Adapting Krmn Vortex Street Phenomena to the Reduction of Induced Drag in Aerospace Vehicles - Thesis Proposal Example

?Client Sur August Karman Vortex Street Phenomena to the Application of Multiple Wingtip Feathered Winglets to the Reduction of Induced Drag In Aerospace Vehicles 1.0 Introduction One of the most important changes in our world is definitely the advancements in aerospace science and technology along with the research that goes with it. In fact, through this research, it has created a large industry of aviation which has changed the fabric of transportation and logistics all across the world. Naturally, the overall aerodynamic and flight stability characteristics of aeronautical vehicles remain the same as the same set of governing equations are applied in the designing and analysis phase. Thus, as a result, the overall shape and operating characteristics of aerospace vehicles resemble each other regardless of the company or the engineers which have worked on that particular design. Hence, similar characteristics pop up in each aerospace vehicle. However, there are many areas of improvement within these constrained design parameters. Especially as the fuel prices across the world soars every day, the concept of lowering the fuel cost for commercial, military as well as research vehicles have become more urgent then ever. Lowering the overall fuel consumption of the aerospace vehicle is an important criteria as it allows the aircraft to operate with less operational costs. Naturally, reduction in fuel consumption can be achieved by altering the various phases in the propulsion system of the aircraft; but the real reduction in fuel can only be achieved by lowering the aerodynamic drag that is acting upon the vehicle itself. This is quite important, as the overall drag force on the aerospace vehicle as well as the geometric zone of effect can have a large impact on the fuel consumption of the aero vehicle. In essence, drag force is nothing more then simple friction caused by the molecules of air (or any other fluid that the solid body may be immersed in) acting upon the surface of the aerospace vehicle. Naturally, this interaction can never become zero as long as viscous flow conditions are present in real life. However, the aerodynamic drag can be reduced by several important design techniques such as streamlining the body, changing the shape and the location of winglets as well as control surfaces on an aerospace craft. In essence, the main objective is to make the flow become as smooth as possible without causing any major obstructions in the flow around the body. This way, the amount of drag force induced by aerodynamic forces can be severely minimized and more importantly the coefficient of lift at the same time can be maximized. In addition, this will also help control the overall flight stability characteristics of the airplane itself. This thesis will concentrate on using Karman Vortex Street Phenomena to the Application of Multiple Wingtip Feathered Winglets for the Reduction of Induced Drag in Aerospace vehicles for the purpose of reducing fuel consumption. In fact, the interaction of aeroelastic forces with the aerodynamic coefficients is an interesting phenomena as well as its effects on the drag. Karman Vortex Street Phenomena will be examined in detail and then with corresponding experimental and computational data, the thesis will show that the application of multiple wingtip feathered winglets will have a measurable effect in reducing the overall drag and hence the overall fuel cost of the aerospace vehicle. The conclusion will be strengthened with data as well as original solution methods which will be introduced in this thesis. 2.0 Karman Vortex Street Phenomena One of the most interesting aerodynamic concepts is the Karman Vortex Street Phenomena. It is essentially a combination of swirling vortices formations that continue to repeat themselves. There are several reasons cited for the presence of Karman Vortex Street and one of the most important reasons is the unsteady and inhomogeneous separation of the flow around bluff solid bodies. As it can be seen in the below figure, Karman Vortex Street will be cyclic in its formation as well as continuation. Figure 1.1 : Karman Vortex Street Formation An interesting thing about Karman Vortex Street Formation is that it can be seen in civil structures as well. Especially in conditions of low speed turbulent flows, Karman Vortex Street phenomena can be seen in tall buildings as well. However, in aerospace structures, it is seen quite frequently as it can easily be found in any large solid body under certain conditions. Moreover, the presence of Karman Vortex Street will also affect aeroelastic interactions as well. Depending upon the vortex shedding frequency, various other aeroelastic vibrations can occur. In essence, the presence of vortex shedding causes drag to be formed which affects any solid bodies in the wake. Drag is necessarily to be avoided since drag component will disturb the overall stability of the plane or the aerospace vehicle; but also it will affect the fuel consumption of the craft itself. As the drag increases the amount of fluid fractioning upon the surface of the aircraft increases and as a result, the momentum of the aircraft is lessened and the momentum boundary layer on the aircraft surface is also changed. The kinetic energy of the craft is transformed in to potential energy and moreover the circulation around the plane is negatively effected in concordance with Jakowski Theorem. Naturally, as long as there is an obstruction on the flow field, vortex shedding is unavoidable. However, through design changes in the exterior part of the aircraft surface, it is possible to change the magnitude of the obstruction and thus it is possible to reduce the effects of drag caused by vortex shedding. Thus, it is the job of the aerospace engineer to incorporate the necessary design changes in order to minimize drag and to reduce the amount of vortex shedding for best results. The general equation for these types of vortex shedding is given by the equation given below for a cylinder. However, regardless of the shape, one of the most important terms that describe this phenomena is the Strouhal number which tries to give a relation between vibration frequency as well as the flow of the velocity as well as the dimensions of the solid body itself. Thus, a corresponding relationship is worked out for such aerospace bodies. 