Evaluation of wave drag reduction by flow control

Aerospace Science and Technology, 2008

A detailed study of transonic flow over three-dimensional bumps has been conducted using experimental measurements and computational simulation. The aim of the investigation is to determine the flow characteristics over these potentially effective flow control devices for wave drag reduction for transonic aircraft wings. Careful qualitative and quantitative matching of the simulation and experimental conditions has considerably improved the agreement between them. In both the experiments and the computation, the shock position in the working section was found to be sensitive not only to the back pressure and the sidewall effects, but also to the incoming boundary layer characteristics. Two turbulence models, a modified Baldwin-Lomax model with enhanced performance in separated flow (curvature model) and a two-equation model (k-ω) have been implemented in the numerical simulation. Overall, both turbulence models gave reasonable results for the uncontrolled and controlled cases regarding the inviscid flowfield and the shock structure. In particular, a pair of streamwise vortices embedded in the boundary layer was captured by the two models. The traces in the surface oil flow pictures from the experiment also suggested the existence of the streamwise vortices. The combined surface and flowfield data provide some further insight into the flow physics on the shock control ramp bump, which is discussed in the paper. In addition, the study also demonstrates the enhanced capability of the algebraic model in capturing separated flow features with lower computational cost as compared to the k-ω model.

1987

A method is described for calculating unsteady transonic flow with viscous interaction by coupling a steady integral boundary-layer code with an unsteady, transonic, inviscid small-disturbance computer code in a quasi-steady fashion. Explicit coupling of the equations together with viscous -inviscid iterations at each time step yield converged solutions with computer times about double those required to obtain inviscid solutions. The accuracy and range of applicability of the method are investigated by applying it to four AGARD standard airfoils. The first-harmonic components of both the unsteady pressure distributions and the lift and moment coefficients have been calculated. Comparisons with inviscid calcualtions and experimental data are presented. The results demonstrate that accurate solutions for transonic flows with viscous effects can be obtained for flows involving moderate-strength shock waves.

2016

Transonic and high subsonic speeds are from the most critical aircrafts flight conditions due to the fact that flying in these speeds causes local supersonic pockets on their lifting surfaces which end up with a shock wave. Shock wave appearance results in drag increment and has the potential to cause flow unsteadiness and buffet. Shock wave control for large flight vehicles can reduce the fuel consumption and/or alternatively increases the flight range. Therefore, shock waves creation, their interaction with boundary layer and their control have been the subject of the wide range of studies. Surface mass transfer using fluid suction or injection is one of the active devices for the control of shock wave-boundary layer interaction as a means of airfoil drag reduction or aerodynamic performance improvement. A parametric study is carried out in order to gain insight into the effects of mass transfer parameters, i.e. the location, flow rate and the inclination angle on the aerodynamic ...

2004

The effects of the Contoured Bump and Surface Cooling flow control techniques on drag reduction and buffeting alleviation over aerofoils are investigated using CFD as a tool. In the range of transonic periodic flow that was found experimentally by McDevitt, Gibb and Okuno, the Tijdeman’s Types A, B and C shock-induced oscillations (SIO) have been numerically identified on NACA0012 and biconvex aerofoils. For biconvex aerofoils, at zero degrees incidence, the Contoured Bump, which was located on the lower surface, reduces the dynamic loads by up to 36% and drag by up to 15%. For NACA0012 aerofoil, at low angles of incidence (α=3.2°) and at Mach number of 0.7, the bump reduces buffeting by 60% and drag by 7.5%. The Surface Cooling has also a significant effect on the shock wave/boundary layer interaction and on skin friction at the surface of the aerofoil. For 18% and 14% biconvex aerofoils, in transonic periodic flow, at temperature rations Tw/T of 0.9 and 0.6 respectively, total dra...

This study proposes a flow over an airfoil which involves shock oscillations at certain critical Mach number in Transonic region. This is due to the shock wave interaction with airfoil boundary layer which results in shock buffeting. In the present study Energy equation has been used for the shock buffet and also the corresponding aerodynamic behaviour of NACA 0012 symmetric airfoil, NACA 2412 cambered airfoil, NASA SC (2)0714 supercritical airfoil, biconvex and double wedge airfoil. Energy computations have been performed at critical Mach number in the transonic region. The results obtained from the computational analysis have been compared with each other. Velocity contours and Mach number variation over the airfoils have been analysed and confirmed with the shock buffeting phenomena. The buffeting generated in airfoils reaches transonic Mach number with Zero AOA.

