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Computational Study of Ailerons in Cross Flows Ground Effects

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Computational Study of Ailerons in Cross Flows Ground Effects cs & Aero ti sp au a n c o e r E e Alsarheed and Sedaghat, J Aeronaut Aerospace Eng 2016, 5:1 n A g f i o n Journal of Aeronautics & Aerospace DOI: 10.4172/2168-9792.1000161 l e a e r n i r n u g o J Engineering ISSN: 2168-9792 Research Article Open Access Computational Study of Ailerons in Cross Flows Ground Effects and Biplanes Configurations Alsarheed M* and Sedaghat A Department of Mechanical Engineering, Australian College of Kuwait Mishref, Kuwait Abstract An aileron, part of the trailing edge of a fixed-wing airplane, is used to control aircraft’s movement around its longitudinal axis (roll). Ailerons have significant impacts on airfoil surface pressure and its lift and drag coefficients. Both panel and finite volume methods were used on a NACA 2412 airfoil with a 20% aileron in a cross flow. The aerodynamic performance of ailerons alone, in a biplane configuration, and in the ground effects has been computationally investigated using both the panel and the finite volume methods. Several parameters were analyzed including the effects of the attack angle of the airfoil, aileron deflection angle, the ride height from the ground, and the characteristics of biplanes. Results of both computational methods are presented and discussed for the aforementioned configurations of NACA 2412 airfoil with aileron. Keywords: Aileron; Biplane; Finite volume; Ground effect; NACA Airfoils Characteristics 2412; Panel method For every action, there is an equal and opposite reaction. According Introduction to Newton's third law [10,11] an airfoil generates lift by diverting the motion of flow over its surface in a downward direction, resulting in an From flying a kite to operating a hypersonic vehicle, achieving and equal upward reaction. Typical airfoil geometry is shown in Figure 1. maintaining a high lift is always critical to a successful performance of a flying craft. Some of the advantages of an efficient lift system include The main parameters of an airfoil, which play a key role in its improved maneuverability, higher wing payload capacity, longer range aerodynamic performance, are the angle of attack, the chord length, for given gross weight, and lower takeoff and landing speeds. Wing and the mean camber line. The chord is the straight line across the components such as trailing edge flaps play a major role in altering airfoil and it is used to measure the airfoil length. The mean camber is the lift in conventional fixed-wing aircrafts. They control lift and extend line halved the airfoil thickness and it is used to measure airfoil curvature. speed range by upward or downward movements causing an increase Common airfoil shapes have been characterized by the National or decrease in lift. Also, the flap deflection can reduce efficiency by Advisory Committee for Aeronautics (NACA) [12]. Each airfoil shape causing or increasing drag, which sometimes is difficult to reduce due is defined by this system by a series of digits that correspond to non- to flow with high Reynolds number. Part of the trailing edge assembly dimensional airfoil properties. The number of digits series varies, but in is a flap called ailerons, which is a French word for ‘little wing’ that is this paper the focus is on four digit airfoils (NACA 2412). used to control the aircraft in its longitudinal axis (roll). This movement A NACA 2412 airfoil has a maximum camber that is 2% of the is called ‘rolling’ or ‘banking’ and it alters the flight path due to the chord length with a maximum thickness of 12% of the chord length, titling of the lift vector. One of the main components that impacts lift located 4/10 of the chord length away from the leading edge. is the airfoil. An airfoil is a cross sectional part designed to generate lift when it is subject to an air flow. Ailerons are not just limited to Computational Modeling aircraft applications. They are also used in high speed levitation trains to control and maintain the hovering in track. A NACA 2412 airfoil configuration with a 20% aileron was used for the panel and finite volume methods, and it is shown in Figure 2a. The The airfoil configuration has been widely studied. There are several chord length of the airfoil is designated by c, the velocity of the uniform studies that optimize the airfoil configuration and assess the maximum flow is denoted by U, δf is the aileron deflection angle, and α represents airfoil lift capabilities [1,2]. Other literatures study the efficiency of the angle of attack. The flow is assumed to be two-dimensional steady the airfoil and the effects of different parameters on it [3]. The camber and incompressible Laminar flow. and the Reynolds number effects have also been studied by many including Levy [4]. Other work done by Birch [5] and Rinoie [6] include calculations of unsteady loading on the airfoil and the effect *Corresponding author: Alsarheed M, Department of Mechanical Engineering, of trailing edge flaps on flow vortex shapes. Also, the performance of a Australian College of Kuwait Mishref, Kuwait, Tel: 965 1828225; E-mail: m.alsarheed@ wing in ground effect and the flow field characteristics has been studied ack.edu.kw in [7,8]. In addition, the lift distribution between biplane wings and Received February 02, 2016; Accepted February 26, 2016; Published February other biplane characteristics have been looked at in [9]. The objective 29, 2016 of this paper is to investigate the performance and the ground effects Citation: Alsarheed M, Sedaghat A (2016) Computational Study of Ailerons in on the ailerons of a biplane configuration. Both panel and finite volume Cross Flows Ground Effects and Biplanes Configurations. J Aeronaut Aerospace Eng 5: 161. doi:10.4172/2168-9792.1000161 methods were used on a NACA 2412 airfoil with a 20% aileron in a cross flow. Numerical measurements of surface pressure distributions Copyright: © 2016 Alsarheed M, et al. This is an open-access article distributed under the terms of the Creative Commons Attribution License, which permits were also obtained to determine the lift and drag coefficients for various unrestricted use, distribution, and reproduction in any medium, provided the configurations. original author and source are credited. J Aeronaut Aerospace Eng ISSN: 2168-9792 JAAE, an open access journal Volume 5 • Issue 1 • 1000161 Citation: Alsarheed M, Sedaghat A (2016) Computational Study of Ailerons in Cross Flows Ground Effects and Biplanes Configurations. J Aeronaut Aerospace Eng 5: 161. doi:10.4172/2168-9792.1000161 Page 2 of 11 Panel method all the points on the airfoil will be constant and the value is denoted by C. Also, for all points r on the airfoil surface, equation (2) becomes: The airfoil was analyzed using a vortex panel method, where the 1 airfoil was approximated by a finite number of panels. As shown in u y− v x − ∫ γ ln r− r ds−= C 0 ∞∞ 0 ( 0 ) 0 (3) Figure 2b, the airfoil surface is divided into 70 panels. There are more 2π panels near the leading and trailing edge to accommodate for the rapid The Kutta condition is an aerodynamics principle and it is changes in flow near these two ends. Each panel had local, uniform and applicable to the trailing edges of the airfoils. This condition states that distributed vortices. Since the velocities can be singular at the center of both the pressure above and below the trailing edge of the airfoil must the vortex, and to avoid singularities, these vortices had infinitesimal be equal. It also requires that the flow be smooth and must leave the strength γ0ds0, where γ0 is the vortex strength per unit length, and ds0 trailing edge in the same direction at both upper and lower edges and represents the length of a small segment of the airfoil. it is expressed as follows: The governing equation for the stream function due to all such γγupper= − lower (4) infinitesimal vortices at a point in space may be given by: For a finite number of N panels, there are equal N unknown vorticity γ 0 ψ = − strengths, denoted by γo, j, as shown in Figure 2(b). To approximate ∫ ln (r r00) ds (1) 2π the line integral over the entire airfoil, it can be integrated as several line integrals over N panels, taking each value of γo as constant and so r0 where is a point in the x-y plane with coordinates x0 and y0 and equation (3) becomes: the distance between the position of the vortex, and the point where the N − velocity is evaluated with coordinates x and y is denoted by u∞ y v∞.x The x-y (5) uy∞∞i− vx i −∑ A ij,0γ j −= C 0 plane is originated at the leading edge of airfoil with x-axis along chord j=1 length and y perpendicular to it. 1 = − Aij, ∫ ln ( ri r00) ds Equation (1) is integrated over all vortices on the airfoil surface 2π and the stream function associated with the free stream is given by Equation (5) can also be solved over the entire airfoil as several line u∞ y − v∞ x , where u∞ and v∞ are the x- and y- components of integrals over the panels for the unknown vorticity strengths. the free stream velocity. Taken these effects into account, the stream function at any point r in space is given by Finite volume method 1 The numerical simulations were conducted using FEMLAB ψγ=u∞∞ y −− v x 00ln ( r − r) dso (2) 6 2π ∫ software with a Reynolds number Re = 3 × 10 . The Reynolds number Re is defined as Since the body itself is a streamline, the stream function value at Re = Uc υ (6) Where U is the velocity of a uniform flow, c is the chord length of the airfoil, and υ denotes the kinematic viscosity The governing non-dimensional equations for the fluid flow are described by the following continuity and Navier-Stokes equations: ∇=.0u ∂u (7) ρ −∇υ.
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