Distributions of the amplitude of controlled disturbances in space and time and their frequency-wave characteristics are obtained from experimental results on weakly nonlinear development of the wave train in the region of a stationary wake inside the boundary layer on a flat plate at the Mach number M = 2. A stationary streamwise disturbance is generated by a pair of weak oblique shock waves. Controlled disturbances are inserted into the flow by a local high-frequency glow discharge located inside the model. The development of controlled disturbances is analyzed on the basis of the linear theory of hydrodynamic stability. Typical resonant wave triplets are identified. It is found that flow inhomogeneity suppresses the mechanisms of interaction of controlled disturbances.

The paper discusses experimental results on the wave train development in a longitudinal trace in the flat plate boundary layer at Mach 2.5. An analysis of the spatiotemporal distributions and frequency-wave spectra of pulsations, as well as their wave characteristics in the linear and weakly nonlinear phase of the wave train development in a homogeneous and inhomogeneous boundary layer at a fixed power of the controlled disturbances source has been carried out. During the analysis of the results, no subharmonic resonance was observed. Expansion into the wave spectrum of a stationary in- homogeneity and experimental data on wave characteristics and disturbance spectra made it possible to propose options for wave interaction for oblique breakdown.

The results of experimental measurements are presented to assess the correlation characteristics of disturbances in the boundary layer of a flat plate with a sharp leading edge and pulsations of the incoming flow in cases of natural pulsations and influence of an N-wave for Mach 2 on the leading edge. The parameters of the cross-correlation of signals from two constant-temperature anemometry (CTA) are obtained digitally. The frequency ranges with apparent correlation between the mass flow rate pulsations measured by the probes are determined.

A passive control method for supersonic boundary-layer transition on a swept wing using longitudinal roughness is proposed. Tests were carried out to examine the effect of distributed roughness on the development of flow peturbations and on the laminar-turbulent transition. The method makes it possible to manipulate the transition in a wide range, bringing it either closer to the leading edge of the wing by 30 % or delaying it by 40%.

The action of a pair of weak shock waves on a supersonic boundary layer on a swept flat plate with a bluntness radius of the leading edge equal to 2.5 mm at the Mach number 2 is experimentally studied. Transverse hot-wire measurements are performed in the boundary layer with a fixed distance of the probe from the model surface. It is found that a change in the sweep angle of the leading edge from 35 to 45 degrees reduces the intensity of the action of the weak shock waves on the boundary layer flow. As the sweep angle of the leading edge increases to 50°, the weak shock waves no longer affect the flow in the supersonic boundary layer on the swept plate.

The present study is aimed at searching for mechanisms of wave interaction in a supersonic boundary layer on a flat plate with the flow being distorted by a streamwise disturbance generated on a sharp leading edge by a pair of weak shock waves. The space and time distributions and the frequency-wave spectra of oscillations and their wave characteristics in the linear and weakly nonlinear phases of wave train development in a nonhomogeneous boundary layer under conditions of a fixed power of the local source of controlled disturbances are analyzed. Based on experimental results, possible variants of wave interaction in a supersonic boundary layer on a flat plate in the region of the longitudinal wake generated by an N-wave are proposed.

This work is devoted to the study of the influence of the small angles of attack and external disturbances on the position of the laminar-turbulent transition in the supersonic boundary layer on swept wings. The experiments were carried out in a low-noise supersonic wind tunnel T-325 of the ITAM SB RAS at Mach numbers M = 2 and M = 2.5 on two wing models with leading edge swept angles χ = 45˚ and χ = 72˚. The measurements were carried out by the hot-wire probe or Pitot tube. External disturbances were generated by a wire placed ahead of the nozzles in the subsonic part of the flow. It was shown that external disturbances and small angles of attack strongly affect the position of the laminar-turbulent transition. The laminar-turbulent transition on a wing model with a subsonic leading edge occurs earlier than on a wing model with a supersonic leading edge at low noise conditions. The influence of leading edge swept angle χ on transition Reynolds number was not observed at noisy conditions.

We present the new (for Mach numbers М = 3 and 3.5) and generalizing (for Mach numbers from 2 to 4) results of experimental investigations on the effect of small angles of attack on laminar-turbulent transition in the supersonic boundary layer on a swept wing with the leading-edge slip angle of 72°. The angle-of-attack variation has a strong effect on the transition Reynolds number. The transition Reynolds number decreases with increase in the Mach number. The measurements were carried out by means of a constant-temperature hot-wire anemometer using the proven procedure of determining the transition location. The eN method is used for the first time for numerically estimating the transition Reynolds numbers in the supersonic boundary layer on a swept wing with the leading-edge slip angle of 72°. The growth of the amplitudes of the steady and unsteady modes of the boundary layer crossflow are calculated in accordance with the linear stability theory, within the framework of the Lees–Lin system of equations. The numerical results indicate that, in accordance with the experimental results, laminar-turbulent transition in the boundary layer on the model swept wing is governed by the growth of stationary modes of the crossflow instability.