The paper focus on exclusive review in the field of Transonic flow over an airfoil. Various research and analysis has been done for the improvement and development of Airfoil in transonic regimes by means of Computational Fluid Dynamics (CFD), Direct Numerical Simulation (DNS) in form of analytical model and various experimental approach has been carried out by researchers in the field of transonic aerodynamics in order to overcome the trouble shoot occurs in form of instability and irreversibility during transonic and subsonic flight. During transonic flow Airfoil of Air vehicles exhibits shock wave in form of instability and If these shock waves are not been analyzed this may leads to tragic failure Since airfoils are subjected to both static and dynamic loads, due to which it pose great challenge to analyze the aerodynamic characteristics of an Airfoil.

Shock Waves, 2015

A shock control bump (SCB) is a flow control method which uses local small deformations in a flexible wing surface to considerably reduce the strength of shock waves and the resulting wave drag in transonic flows. Most of the reported research is devoted to optimization in a single flow condition. Here, we have used a multi-point adjoint optimization scheme to optimize shape and location of the SCB. Practically, this introduces transonic airfoils equipped with the SCB which are simultaneously optimized for different off-design transonic flight conditions. Here, we use this optimization algorithm to enhance and optimize the performance of SCBs in two benchmark airfoils, i.e., RAE-2822 and NACA-64A010, over a wide range of off-design Mach numbers. All results are compared with the usual single-point optimization. We use numerical simulation of the turbulent viscous flow and a gradient-based adjoint algorithm to find the optimum location and shape of the SCB. We show that the application of SCBs may increase the aerodynamic performance of an RAE-2822 airfoil by 21.9 and by 22.8 % for a NACA-64A010 airfoil compared to the no-bump design in a particular flight condition. We have also investigated the Communicated by A. Hadjadj.

CIT06-0594, 2006

This work is a direct numerical simulation of the transonic laminar flow in airfoils with high amplitude plunging motions. The problem is solved for a non-inertial system of reference that is moving with the airfoil, and for this reason, the associated pseudo-force is included as a source in the momentum equation and the work done is also included as a source in the energy equation. This methodology allows the solution of high amplitude plunging motions, since the problem is solved from a non-inertial frame of reference that is moving with the airfoil and, for this reason, no grid deformation is needed to account for the motion. The compressible Navier-Stokes equations are solved using the skew-symmetric form of Ducros’ shock-capturing algorithm, with fourth order accuracy in space and third-order accuracy in time. Five cases are studied: the static airfoil and the plunging motions with amplitudes of 2.5%, 13%, 22% of the airfoil chord. For all the cases, the Reynolds number is 10,000, the Mach number of the free flow is 0.8 and the plunging frequency has same value of the vortex emission frequency of the static case. The numerical results show a very complex and unsteady interaction between the boundary layer, the detached vortex wake and the transonic shock-wave system for the four cases studied. There are also some characteristic shock phenomena at the last two cases.

Physics of Fluids

Investigations have been performed via implicit large eddy simulations to study the overall effects of exciting a flow field by thermal (wall-heating and wall-cooling) and vortical (with high and low frequencies) actuation. The actuator is placed on the suction surface of a natural laminar flow (SHM-1) airfoil having an angle of attack of $\alpha = 0.38^o$ (cruise setting). Oncoming flow has a Mach number of $0.72$, and a Reynolds number based on chord of $Re = 16.2 \times 10^6$, for which a complex shock system is formed on the suction surface. Vorticity dynamics of the flow is studied using time series of vorticity at different locations above the suction surface and instantaneous contour plots of vorticity in the domain. An inspection of the flow using snapshots of $\nabla \rho $ and $\nabla (\rho T)$ is done to characterize the numerical Schlieren. Comparative effects of the various forms of excitation on the shock-boundary layer interactions (SBLI), have been analyzed using tim...

Fluid Mechanics and its Applications, 1999