IIT Madras is a public technical and research university located in Chennai, Tamil Nadu. Founded in 1959, it is recognised as an Institute of National Importance.

IIT Madras has been ranked as the top engineering institute in India for four years in a row by the National Institutional Ranking Framework of the MHRD

It currently offers undergraduate, postgraduate and research degrees across 16 disciplines in Engineering, Sciences, Humanities and Management. About 596 faculty belonging to science and engineering departments and centres of the Institute are engaged in teaching, research and industrial consultancy.

IIT Madras

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. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

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Department of Applied Mechanics

54th AIAA Aerospace Sciences Meeting

Prediction of post-stall aerodynamic characteristics of wing(S) with separated flow modeled as a single nascent vortex

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, Ð“ 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 Ã— MN equations are solved for Ð“ using LU decomposition. Â© 2014, American Institute of Aeronautics and Astronautics Inc. All rights reserved.

52nd Aerospace Sciences Meeting

A Numerical Study of the Unsteady Motion of a Wing using N-Body Approximation

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 flflying in echelon formation. Results like CL -α from the current method are compared with experiment. Additional results like section Cl distributions over wing spans, CM - α and CLW αtw (i.e. CL of the leading aircraft for diαerent angles of attack of the trailing aircraft) and analysis for (-)ve y-oαset, which are available only from the current numerical method are reported that supplement the experimental results. Detailed post-stall numerical analysis and the eαect of chord-wise, span-wise and vertical oαsets on the aerodynamics of the formation are also reported.

52nd AIAA  Aerospace Sciences Meeting -  AIAA Science and Technology Forum and Exposition, SciTech 2014

A numerical study of the aerodynamics of cessna 172 aircrafts in echelon formation

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. C<inf>L</inf>, C<inf>Di</inf>, 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. © 2014, American Institute of Aeronautics and Astronautics Inc. All rights reserved.

A numerical study of the unsteady motion of a wing using N-body approximation

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 C<inf>L</inf>- α from the current method are compared with experiment. Additional results like section C<inf>l</inf>distributions over wing spans, C<inf>M</inf>- α and C<inf>LW</inf>- α<inf>tw</inf>(i.e. C<inf>L</inf>of 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. © 2014, American Institute of Aeronautics and Astronautics Inc. All rights reserved.

A numerical study of the aerodynamics of Cessna 172 aircrafts in echelon formation

Low Reynolds number flows are important especially for MAV applications due to the presence of the LSB, which affects airfoil performance. The transition model of Fluent12 is used to extract the local velocity profiles and their second derivatives for flow past NACA0012 airfoil at Re=0.34X106. These form the input to a numerical linear stability code developed to understand the T-S wave characteristics in the LSB. The code solves the Orr-Sommerfeld equation locally for NACA0012 mean profiles at ##α=9° with parallel ow approach. The OSE is solved as a spatial stability problem where the temporal frequency is a real number and the eigen values are the spatial wave numbers. A complex Newton- Raphson technique is used for the Eigen value search and eN method is used to predict transition. The growth of the most unstable wave is such that the amplification ratio reaches a predefined number e9. Based on eN method the T-S wave frequency and phase speed are predicted. The LSB is observed using the Cp and velocity vector plots from Fluent12.

Linear stability analysis of laminar separation bubble over NACA0012 airfoil at low reynolds numbers

Numerical simulation of unsteady motion of wing through undisturbed uid 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, γ 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-1 numerical singularity is obtained from Rankine vortex model. No-normal flow boundary condition is imposed over the surface of mean camber and the resulting MN × MN equations are solved for γ using LU decomposition.

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.

51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition 2013

A numerical study of tandem pitching airfoils

An iterative decambering approach for the post stall prediction of wings using known section data as inputs is presented. The method can currently be used for incompressible flow and can be extended to compressible subsonic flow using Mach number correction schemes. A detailed discussion of past work on this topic is presented first. Next, an overview of the decambering approach is presented and is illustrated by applying the approach to the prediction of the two-dimensional Cl and Cm curves for an airfoil. The implementation of the approach for iterative decambering of wing sections is then discussed. A novel feature of the current effort is the use of a multidimensional Newton iteration for taking into consideration the coupling between the different sections of the wing. The approach lends itself to implementation in a variety of finite-wing analysis methods such as lifting-line theory, discrete-vortex Weissinger's method, and vortex lattice codes. Results are presented for a rectangular wing for α from 0 to 25 deg. The results are compared for both increasing and decreasing directions of α, and they show that a hysteresis loop can be predicted for post-stall angles of attack.

41st Aerospace Sciences Meeting and Exhibit

An Iterative Decambering Approach for Post-Stall Prediction of Wing Characteristics from Known Section Data

A novel scheme is presented for an iterative decambering approach to predict the post-stall characteristics of wings using known section data as inputs. The new scheme differs from earlier ones in the details of how the residual is computed. With this scheme, multiple solutions at high angles of attack are brought to light right during the computation of the residual for the Newton iteration. As with earlier schemes, multiple solutions are obtained for wings at high angles of attack and the resulting converged solution depends on the initial conditions used for the iteration. In general, the new scheme is found to be more robust at achieving convergence. Results are presented for a rectangular wing with two different airfoil lift curves and for a wing-tail configuration.

AIAA Paper

Post-Stall Prediction of Multiple-Lifting-Surface Configurations Using a Decambering Approach

A novel scheme is presented for an iterative decambering approach to predict the post-stall characteristics of wings using known section data as inputs. The new scheme differs from earlier ones in the details of how the residual in the Newton iteration is computed. With earlier schemes, multiple solutions are obtained for wings at high angles of attack as the final converged solution depends on the initial conditions used for the iteration. With this scheme, multiple solutions at high angles of attack are brought to light right during the computation of the residuals for the Newton iteration. In general, the new scheme is found to be more robust at achieving convergence. Experimental validation is provided using experimental airfoil lift curves from Naik and Ostowari for three different aspect ratios of rectangular wings. Results are presented from a study of the stall characteristics of wings of different planform shapes and two configurations of a wing-tail and a wing-canard configuration. Results are also presented from a study to investigate possible asymmetric lift distributions when the iterations were started with an initial asymmetric distribution of the decambering.

Fulltext

Poststall Prediction of Multiple-Lifting-Surface Configurations Using a Decambering Approach

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.

29th AIAA Applied Aerodynamics Conference 2011

Unsteady aerodynamics of multiple airfoils in configuration

An iterative decambering approach for the post stall prediction of wings using known section data as inputs is presented. The method can currently be used for incompressible flow and can be extended to compressible subsonic flow using Mach number correction schemes. A detailed discussion of past work on this topic is presented first. Next, an overview of the decambering approach is presented and is illustrated by applying the approach to the prediction of the two-dimensional Cl and Cm curves for an airfoil. The implementation of the approach for iterative decambering of wing sections is then discussed. A novel feature of the current effort is the use of a multidimensional Newton iteration for taking into consideration the coupling between the different sections of the wing. The approach lends itself to implementation in a variety of finite-wing analysis methods such as lifting-line theory, discrete-vortex Weissinger's method, and vortex lattice codes. Results are presented for a rectangular wing for α from 0 to 25 deg. The results are compared for both increasing and decreasing directions of α, and they show that a hysteresis loop can be predicted for post-stall angles of attack. © Ashok Gopalarathnam by Rinku Mukherjee. Published by the American Institute of Aeronautics and Astronautics, Inc.