Pavel PA

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Papers by Pavel PA

Response of the Shock Wave/Boundary Layer Interaction to Disturbances Induced by the Plasma Discharge

Aerospace

The paper focuses on the investigation of unsteady effects in shock wave/boundary layer interacti... more The paper focuses on the investigation of unsteady effects in shock wave/boundary layer interaction. The study was carried out using a flat plate model subjected to a free stream Mach number of 1.43 and a unit Reynolds number (Re1) of 11.5 × 106 1/m. To generate two-dimensional disturbances in the laminar boundary layer upstream of the separation region, a dielectric barrier discharge was employed. The disturbances were generated within the frequency range of 500 to 1700 Hz. The Strouhal numbers based on the length of the separation bubble ranged from 0.04 to 0.13. The measurements were carried out using a hot-wire anemometer. Analysis of the data shows that disturbances in this frequency range mostly decay. The maximum amplitudes of perturbations were observed at frequencies of 1250 Hz and 1700 Hz.

format_quoteLarge disturbances develop at frequencies of 1250-1700 Hz corresponding to St ≈ 0.1-0.13, with lower frequencies leading to diminished amplitudes.format_quote

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Effect of Blowing/Suction through Porous Surface at Supersonic Mach Number

The effect of blowing/suction through perforated surfaces on the turbulent boundary layer was stu... more The effect of blowing/suction through perforated surfaces on the turbulent boundary layer was studied experimentally at Mach number M = 1.4. In the experiments, several different porous surfaces were investigated in a wide range of the blowing/suction mass flow rate. The distributions of the integral parameters of boundary layer were obtained depending on the type of porosity and mass flow rate. Data analysis has showed a significant influence of shock waves generated near the holes on the turbulent boundary layer. The integral effect from shock waves on boundary layer depends on set of parameters, which makes search of optimal parameters of porosity difficult.

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Estimation of the Flow Parametrs in Hypersonic Wind Tunnels

<jats:p>The paper describes the algorithm of the flow parameters calculation for hypersonic... more <jats:p>The paper describes the algorithm of the flow parameters calculation for hypersonic wind tunnels taking into account the real gas properties using air and carbon dioxide as a working gas. The results of the experimental measurements of the flow velocity at the contoured nozzle exit in the hypersonic wind tunnel IT-302M have been carried out for verification of the algorithm</jats:p>

Effect of nose-tip roughness on laminar-turbulent transition

The effect of the distributed roughness applied to the blunted nose-tip of the cone on the positi... more The effect of the distributed roughness applied to the blunted nose-tip of the cone on the position of the laminar-turbulent transition is considered. Study is performed at Mach number 5.95. It was found that the position of the roughness plays an important role in the transition process. It was obtained that the roughness has a greater effect on the transition when covers area downstream of the sonic line.

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WP-1 Reference Cases of Laminar and Turbulent Interactions

In order to be able to judge the effectiveness of transition induction in WP-2, reference flow ca... more 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 thi...

Plasma Control of Transonic Shock Wave/Boundary Layer Interaction

The study is relevant to plasma flow control applications, namely, delay of buffet onset of trans... more The study is relevant to plasma flow control applications, namely, delay of buffet onset of transonic airfoils using plasma discharges. At transonic speeds the region of supersonic flow arises on the suction side of the airfoil ended by the strong shock wave. This leads to appearance of additional wave drag of the wing decreasing the aircraft performance in transonic regime. Moreover the interaction of the shock wave with boundary layer produces flow separation and induces flow instabilities. The flow oscillations called “transonic buffet” limit flight Mach number and wing loads reducing the flight efficiency. The wave drag and arising of the flow instability are strongly connected. Therefore, development of the flow control techniques capable of decreasing the wing wave drag and suppressing/delaying buffet oscillations is highly desirable for increasing flight efficiency of civil aircrafts.

Analysis of Upstream Conditions Effect on Shock Wave–Boundary Layer Interaction at Moderate Mach Number

Modern trends in transonic aviation are connected with reducing viscous drag. Therefore commercia... more Modern trends in transonic aviation are connected with reducing viscous drag. Therefore commercial airplanes of the next generation may be equipped with a laminar wing [1, 2]. Generic feature of the transonic flow near the airfoil is presence of local supersonic zone ended by a shock wave. Laminar boundary layer has weak resistance to adverse pressure gradients especially ones produced by shock waves, in contrast to the turbulent boundary layer [3]. This leads to significant flow separation that can eliminate all advantages of laminar flow. Therefore it is necessary to estimate the difference between steady and unsteady parameters of the separation zone for various positions of the laminar–turbulent transition relative to the zone of shock wave boundary layer interaction (SWBLI) for moderate Mach numbers. Flow separation on airfoil not only reduce the lifting force and increases drag, but also leads to additional adverse events, including buffeting [3]. It is well known that separat...

WP-2 Basic Investigation of Transition Effect

An important goal of the TFAST project was to study the effect of the location of transition in r... more An important goal of the TFAST project was to study the effect of the location of transition in relation to the shock wave on the separation size, shock structure and unsteadiness of the interaction area. Boundary layer tripping (by wire or roughness) and flow control devices (Vortex Generators and cold plasma) were used for boundary layer transition induction. As flow control devices were used here in the laminar boundary layer for the first time, their effectiveness in transition induction was an important outcome. It was intended to determine in what way the application of these techniques induces transition. These methods should have a significantly different effect on boundary layer receptivity, i.e. the transition location. Apart from an improved understanding of operation control methods, the main objective was to localize the transition as far downstream as possible while ensuring a turbulent character of interaction. The final objective, involving all the partners, was to b...

Experimental Study of Shock Wave / Turbulent Boundary Layer Interaction

Siberian Journal of Physics

Experimental study of separated flow in a zone of oblique shock wave / turbulent boundary layer i... more Experimental study of separated flow in a zone of oblique shock wave / turbulent boundary layer interaction was carried out for Mach number 2 and Reynolds number Reθ = 2,7÷3,5 × 103. Streamwise pressure distribution on the model surface was obtained, Schlieren and oil-flow visualizations were performed. The paper gives detailed data of hot-wire anemometry measurements in upstream boundary layer, interaction and recovery regions. Unsteady nature of separated zone and reflected shock wave was discovered. The effect of side walls on quasi 2D structure of separated flow is described.

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