3.0 Vortex and Drag Effects In this literature survey, the main scope is to show the interrelation between vortex and drag effects along with its effect on fuel consumption. The work of Champigny shows that the two main irregularities in the flow field, namely a slight body sideslip and turbulence, are directly responsible for creating asymmetric vortices in the flow. However, interesting work that must be pointed out in Champigny’s work is that the asymmetric vortices in the flow are not temporal in nature. This means that stable and steady nature of vortices are not maintained in the flow. Figure 3.1 : Vortex flows on a body of revolution at high angles of attack Many aerodynamicists have related the regular unsteady two dimensional von Karman vortex street phenomena to the steady 3D vortex array, by using the principle of space- time equivalence. By this space time equivalence principle, flow development overall is related to time. It is measured from the beginning of an impulsive two-dimensional motion or as an alternative, from the instant a fluid particle makes contact with a three dimensional body. In the latter situation, time is defined as the distance that is travelled along the body and the axial component of free-stream velocity, as shown in section B-B in the above Figure. The steady asymmetric induced flow on the overall body results in the generation of steady side forces as well as yawing moments on the flow body. Since the asymmetric flow itself is steady, this causes the side force distribution along the length of the body to become sinusoidal and each maximum in this system corresponds to the detachment of a vortex sheet from the body. If the analysis is being conducted on a pointed, slender body, then the corresponding steady vortex asymmetry usually begins at the nose of the slender body and in addition the frequency at which the vortices are shed, increases with the angle of attack of the body. the angle of attack at which steady asymmetric vortices develop is dependent on the cone-half angle ?c. Asymmetric vortex development occurs when the angle of attack is approximately double the total included angle at the apex (?AV ? 2?A). However, on the other hand, slightly blunted bodies steady vortex asymmetry usually begins at the aft end of the body and with a further increase in angle of attack, the asymmetry becomes stronger and moves forward until it reaches the nose tip of the body at high angles of attack. Alternate vortex shedding does not occur as readily, and thus side force cells are much larger and can cover the entire cylindrical body. For slightly blunted bodies the onset angle of attack for steady asymmetric vortices is determined by the overall body fineness ratio. As it can be seen from the below Figure, there is a specific correlation between the angle of attack and the vortices induced. By changing the angle of attack, symmetric as well as asymmetric vortex formation can be induced and this can be instrumental in helping to reduce the overall drag component and thus the overall fuel consumption. Especially, asymmetric vortices can induce more drag on the aerospace vehicle and this can cause even more fuel consumption problems as the aerospace vehicle will need to use more fuel in order to overcome friction. Figure 3.1 : Vortex flows on a body of revolution at high angles of attack 4. Motivation The main aim of the thesis is to investigate the adaption of the Karman vortex street influence of multiple winglets using the application of multiple wingtip feathered winglets to the reduction of induced drag in aerospace vehicles. The study aims to optimize the CL3/2/CD parameter. The investigation will be performed both numerically and experimentally beginning with airfoils and winglet shape design Numerical simulations have the potential of greatly reducing design costs while providing a detailed understanding of the complex aerodynamics associated with each change. It can lead to accurate determination of aerodynamics; critical to the low-cost development of new advanced guided projectiles, and various other aerospace craft. Improved computer technology and state-of-the-art numerical procedures now enable solutions to complex, 3-D problems associated with complex aerodynamics. Using various techniques in computational fluid dynamics, it is possible to investigate the idea of using the application of multiple wingtip feathered winglets to the reduction of induced drag in aerospace vehicles. By utilizing suitable meshing techniques, the solid body can be geometrically designed and then meshed with the appropriate mathematical techniques. This way full Navier Stokes analysis can be performed on the body as well as the induced drag due to multiple winglets around the solid body. This thesis concentrates on around the use of multiple winglet vanes, as opposed to the current singular (upright) winglet/wingtip