Rinku Mukherjee

This paper investigates an improvement of the aerodynamic performance of a rectangular wing by re-designing its camberline to control the laminar separation of its boundary layer. This is experimentally implemented using an Aluminium secondary skin on the wing surface, which aligns itself to the separated boundary layer, such that the flow remains attached to it, which otherwise would have separated on the baseline configuration. The shape of the skin, which is now regarded as the effective flow surface, is essentially a decambered version of the baseline shape of the wing and is predicted numerically using an in-house code based on two linear functions that account for the local deviation of camber by accounting for the difference in coefficients of lift and pitching moments. Aerodynamic characteristics of the effective decambered configurations using numerical analysis, CFD, and wind tunnel experiments are reported. Results indicate that significant improvement in aerodynamic performance can be achieved for laminar separation control through this active flow surface.

The surface of a rectangular wing is morphed numerically at high angles of attack such that it still operates at the reduced coefficient of lift at which the baseline wing operates but while the flow is separated on the baseline wing, it remains attached on the morphed wing. The aerodynamic characteristics of the baseline wing are obtained experimentally and that of the morphed wing is obtained numerically. The morphed surface at high angles of attack is obtained using a novel ‘decambering’ technique, which accounts for the deviation of the coefficients of lift and pitching moment from that predicted by the potential flow. Two wings with different airfoil sections, N ACA0012 and N ACA4415 are tested and compared at high α. Numerical morphing of wing surface for design coefficient of lift (CL ) (in terms of percentage increment) is presented at angles of attack 50 and 150 . This concept of design CL of a morphing wing is one of the possible solutions to fly at different flight conditions with corresponding targets and maneuvering requirements. A significant addition to the present numerical approach is to provide some comparison of the flow separation behavior with CFD at the root section of both the wing sections. The effects of morphed surfaces on drag penalty, coefficient of lift and post-stall angles of attack are studied and compared in terms of aerodynamic performance.

A lumped vortex model coupled with a vortex dissipation and vortex core criteria is used to study the unsteady flow past two airfoils in configuration, each of which is impulsively set into motion. The unsteady wakes of both airfoils are modeled by discrete vortices and timestepping is used to predict the individual wake shapes. The coupled flow is solved using a combined "zero-normal flow" boundary condition and Kelvin condition which results in (2N +2)X(2N+2) equations. Results are presented showing the effect of airfoil-airfoil and airfoilwake interaction on the aerodynamic characteristics of the configuration for the airfoils moving with zero and non-zero relative velocity. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

This paper identifies the laminar separation bubble at the root or span-wise midsection of a rectangular wing using direct surface pressure measurements and analyses their behavior. The locations of separation, transition, and reattachment are identified from surface pressure measurements and oil flow visualizations. Surface oil flow visualization results also clarified the wing-tip and separation bubble interactions near the leading edge of the wing. The transition structure and turbulence characteristics in the separated shear layer locations are studied using Laser Doppler Velocimetry. Time series analysis is carried out to distinguish the flow patterns of transition and later transition locations along with chordwise locations of the root section of the wing.

Numerical simulation of unsteady motion of wing through undisturbed fluid is formulated using Unsteady Vortex Lattice Method. Singularities in the form of vortex ring elements are distributed over the surface of the wing mean camber and remain fixed in space. The wake vortex ring singularities are free to deform under the action of velocity field induced by the combination of moving wing and deforming wake vortex ring elements. CL, CDi, Г distribution etc., are calculated for different cases of unsteady motion. A free-wake algorithm is developed based on the higher order numerical integration schemes. Minimum cut-off distance to avoid r-1numerical singularity is obtained from Rankine vortex model. No-normal ow boundary condition is imposed over the surface of mean camber and the resulting MN × MN equations are solved for Г using LU decomposition.

A numerical iterative vortex lattice method is developed to study flow past wing(s) at high angles of attack where the separated flow is modelled using NY nascent vortex filaments. The wing itself is modelled using NX × NY bound vortex rings, where NX and NY are the number of sections along the chord and span of the wing respectively. The strength and position of the nascent vortex along the chord corresponding to the local effective angle of attack are evaluated from the residuals in viscous and potential flow, i.e. (Cl)visc - (Cl)pot and (Cm)visc - (Cm)pot. Hence, the 2D airfoil viscous Cl - α and Cm - α is required as input (from experiment, numerical analysis or CFD). Aerodynamic characteristics and section distribution along span are predicted for 3D wings at a high angle of attack. Effect of initial conditions and existence of multiple solutions in the post-stall region is studied. Results are validated with experiment.

A lumped vortex model coupled with a vortex dissipation and vortex core criteria is used to study the unsteady flow past tandem pitching airfoils. The unsteady wakes of both airfoils are modeled by discrete vortices and time-stepping is used to predict the individual wake shapes. The coupled flow is solved using a combined "zero-normal flow" boundary condition and Kelvin condition which results in (2N+2)X(2N+2) equations. Commercial software FLUENT is also used to study the flow tandem pitching airfoils. Results are presented showing the effect of airfoil-airfoil and airfoil-wake interaction on the aerodynamic characteristics of the tandem airfoils pitching in or out of phase and also when only the leading airfoil pitches and the trailing airfoil is stationary. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

A vortex-lattice numerical scheme that uses a novel decambering technique to predict post-stall aerodynamic characteristics is used to study the aerodynamics of tandem Cessna 172 aircrafts flying in echelon formation. Results like CL- α from the current method are compared with experiment. Additional results like section Cldistributions over wing spans, CM- α and CLW- αtw(i.e. CLof the leading aircraft for different angles of attack of the trailing aircraft) and analysis for (-) ve y-offset, which are available only from the current numerical method are reported that supplement the experimental results. Detailed post-stall numerical analysis and the effect of chord-wise, span-wise and vertical offsets on the aerodynamics of the formation are also reported.

Comparisons of the development of flow over a cylinder with a 20° cone nose and a cylinder with an ogive nose, which represent typical heat-shield configurations are studied using CFD and experiments at transonic Mach numbers. The Cp plots are studied to locate expansion or separation. Experiments are carried out at M = 0.8, 0.9, 0.95 and 1.1 and Re ≈ 2.45 × 106. Computations are carried out using the commercial package, FLUENT 6.3. Inadequate spatial resolution of pressure ports in experiments as well as limitations of the CFD tool result in some differences in experimental and CFD results.

A numerical iterative vortex lattice method is developed for post-stall flow past wing(s) where the separated flow is modeled using NY nascent vortex flaments. The wing itself is modeled using NX x NY bound vortex rings, where NX and NY are the number of sections along the chord and span of the wing respectively. The strength and position of the nascent vortex along the chord corresponding to the local e_ective angle of attack are evaluated from the residuals in viscous and potential ow, i.e. (Cl)visc– (Cl)potand (Cm)visc– (Cm)pot. Hence, the 2D airfoil viscous Cl– α and Cm– α is required as input (from experiment, numerical analysis or CFD). Aerodynamic characteristics and section distribution along span are predicted for 3D wings at post-stall angles of attack. Effect of initial conditions and existence of multiple solutions in the post-stall region is studied. Results are validated with experiment.