Stationary disturbances in a free supersonic flow caused by two-dimensional roughness in a turbulent boundary layer on a wall were investigated using numerical simulation in the FlowVision software package. Calculations were performed for cases where the roughness heights were significantly smaller than the thickness of the boundary layer. In order to reduce the numerical oscillations, the computational grid has been adapted to the perturbation fronts by means. It was found that a disturbance in the form of a small amplitude N-wave is formed in the flow above the boundary layer. The effect of roughness height on disturbances formed in a free flow was studied. It was found that as the roughness height increases, there is an observed increase in the amplitude and spatial scale of the disturbance. It was also found that the gradient of the flow parameters between the disturbance fronts remains practically unchanged for all roughness heights considered. The numerical results were verified with experimental data. A strong agreement was achieved between the simulation and experimental result.

Views Icon Views Article contents Figures & tables Video Audio Supplementary Data Peer Review Share Icon Share Twitter Facebook Reddit LinkedIn Tools Icon Tools Reprints and Permissions Cite Icon Cite Search Site Citation A. A. Yatskikh, A. D. Kosinov, N. V. Semionov, Yu. G. Ermolaev, L. V. Afanasev, V. L. Kocharin, A. V. Shmakova; Experimental studies of pulsations of the boundary layer of swept wings with periodic roughness at M = 4. AIP Conference Proceedings 16 February 2023; 2504 (1): 030068. https://doi.org/10.1063/5.0132320 Download citation file: Ris (Zotero) Reference Manager EasyBib Bookends Mendeley Papers EndNote RefWorks BibTex toolbar search Search Dropdown Menu toolbar search search input Search input auto suggest filter your search All ContentAIP Publishing PortfolioAIP Conference Proceedings Search Advanced Search |Citation Search

Results of experimental studies are applied to compare the spatiotemporal amplitude distributions of controlled fluctuations in a linear and a weakly nonlinear phase of the development of a wave train in a homogeneous and an inhomogeneous boundary layer on a flat plate at a Mach number M = 2. (The flow is inhomogeneous due to a longitudinal steady perturbation generated by a pair of weak shock waves. Controlled perturbations are generated locally from the surface at a fixed power of a high-frequency glow discharge inside the model.) It is revealed that the inhomogeneity of the flow changes the mechanism of interaction of subharmonic perturbations. It is shown that the spatial synchronism of fluctuations of subharmonic frequency is violated in the center of the wave train.

The work is devoted to the development of a new experimental method for entering controlled disturbances with a given frequency-wave structure into a supersonic boundary layer. Experimental data on the formation of disturbances from two pulsed sources (pulsed glow discharge) in the laminar flat-plate boundary layer at the Mach number equal to 2 are given. The experiments were carried out in the T-325 wind tunnel of ITAM SB RAS. The localized sources were spaced out at the same distance from the leading edge of the plate at 6 mm from each other spanwise. Flow pulsations were measured using a hot-wire probe of constant temperature anemometer, the signal was recorded synchronously with ignition of the discharges. This made it possible to distinguish the pulsations from the discharges against the background of random uncontrolled “natural” pulsations of the boundary layer. The spatial-temporal and frequency-wave structure of the generated disturbances from a single or two discharges, operating synchronously or with a time delay, are analyzed. It is found that the maximum difference in the structure of disturbances from one and two sources is observed in the central region, while at the side boundaries of the disturbances, the pulsations are close in all considered cases. In the wave spectra of disturbances from two discharges, nodes and antinodes are formed. Their position is determined by the distance between the sources and the time delay in their operation.

An experimental study of the influence of the leading-edge radius of a blunt plate on the response of boundary layer on this plate to an N-wave at Mach 2 was carried out. Three flat-plate models with different leading-edge bluntness radii, r = 0.05, 0.5, and 2.5 mm, were used in the experiments. The oncoming-flow disturbances were created using a generator provided on the sidewall of the test section of the T-325 wind tunnel of ITAM SB RAS. It is found that behind the N-wave in the oncoming flow, there forms an extended region with an increased level of flow pulsations whose spectrum contains the amplitudes increased both in low-frequency and high-frequency range of the spectrum compared to the undisturbed freestream. It was shown that, under experimental conditions, the flow non-uniformity generated by the N-wave can exert a greater influence on the flat-plate boundary layer laminar-turbulent transition with increasing leading-edge bluntness radius.

Views Icon Views Article contents Figures & tables Video Audio Supplementary Data Peer Review Share Icon Share Twitter Facebook Reddit LinkedIn Tools Icon Tools Reprints and Permissions Cite Icon Cite Search Site Citation L. V. Afanasev, A. D. Kosinov, A. A. Yatskikh, N. V. Semionov, Yu. G. Yermolaev, V. L. Kocharin; Experimental study of the relationship between the pulsations of the oncoming flow and the disturbances of the boundary layer of a sharp plate at Mach number M=2.5. AIP Conference Proceedings 16 February 2023; 2504 (1): 030001. https://doi.org/10.1063/5.0132410 Download citation file: Ris (Zotero) Reference Manager EasyBib Bookends Mendeley Papers EndNote RefWorks BibTex toolbar search Search Dropdown Menu toolbar search search input Search input auto suggest filter your search All ContentAIP Publishing PortfolioAIP Conference Proceedings Search Advanced Search |Citation Search

Experimental results of the interaction of disturbances from two point sources of periodic controlled pulsations in the supersonic boundary layer of a swept wing at a Mach number of 2 are presented. It is shown that as a result of the summation of disturbances, an interference pattern is formed and, for certain values of the wave numbers, it is possible to suppress the corresponding components of unstable waves.