vane, to cause reduction in induced drag. By causing a reduction in induced drag, it would become possible to reduce the overall fuel consumption which seems to be a major concern for aerospace vehicles. With the high rising costs of fuel every day, it is essential to cut back on fuel without compromising on the efficiency of the aircraft or spacecraft. Naturally, due to constraints in design, the main shape of the aerospace craft can not be altered dramatically as the equations that govern flight require certain shapes and dimensions in order to obtain lift. However, by working with secondary surfaces such as winglet vanes, it is possible to reduce the overall drag formation on the aerospace vehicle without compromising on the lift capability of the aerospace craft. By controlling the amount of induced drag, some percentage of fuel consumption can be saved, which can be a competitive edge for that particular aerospace craft. The current trend is to lower fuel consumption in both commercial as well as in military crafts. Especially, in commercial crafts, this can become a competitive edge, which can cause a certain design to get selected. Hence, the main motivation of this thesis is to achieve this fuel consumption through the investigation of multiple winglet vanes and this thesis will try to demonstrate this by using computational work as well as experimental work. 5. Methodology In order to achieve the desired results in this thesis, a twofold approach will be used. The first method that is chosen would be to use Computational Fluid Dynamics or CFD to investigate the problem. Computational fluid dynamics (CFD) in essence is the numerical simulation of flow fields through the approximate solution of the governing partial differential equations for mass, momentum, and energy conservation coupled with the appropriate relations for thermodynamic and transport properties. This way it is possible to create a virtual lab and analyze the effect of flow on the solid surface of the craft and see the various effects of the variation. CFD has emerged as a critical technology for the aerodynamic design and assessment of aircraft and missiles. Improved computer technology and state of the art numerical procedures enable solutions to complex, 3-D problems associated with projectile and missile aerothermodynamics. Thus optimal performance parameters can be designed and tested using CFD analysis. Through the utilization of CFD, it becomes possible to test various parameters of the suggested design to see how multiple winglets will reduce the overall drag on the aerospace vehicle. Moreover, several different designs can be used and tested under virtual lab conditions so that the optimal configuration can be reached and proven under computational conditions. On the second phase of the investigation, some experimental studies will be carried out. This way, the numerical results of the computational investigation can be checked and verified with the experimental results. This way, the suggested design with the optimal parameters from your computational investigation would be tried out with experimental studies. Hence, the optimal configuration for the corresponding fuel reduction to achieved. 5.1 Geometric Modeling This is the most important part of the project as proper modeling of the body to be analyzed must be modeled in detail and according to scale in order to get reliable results. Here various modeling software can be used ranging from CATIA to using Geometric Modeler in ANSYS. Some basic models can be first done in GAMBIT to prepare for post processing in FLUENT. The model will incorporate the basic aerospace structure in its most primal skeletal form as well as the winglet configuration that is to be tested. The structure must be designed in such a way that the changes in the configuration can be done with minimum amount of work. Hence, universal vertices will need to be used, so that these changes can be done quickly. 5.2 Meshing Once the design is made, it is important to mesh the system in the most optimal manner to create a stable and converged solution during the post processing phase. All CFD solutions require appropriate grids with sufficient grid density in the regions of high flow gradients. Generally the region of high flow gradients is known in advance due to theoretical aerothermodynamics. The type of meshing that is chosen would be more denser near the winglets and less coarse toward the middle of the design shape. Hence, as boundary conditions are reached, the denser the mesh will have to be. The far flow field will also need sufficient triangular meshing to get the optimal result and the near field will need high level of meshing in order to be successful. However, the meshing must not be done too fine to create extremely long computational times. This is essential since the work will require several design shapes to be tested before it can be possible to create the optimal design. Once optimal design is reached, then a denser meshing scheme can be used for the post processing. Moreover, wrong meshing can also cause stability issues in post processing as well, so that also has to be taken into account for best results. In the meshing, the two-layer zonal model will be used for the near wall equations. The computational domain extends about 4D from the missile body. An outflow boundary condition will be used downstream, a pressure inflow (with freestream conditions) boundary condition will be used upstream, and a far-field pressure (nonreflecting) boundary condition will be used for the outer boundary. A no-slip wall boundary condition was used for all solid surfaces. Sample mesh distribution can be seen in