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In considering the principles of the design of low‐drag aerofoils, we saw that we should like the velocity outside the boundary layer to rise to a position as far back along the wing as possible, but that we are hindered by the danger of turbulent separation if the rate of velocity decrease at the back of the aerofoil is too great; the danger increases, roughly, when the thickness of the aero‐foil increases and when the position of maximum suction is moved further back. If, however, the whole of the pressure recovery, or velocity decrease, is concentrated over a very narrow interval along the chord, over which the boundary layer, or as much of it as necessary, is sucked away to stop separation, all danger of separation is avoided, and we can have, if we wish, a favourable velocity gradient over the whole of the rest of the chord. More specifically, if the upstream boundary layer, or part of it, is sucked in through a slot, there must be a streamline that divides the air crossing the slot from that entering it, and that streamline must meet the surface at a stagnation point (fig. 10). With the stagnation point behind the slot the flow in the boundary layer along the surface is reversed in direction between the stagnation point and the slot, and all danger of separation is avoided with a well‐designed slot if the design of the aerofoil is such that the pressure recovery takes place wholly between the upstream lip of the slot and the stagnation point. Up to now suction aerofoils have actually been designed with a discontinuous drop in velocity; the velocity rises to a given position on the chord, drops discontinuously, and thereafter rises, stays constant, or at any rate docs not decrease rapidly enough to produce any danger of separation, to the trailing edge; the velocity distribution has this character on both the top and bottom surfaces.

This paper aims to describe a proposal for an innovative method of normal shock wave–turbulent boundary layer interaction (SBLI) and shock-induced separation control.

The concept is based on the introduction of a tangentially moving wall upstream of the shock wave and in the interaction region. The SBLI control mechanism may be implemented as a closed belt floating on an air cushion, sliding over two cylinders and forming the outer skin of the suction side of the airfoil. The presented exploratory numerical study is conducted with SPARC solver (steady 2D RANS). The effect of the moving wall is presented for the NACA 0012 airfoil operating in transonic conditions.

To assess the accuracy of obtained solutions, validation of the computational model is demonstrated against the experimental data of Harris, Ladson & Hill and Mineck & Hartwich (NASA Langley). The comparison is conducted not only for the reference (impermeable) but also for the perforated (permeable) surface NACA 0012 airfoils. Subsequent numerical analysis of SBLI control by moving wall confirms that for the selected velocity ratios, the method is able to improve the shock-upstream boundary layer and counteract flow separation, significantly increasing the airfoil aerodynamic performance.

The moving wall concept as a means of normal shock wave–turbulent boundary layer interaction and shock-induced separation control has been investigated in detail for the first time. The study quantified the necessary operational requirements of such a system and practicable aerodynamic efficiency gains and simultaneously revealed the considerable potential of this promising idea, stimulating a new direction for future investigations regarding SBLI control.

WRITING an introduction to an article by Mr S. B. Gates on Trailing‐Edge Flaps, which appeared in these columns in 1937, the Editor felt constrained to admit his bewilderment over the number and variety of types of high‐lift aid which then existed. Without intending any disrespect, I imagine that the progress of years must have added to his embarrassment. It has certainly added to the number of devices in use and under test.

The purpose of this paper is to study the effect of non‐uniform double slot suction (injection) into a steady laminar boundary layer flow over a yawed cylinder when fluid properties such as viscosity and Prandtl number are inverse linear functions of temperature. Non‐similar solutions have been obtained from the starting point of the streamwise co‐ordinate to the exact point of separation.

The governing equations are tackled by the implicit finite difference scheme in combination with the quasi‐linearization technique. Quasi‐linear technique can be viewed as a generalization of the Newton‐Raphson approximation technique in functional space. An iterative sequence of linear equations is carefully constructed to approximate the nonlinear equations for achieving quadratic convergence and monotonicity. The quadratic convergence and monotonicity are unique characteristics of the quasilinear implicit finite difference scheme, which makes this scheme superior to built‐in iteration of upwind or finite amplitude techniques.

The results indicate that the separation can be delayed by non‐uniform double slot suction and also by moving the slot downstream. However, the effect of non‐uniform double slot injection is just the opposite. Yaw angle has very little affect on the location of the point of separation.

This analysis is useful in understanding many boundary layer problems of practical importance for undersea applications, for example, in suppressing recirculating bubbles and controlling transition and/or separation of the boundary layer over submerged bodies.

