Challenge

Flow over an airfoil

SS Manoj Kumar25 Jun 2020

Flow over an airfoil Using Solid Works

Abstract:

In this Project, CFD analysis of the two-dimensional subsonic ow over a NACA 0017 airfoil at various angles of attack and compare the lift and drag force. The geometry of the airfoil is created using SOLIDWORKS. CFD analysis is carried out using Flow Simulation Wizard at various angles of attack from 0° to 10° with an interval of 2°. Velocity distributions over NACA-0017 airfoil for different angles of attack are presented. Drag Force and Lift Force are plotted and compared.

1. Introduction:

Airfoil is de ned as the cross-section of a body that is placed in an airstream in order to generate useful aerodynamic force. Computational Fluid Dynamics (CFD) has fascinated researchers, as it is a fast, accurate and reliable method of analyzing the variation of hydrodynamic properties of the ow over and within a body. Computational Fluid Dynamics (CFD) is a branch of fluid mechanics that uses numerical methods and algorithms
to solve complex problems involving fluid flow.

Angle of attack is the angle between the oncoming air or relative wind and a reference line on the airplane or wing. Sometimes, the reference line is a line connecting the leading edge and trailing edge at some average point on the wing.

Effect of angle of attack:
Increases as the angle of attack increase the lift force increase proportionately. And a certain point, called critical angle of attack, upper surface ow is more separated and the airfoil or wing is producing its maximum coecient of lift.
The study of airfoil behavior to varying angle of attacks is made in order to determine the best possible dimensions for practical conditions. In most airplanes, this is about 16 to 18 degrees. During a cruise, the requirement is to maintain the angle of attack where the total drag on the airplane is minimum so as to use the minimum thrust and therefore, minimum fuel consumption. This could occur at an angle of about 3 to 5 degrees.

The present study aims to analyze the ow eld for NACA-0017 airfoil at various angles of attack and also to compare the lift and drag force.

2. Geometry and Mesh Generation:

3. Inputs and Boundary Conditions:

The problem consists of ow around an airfoil at various angles of attack (0, 2, 4, 6, 8, 10 degrees). The inputs and boundary conditions are presented in Table 1.

Table 1. Inputs and boundary conditions

4. Procedure:

i. An airfoil is created by using convert entities then line command.

ii. For the different angle of attacks rotate entities is used.

iii. The airfoil is extruded to 0.5 mm. Now save the file.

iv. The flow simulation is created using the wizard menu.

v. The analysis type is selected as external. Time-dependent is selected.

vi. Total analysis time is selected as 5 sec and output time is 1sec.

vii. The uid is selected as air and velocity parameter over X-direction is selected as 600m/s and then click finish.

viii. Now a computational domain is created by using the 2D option.

ix. The computational domain is expanded along both y-directions and x-direction in such a way that the length after the wake region should be 5 times the
length ahead of wake region.

x. The Global mesh is de ned and the local mesh is de ned by selecting the surface. It is re ned up to a level of 3.

xi. The goals option is selected then force x and force y selected after selecting the surface.

xii. Run the simulation.

xiii. Now a cut plot is made along x-direction and velocity over the airfoil is selected.

xiv. Color re ning is given 200.

xv. A goal plot is plotted between drag force along the x-direction and lift along the y-axis.

xvi. The above procedure is followed for different attack angles and the following results are used to plot the necessary graphs.

5. Results:

Results are computed and studied on different angle of attacks.

let's run one by one on the different angle of attacks.

Here,

Integrated force along the x-axis is Drag force.

Integrated force along the y-axis is Lift force.

Case 1: Angle of attack at 0°

Results calculated.