the figure below. Figure 5.1: Sample Meshing on aMissile Geometry 5.3 Post Solver Stage The preliminary simulations will be done with a commercial flow solver FLUENT version 6.2. Fluent is a solver based on finite volume method. In control volume method, the computational domain is discretized into a finite set of control volumes or cells. The general conservation equations for mass, momentum and energy in integral form are discretized into a system of algebraic equations. All algebraic equations are then solved numerically to render the solution flow field. The implicit, compressible, unstructured-mesh scheme will be used to solve the three-dimensional, time-dependent, Reynolds-Averaged Navier-Stokes (RANS) equations. In the implicit scheme, each equation in the coupled set of governing equations is linearized implicitly with respect to all dependent variables in the set, resulting in a block system of equations. A block Gauss-Seidel, point implicit linear equation solver is used with an algebraic multi-grid method to solve the resultant block system of equations. The coupled set of governing equations is discretized in time and time marching proceeds until a steady state solution is reached. For the computation of high gradient flow fields, a hypersonic flow solver will be developed with FORTRAN programming language using an unconditionally stable McCormack technique. This technique is based on Predictor- Corrector formulation. In this, the governing partial differential equations are discretized with finite difference formulation. A second order accurate discretization will be done. This will be done to capture high pressure and temperature gradients in the flow field. A shock capturing approach will be used. 5.5 Turbulence Modeling Turbulence is the unsteady, irregular motion of fluid particle in which transported quantities fluctuate in time and space. It is one of the two unsolved mysteries of physics that existed at the start of twentieth century, the other being quantum physics. The latter has been resolved successfully but the former remains a challenge for physicists and mathematicians. There are various techniques for the numerical prediction of turbulent flows ranging viz. Reynolds Averaged Navier-Stokes (RANS), large eddy simulation (LES) and direct numerical simulation (DNS). DNS attempts to resolve all scales of turbulence from the largest to the smallest by solving the Navier-Stokes equations directly. LES attempts to model the smaller, more homogenous scales, while resolving the larger energy containing scales, thus making grid refinement for LES less than that for DNS. A derivative of LES is the detached-eddy simulation (DES), which is a hybrid approach combining the advantages of LES and RANS into one model. For the DES approach, RANS is used in the boundary layer, where it performs well and LES is then used in the separated regions where its ability to predict turbulence length scales is important. The grid requirement for DES is lesser than that of LES. For the RANS approach, the equations have been averaged over a time-scale, which is small in relation to the aerodynamic time-scale but large in comparison to the time-scale of the turbulent eddies. The RANS approach attempts to solve the time-averaged flow, which means that all scales of turbulence must be modeled. Turbulence models are semi-empirical formulations that are used to close the RANS equations by approximating the Reynolds stress terms. Reynolds stresses are modeled in two ways, namely eddy viscosity models and shear stress transport models. The Spalart-Allmaras turbulence model is a one-equation model based on the transport eddy viscosity and was designed for aerospace applications. It predicts flow separation very well. Other RANS based turbulence models are two equation models. The standard k-? turbulence model is eddy viscosity model for incompressible and compressible turbulent flows. It is a high Reynolds number model and is not meant to be used in the near wall regions where viscous effects are greater that the effects of turbulence. The standard k-? turbulence model has been implemented in FLUENT by means of wall functions. In the current work, two RANS based models viz. Spalart-Allmaras and standard k-? turbulence model will be implemented for preliminary computations. After confirming the suitability of one of the models, one of them will be implemented for final computations. The LES turbulence model was formulated for solving unsteady cyclic and vortical flows and should be chosen for modeling steady asymmetric vortex flow. The LES turbulence model will be implemented only if time permits. A major challenge in aerodynamic design is the accuracy of turbulence models for simulations of complex turbulent flows for example high angle of attack. Development of improved turbulence models has increased in the last decade due to the technological requirements of present aerodynamic systems, aided by advances in computers and numerical simulation capabilities. A variety of researchers have proposed methods for adapting algebraic turbulence models for high angles of attack. 