The purpose of this paper is to study the effects of boundary‐layers bleeding on performance parameters of hypersonic inlets.

The inner flowfield of a hypersonic inlet at different bleeding rates is simulated with a Reynolds‐averaged Navier‐Stokes solver using a renormalization group k‐ε turbulence model.

In contrast with no bleeding, the performance parameter of hypersonic inlets without backpressure is reduced slightly, but the flow uniformity is improved. The interaction between boundary layers and shocks is weakened at the action of the bleeding, which leads to that the boundary‐layers separations at the entrance of the isolator caused by the high‐backpressure occur later, and it can improve the maximum backpressure ratio of hypersonic inlets. With the bleeding rate increasing, the maximum backpressure ratio of hypersonic inlets is added, while the total‐pressure recovery coefficient and mass‐captured coefficient are reduced.

This paper is a useful reference to the design and performance improvement of hypersonic inlets and propulsion systems.

This study aims to investigate the prevalence of shock boundary layer interaction (SBLI) in air-breathing intake system is highly undesirable since this leads to high pressure gradients, typical stream mutilation and pressure drop. A novel flow control mechanism is incorporated in this research holding an array configuration of passive flow control device (micro ramps [MR]) that is adapted to improve the boundary layer stability.

Two geometric variants of the MR, namely, MR40 and MR80 is considered which reduce the pressure drop during SBLI. The incidence oblique shock wave angle of 34° is considered for the modelling. Large eddy simulation (LES) turbulence model was used with subgrid models of Wall modelled LES, Smagorinsky–Lilly to compute the unsteady effects of SBLI control using micro vortex generators. The unsteady results are compared with steady Reynold’s average Naviers–Stoke’s equation for calibrating the turbulence models.

The array configuration of MR80 reduces the pressure drop by 22% as compared with no ramp configuration and also reduces the flow distortion in hypersonic inlet. The most affected region of the MR is in the vicinity of center-line. Quantitative results prove that the upstream influence of the shock waves has been largely reduces by MR80 array configuration as compared to single MR80 pattern configuration. Different vortex structures found in the experiments was exclusively predicted using LES.

This paper substantiates the requirement of MR array configuration for transferring the momentum from free stream to the boundary layer and thereby energizing the boundary layer. This process of energization delays the flow separation in hypersonic flow.

In this paper, the applicability of shear stress transport k-ω model along with the intermittency concept has been investigated over pitching airfoils to capture the laminar separation bubble (LSB) position and the boundary layer transition movement. The effect of reduced frequency of oscillations on boundary layer response is also examined.

A two-dimensional computational fluid dynamic code was developed to compute the effects of unsteadiness on LSB formation, transition point movement, pressure distribution and lift force over an oscillating airfoil using transport equation of intermittency accompanied by the k-ω model.

The results indicate that increasing the angle of attack over the stationary airfoil causes the LSB size to shorten, leading to a rise in wall shear stress and pressure suction peak. In unsteady cases, both three- and four-equation models are capable of capturing the experimentally measured transition point well. The transition is delayed for an unsteady boundary layer in comparison with that for a static airfoil at the same angle of attack. Increasing the unsteadiness of flow, i.e. reduced frequency, moves the transition point toward the trailing edge of the airfoil. This increment also results in lower static pressure suction peak and hence lower lift produced by the airfoil. It was also found that the fully turbulent k-ω shear–stress transport (SST) model cannot capture the so-called figure-of-eight region in lift coefficient and the employment of intermittency transport equation is essential.

Boundary layer transition and unsteady flow characteristics owing to airfoil motion are both important for many engineering applications including micro air vehicles as well as helicopter blade, wind turbine and aircraft maneuvers. In this paper, the accuracy of transition modeling based on intermittency transport concept and the response of boundary layer to unsteadiness are investigated.

As a conclusion, the contribution of this paper is to assess the ability of intermittency transport models to predict LSB and transition point movements, static pressure distribution and aerodynamic lift variations and boundary layer flow pattern over dynamic pitching airfoils with regard to oscillation frequency effects for engineering problems.

The purpose of this paper is to investigate and compare the effects of rod and vane-type vortex generators for wind turbine applications. In large wind turbine rotors, an attached flow at all sections along the span direction is difficult to achieve which leads to an increase in aerodynamic losses, noise generation, and fatigue stress. Therefore, flow control strategies such as vortex generators (VGs) are beneficial to improve performance.

