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 uses a numerical post-stall predictive tool based on ‘decambering’ approach to study the aerodynamic characteristics of a lead-trail formation in pre and post-stall flow conditions. A basic lead-trail formation consisting of 2 wings and an extended formation consisting of 5 wings are studied with a view to the possibility of fuel savings, increase in range of operation, delayed flow separation and efficient positioning of the wings with respect to each other. Whether increasing the number of wings in a configuration is more useful is also looked into. The optimum operational angles of attack for maximum advantage in terms of fuel efficiency of all wings is studied including post-stall angles of attack. Numerical results for CL, CDi, section Cl distribution and their dependence on vertical offsets and angle of attack are reported.

Blunt-nosed launch vehicles featuring large nose-cone angles experience high levels of pressure fluctuations over the payload region at transonic Mach numbers due to shock-wave/boundary-layer interactions. The main cause for this phenomenon appears to be flow instability associated with separated flow and formation of a vortex pair. Interactions between the induced velocity of the vortex pair and oncoming mean flow cause oscillations of the λ-shock system and high levels of fluctuating pressures. An aerospike causes flow separation at the nose and reattachment of the shear layer downstream, energizing the boundary layer. Consequently, flow separation and vortex formation are prevented: shock oscillations are stabilized. Dramatic reductions in the pressure fluctuations, around 95–35% in the low-frequency range, are observed along the payload region at small angles of attack. The observations are based on wind-tunnel tests involving unsteady pressure measurements, surface-flow patterns, and high-speed shadowgraph recordings on a blunt nose cone with various cone angles.

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.

Two-dimensional Direct Numerical Simulation (DNS) of Rayleigh–Bénard Convection of Jet-A fluid (Pr=19.17) is performed in a square cavity. The Rayleigh number of this study is Ra=107 with a temperature difference ΔT=40 K between hot and cold plates to account for non-Oberbeck–Boussinesq (NOB) effect. The Jet-A fluid has higher viscosity and lower thermal conductivity compared to water and the Prandtl number is approximately 4.4 times higher than water at mean temperature of 40 °C. The fluid properties of the working fluid are defined as a polynomial function of temperature. The Boundary layer thicknesses, the Centre line temperature, Flow reversals and the Proper Orthogonal Decomposition (POD) modes are investigated. We report the standard and cessation type flow reversals for the Jet-A fluid with Prandtl number of around 19. The power spectral density (PSD) of velocity follows the Bolgiano–Obukhov (BO) scaling.

The surface of a rectangular wing is morphed at high angles of attack such that it continues to operate at the reduced coefficient of lift (Cl) at which the baseline wing operates, but unlike the baseline wing, where the flow is separated, the flow remains attached on the morphed wing. A morphed surface is also generated to operate at a local design 2D (two-dimensional) Cl, which is obtained by incrementing the baseline Cl by a percentage at pre and post-stall angles of attack. The morphed surface is generated numerically using a novel ‘decambering’ technique, which accounts for the deviation of the coefficients of lift and pitching moment from that predicted by potential flow, analytically, using CFD and implemented experimentally by attaching an external Aluminium skin to the leading edge of the wing. Two different wing sections, NACA0012 and NACA4415, are tested on a rectangular planform. The effect of morphing on the aerodynamic performance is discussed, and aerodynamic characteristics are reported. Results indicate that significant improvement in aerodynamic performance is achieved at high angles of attack, especially at post-stall through this active morphed flow surface.

Experimental investigation on two rectangular wings with NACA0012 and NACA4415 profiles is performed at different Reynolds numbers to understand their aerodynamic behaviours at a high α regime. In-house developed numerical code VLM3D is validated using this experimental result in predicting the aerodynamic characteristics of a rectangular wing with cambered and symmetrical wing profile. The sectional coefficient of lift ((Formula presented.)) obtained from the numerical approach is used to study the variation in spanwise lift distribution. The lift and moment characteristics obtained from wind tunnel experiments are plotted, and change in the maximum coefficient of lift ((Formula presented.)) and stall angle (α stall) are studied for both of the wing sections. A significant addition to the novelty of the present experiments is to provide some comparison of the numerical induced drag coefficient, (Formula presented.) with experimentally fitted model coefficients using least square technique. A novel method is used to examine the aerodynamic hysteresis at high angles of attack. The area included in the lift- Re curve loop is a measure of aerodynamic efficiency, and its variation with angle of attack and wing plan forms is studied.