5.6 Validation of Results It is a saying in scientific community, “everybody believes experimental results except the person who does the experiment and nobody believes the computational results except the person who produces the result”. So, before deriving any conclusion all the results produced from the CFD analysis will be validated against the experimental data for similar configurations and flow conditions, existing in the literature. Apart from experimental validation, a code validation studies will be carried out in the following areas of importance: Mesh size sensitivity Outlet boundary position Half symmetry Mesh Design Sensitivity All computational fluid dynamics (CFD) models require appropriate grids, with sufficient grid density in regions of high flow gradients. The problem lies in determining where these critical regions exist. In a vertical flow field, high flow gradient regions exist in the boundary layer, regions of shear layer separation and the primary and secondary vortices. A grid resolution study will be performed to minimize the error induced by the spatial resolution of the mesh. In order to validate the chosen mesh size, the number of grid points in the circumferential, radial and axial directions will be first halved and then doubled. The results obtained from these two simulations will then be compared to that of the original mesh. 5.6.2 Outlet boundary position In order to set the outlet boundary at atmospheric pressure, the outlet boundary has to be placed sufficiently far from the base of the body, so that it has no influence or a very weak influence on the upstream flow. 5.6.3 Half Symmetry For symmetric flow fields it is possible to model only a portion of the geometry to obtain a solution which is representative of the whole geometry. The advantage of modeling a portion of the geometry is that the number of grid points and thus the size of the overall mesh is reduced. In this way the computational time required to obtain a reasonably accurate solution is also reduced. An investigation is under taken to determine if the normal force and pitching moment coefficients obtained from a full model simulation and a half model simulation would be similar. This would justify the use of a half model for further symmetric flow field simulations. 6.0 Experimental Stage Once the final verified results are received through CFD efforts, then based upon the computational results, the optimum configuration can be made. It is possible to construct parts of the body and the winglet, so that testing in the wind tunnel can be done in order to see the results in real life. Coefficient of Lift and Coefficient of Drag will need to be calculated based upon the wind tunnel results and the main aim would be to reduce the induced drag. This way, through a further solution algorithm, the amount of fuel reduction in the operational cost can be calculated as well. Naturally, the CFD results, as well as the experimental results will need to coincide before it can be said that the results of this research are reliable. Statistical error analysis will also need to be completed, before a certain statement can be made at this stage. Once the experimental results have been verified through statistical techniques as well as through validation by computational techniques; then the necessary design recommendations can be made in detail. These design recommendations can then undergo the same procedures in this thesis, so that the computational and the experimental stages can be re-performed in order to verify the results. Thus, it is hoped at this stage that the necessary results will be obtained and the improvements of using a single winglet configuration can be proved with the appropriate recommendations for the reduction go drag and fuel in aerospace vehicles. 7.0 Conclusion Overall, this thesis will be a mixture of theoretical and experimental results, to create the optimal design parameters for reducing the induced drag through the utilization of the Karman Vortex Street phenomena. The current work presents a diverse array of passive flow control methods that have been implemented for the control of aerodynamic drag, side force and aerodynamic heating of slender bodies at low and high angles of attack for different flow regimes. Using single winglet panes will create a reduced drag and reduced fuel utilization as it will be demonstrated in this thesis. Both the computational part as well as the experimental modules should coincide in their results, so that appropriate recommendations can be made for future design prospects in aerospace vehicles. In time, this will hopefully become the standard norm for increasing fuel efficiency. Works Cited Anderson, J. Computational Fluid Dynamics. New York: Mc Graw Hill, 2002, Print Anderson, J. Compressible Flow. New York: Mc Graw Hill, 2002, Print Anderson, J. Fundamentals of Aerodynamics: New York: Mc Graw Hill, 2002, Print R.M. Cummings, J.R. Forsthye, S.A. Morton, and K.D. Squires. Computational Challenges in High Angle of Attack Flow Prediction. Progress in Aerospace Sciences, 39:369-384, 2003. D. Degani and L.B. Schiff. Numerical Simulation of the Effect of Spatial Disturbances on Vortex Asymmetry. AIAA Journal, 29(2):344-352, February 1991 J.E. Fidler. Active Control of Asymmetric Vortex Effects. J. of Aircraft, 18(4):267-272, April 1981. J.E. Fidler and M.C. Bateman. Asymmetric Vortex Effects on Missile Configurations. Journal of Spacecraft and Rockets, 12(11):674-681, November 1975. Fung C, Theory of Aeroelasticity. Mc Graw Hill, New York, 1978, Print T.T. Ng. Effect of a Single Strake on the Forebody Vortex Asymmetry. Journal of Aircraft, 27(9):844-846, September 1990. T.T. Ng and G.N. Malcolm. Forebody Vortex Control using Small Rotatable Strakes. Journal of Aircraft, 29(4):671-678, July-August 1992. D.M. Rao, C. Moskovitz, and D.G. Murri. Forebody Vortex Management for Yaw Control at High Angles of Attack. Journal of Aircraft, 24(4):248-254, April 1987. R. C. Mehta. Numerical heat transfer study over spiked-blunt body at Mach 6.80, AIAA paper 2000-0344, January 2000. Newman, Isadore and Carolyn R. Benz. Qualitative-Quantitative Research Methodology: Exploring The Interactive Continuum. Washington: SIU Press, 1998. Print Read More