The benefits of the application of rod-type vortex generators (RVGs) to control flow separation on a wind turbine airfoil are assessed numerically using computational fluid dynamics (CFD). The validation of the computational model is conducted against the experimental data available for the DU96-W-180 wind turbine airfoil equipped with 44 RVGs. In addition, a revised wind tunnel angle of attack (AoA) calibration procedure (based on CFD) is proposed that is applicable for separated flows. A comparison of the RVGs to the conventional vane-type vortex generators (VVGs) is presented for inflow velocity of 30 m/s and AoA leading to significant flow separation. A parametric evaluation of the geometric characteristics of both types of VGs is conducted to quantify the generated streamwise vortices.

The comparison of the induced flow structures and aerodynamic efficiency enhancements proves that RVGs may be used as an alternative to the more conventional VVGs applied on wind turbine blades for boundary layer separation control.

A new type of VG (rod) has been investigated and compared against conventional VG (vanes) for wind turbine applications.

The purpose of this paper is to study the steady mixed convection boundary layer flow of a nanofluid past a horizontal circular cylinder in a stream flowing vertically upwards for both cases of a heated and cooled cylinder.

The resulting system of nonlinear partial differential equations is solved numerically using an implicit finite‐difference scheme known as the Keller‐box method. This method is very efficient for solving boundary layer problems.

The solutions for the flow and heat transfer characteristics are evaluated numerically for various values of the parameters, namely the nanoparticle volume fraction φ and the mixed convection parameter λ at Prandtl number Pr=1 and 6.2. Three different types of nanoparticles considered are Cu, Al₂O₃ and TiO₂ by using water‐based fluid with Pr=6.2. It is found that for each particular nanoparticle, as the nanoparticle volume fraction φ increases, the skin friction coefficient and heat transfer rate at the surface also increase, and it also leads to the increment of the value of mixed convection parameter λ which first gives no separation.

The results of this paper are valid only up to the value of λ=λ₀ (<0) below which a boundary layer solution does not exist.

The results obtained can be used to explain the characteristics and applications of nanofluids, which are widely used as coolants, lubricants, heat exchangers and micro‐channel heat sinks. Nanofluids usually contain the nanoparticles such as metals, oxides, or carbon nanotubes, whereby these nanoparticles have unique chemical and physical properties.

The results of this paper are important for the researchers working in the area of nanofluids. The paper is well prepared and presented. The results are original, new and important from both theoretical and application point of views.

Reynolds number in small-size low-pressure turbines (LPT) can drop below 2.5 · 10⁴ at high altitude cruise, which in turn can lead to laminar boundary layer separation on the suction surface of the blades. The purpose of this paper is to investigate the potential of an alternate current (AC)-driven Single Dielectric Barrier Discharge Plasma Actuator (AC-SDBDPA) for boundary layer control on the suction side of a LPT blade, operating at a Reynolds number of 2 · 10⁴.

Experimental and numerical analyses were conducted. The experimental approach comprised the actuator testing over a curved plate with a shape designed to reproduce the suction surface of a LPT blade. A closed loop wind tunnel was employed. Sinusoidal voltage excitation was tested. Planar velocity measurements were performed by laser Doppler velocimetry (LDV) and particle image velocimetry (PIV). The device electrical power dissipation was also calculated. Computational fluid dynamics (CFD) simulations using OpenFOAM© were conducted, modelling the actuator effect as a body force calculated by the dual potential algebraic model. Unsteady RANS (Reynolds Averaged Navier-Stokes equations), also known as URANS approach, with the k-ε Lam-Bremhorst Low-Reynolds turbulence model was used.

The AC-SDBDPA operation brought to a reduction of the separation region; in particular, the boundary layer thickness and the negative velocity values decreased substantially. Moreover, the flow angle in both the main flow and in the boundary layer was reduced by the plasma control effect. The actuation brought to a reduction of the 17 per cent in the total pressure loss coefficient. The pressure coefficient and skin friction coefficient distributions indicated that under actuation the reattacnment point was displaced upstream, meaning that the flow separation was effectively controlled by the plasma actuation. Adopting slightly higher actuation parameters could bring to a full reattachment of the flow.

The work underlines the potentialities of an AC-SDBDPA to control separation in LPTs of aeroengines.

The present work sets a methodological framework, in which the validated procedure to obtain the body force model combined with CFD simulations can be used to study a configuration with multiple actuators allocated in array without requiring further experiments.