion opening system model has been established to analyse combustion oscillation mechanisms, which theoretically demonstrates that the three factors mentioned above can destroy the balance of heat release and dissipation, causing the system cannot self-stabilise once certain temperature fluctuation thresholds in sensitive areas are exceeded. At the same time, the auto-ignition model excludes the possibility of flame flashback generated by auto-ignition effect.

The achievement of laminar flow in the boundary layer at high-speed cruise conditions may further, in addition to shock-wave control, reduce the drag and extend the range of military fighter aircraft. To this end, a further investigation on transitional boundary-layer flow of fighter wings is needed due to different configurations from the wings used on conventional transport aircraft. In this paper, wind tunnel experiments and numerical simulations were conducted on three-dimensional transition of thin diamond-shaped wings used on advanced fighter aircraft at tran/supersonic design points. A newly proposed correlation of crossflow transition which includes the effect of surface roughness was introduced into the γ-Reθt transition model. Predicted results were in good agreement with flow visualizations. Results showed that the strength of the crossflow component grew rapidly around the leading edge because of the severe flow acceleration, just as the same as wings with a large aspect ratio. However, there seemed no regular pattern of instability-dominance variation in span-wise for a diamond configuration. The dominance of different instability mechanisms strongly depended on the local pressure distribution. Hereby, the research recommended a “roof-like” shape of pressure distribution to suppress both crossflow and Tollmien-Schlichting (T-S) instabilities. Besides, a sharp suction peak with a serious pressure rise should be cut off to avoid stronger instabilities. Further discussions also revealed an independence of the unit Reynolds number when transition was triggered by T-S instabilities. Aerodynamic force comparisons indicated that further benefit on drag reduction could be expected by including the three-dimensional transition effect into a wing design process. Keywords: Boundary layer transition, Fighter aircraft design, Supersonic aircraft wings, Transonic wing aerodynamics, Wind tunnel measurements