Challenge

Flow over an airfoil

VVivek Singh16 Jun 2020

Numerical and Experimental Investigation by Flow over NACA0017 airfoil.

Aim:

To determine model flow over a NACA0017 airfoil. For the angle of attacks 0,2,4,6,8,10 compare the lift and drag forces.

Abstract:

Airfoil is the most basic unit of an aircraft wing, tail, or propeller blade the main function of the airfoil is the creation of a major part of the lift and drag of the aircraft as a whole. This airfoil generates a lifting force that acts at right angles to the airstream and a dragging force that acts in the same direction as the airstream.

There are many aero-foil profiles available on the NACA database, these are unique in characteristics. These characteristics are obtained by varying one or more parameters of the aerofoil shown in the below figure.

Airfoil Terminologies:

Leading-edge the front edge of an airfoil.
Trailing edge—the rearmost edge of an airfoil.
The chord line is the line joining the leading and trailing edge of the airfoil profile
Blade span—the length of the rotor blade from the center of rotation to the tip of the blade.
The Camber line is the curve formed by the locus of the points halfway between the upper and the lower surface.
Mean camber line—a line drawn halfway between the upper and lower surfaces of the airfoil. The chord line connects the ends of the mean camber line. Camber refers to the curvature of the airfoil and may be considered curvature of the mean camber line. The shape of the mean camber is important for determining the aerodynamic characteristics of an airfoil section. Maximum camber (displacement of the mean camber line from the chord line) and its location help to define the shape of the mean camber line. The location of the maximum camber and its displacement from the chord line is expressed as fractions or percentages of the basic chord length. By varying the point of maximum camber, the manufacturer can tailor an airfoil for a specific purpose. The profile thickness and thickness distribution are important properties of an airfoil section.
Induced flow—the downward flow of air through the rotor disk.
Resultant relative wind—relative wind modified by the induced flow.
The angle of attack (AOA)—the angle measured between the resultant relative wind and chord line.
The angle of incidence (AOI)—the angle between the chord line of a blade and the rotor hub. It is usually referred to as the blade pitch angle. For fixed airfoils, such as vertical fins or elevators, angle of incidence is the angle between the chord line of the airfoil and a selected reference plane of the helicopter.
Resultant relative wind—relative wind modified by the induced flow.
Center of pressure—the point along the chord line of an airfoil through which all aerodynamic forces are considered to act. Since pressures vary on the surface of an airfoil, an average location of pressure variation is needed. As the AOA changes, these pressures change, and the center of pressure moves along the chord line.

Total Aerodynamic force - It is the resultant force of the force components Lift and Drag.

Point of Impact - The point at which the air separates to flow about the airfoil. A high-pressure area or stagnation point is formed at the point of impact.

Theory:

The lift force of the airfoil depends on the angle of attack (AOA). Generally, Airfoil requires a positive angle of attack for lift (But a cambered airfoil can generate lift at zero angles of attack as well). Particularly subsonic airfoils are more round leading edge making the airfoil very insensitive to the angle of attack. Supersonic airfoils on the other hand have a sharper leading edge which is very sensitive to the angle of attack. So, in modern-day aviation, airfoils are selected on various parameters each of them optimized for their area of application.

Drag coefficient: Resistive force against the body caused by the body moving through the fluid. It is in the opposite direction to the velocity of the flow. The relative velocity between body and flow is called the drag coefficient. `Cd = frac{2Fd}{rho*A*v^2}` `

, where Fd = Drag force

ρ = mass density of fluid,

v = flow speed of the object relative to the fluid.

A = Reference Area

Lift Coefficient: Lift Coefficient is a dimensionless coefficient that relates the lift generated by a lifting body to the fluid density around the body, the fluid velocity, and an associated reference area A lifting body is a foil. `Cl = frac{2Fl}{rho*A*v^2}`
, where FL = Lifting force

ρ = mass density of fluid,

v = flow speed of the object relative to the fluid.

A = Reference Area

Objective: In this report, a detailed study is done on varying angles of attack (0,2,4,6,8,10) and its effect on lift forces (FL) and drag forces (Fd).

Simulation Setup:

Initially download the sample NACA file. Then create a new part file and insert the curve by clicking at insert command Curve Curve through XYZ points.
A dialog box opens browse the .txt file of the downloaded curve of NACA0017.
After inserting the curve, use the convert entity option to convert the curve into the sketch.
Use a line command at the trailing edge to make the sketch a close contour.
Now use the command of “Move entities” to rotate entities for varying the Angle of Attack (AOA). [So, given in this report the curve(spline) is rotated at angles 0,2,4,6,8,10] w.r.t origin.
After setting the angle of attack (AoA)the sketch of this curve is extruded by 0.5meter.
Then by setting the Flow simulation wizard we set up an external flow and the fluid that is selected is Air which flows across X-direction over the airfoil with 600 m/s.
Automatically a computational domain is created. Now set the computational domain in 2-dimensional in XY plane and adjust (Xmin, Xmax) such that it is 5-10 times of chord length and 10-15 times chord length respectively, and (Ymin, Ymax) with at least 6-8 times of chord length.
Now add Global Goals

Force in X and Y direction.
Frictional in X and Y direction.

Also, add Surface Goals by selecting the face of the airfoil

Surface Goals Force in X (drag force)
Surface Goals Force in Y (Lift force)

Run the simulation
After the solver comes to a stalling situation add a cut-plot velocity in XY direction to the airfoil.

Simulation Results:

Case – 1: When the angle of attack (AOA) of Airfoil is 0 degrees.

The Velocity Cut-plot is the lowest at the front side and is maximum at the tail side. This is due to the boundary layer formation across the airfoil. As the (AoA = 0) velocity profiles will be the same on both the sides of wing.

Pressure plot across the airfoil when it is in (AoA = 0) shows as the airfoil is symmetric and the pressure pattern in the above and below the airfoil is the same.

SG Force(X) – It is the drag force that is faced by the airfoil and the average value is 796.15N

SG Force(Y) – It is the lift force that is felt by the airfoil and the average value is -2.9N.

This means at 0 degrees the total Lift force is 0(zero) N. Because of the same pressure found at both the regions of the wing creating no lift force.

Case – 2: When the angle of attack (AOA) of Airfoil is 2 degrees.