Flow Simulation over a NACA 0017 Airfoil using Solidworks
Aim:- Flow Simulation over a NACA 0017 Airfoil using Solidworks
Objective:-
- Model the flow over a NACA0017 airfoil
- For the angle of attacks 0,2,4,6,8,10 compare the lift and drag forces
Introduction:-
An airfoil or aerofoil is the cross-sectional shape of a wing, blade (of a propeller, rotor, or turbine), or sail (as seen in cross-section).
An airfoil-shaped body moving through a fluid produces an aerodynamic force. The component of this force perpendicular to the direction of motion is called lift. The component parallel to the direction of motion is called drag. Subsonic flight airfoils have a characteristic shape with a rounded leading edge, followed by a sharp trailing edge, often with asymmetric curvature of upper and lower surfaces.
The lift on an airfoil is primarily the result of its angle of attack. When oriented at a suitable angle, the airfoil deflects the oncoming air (for fixed-wing aircraft, a downward force), resulting in a force on the airfoil in the direction opposite to the deflection. This force is known as aerodynamic force and can be resolved into two components: lift and drag. Most foil shapes require a positive angle of attack to generate lift, but cambered airfoils can generate lift at zero angle of attack. This "turning" of the air in the vicinity of the airfoil creates curved streamlines, resulting in lower pressure on one side and higher pressure on the other. This pressure difference is accompanied by a velocity difference, via Bernoulli's principle, so the resulting flow field about the airfoil has a higher average velocity on the upper surface than on the lower surface. The lift force can be related directly to the average top/bottom velocity difference without computing the pressure by using the concept of circulation and the Kutta–Joukowski theorem.
Airfoil Terminology
An airfoil has two edges namely leading edge and trailing edge, the straight line connecting the leading and trailing edge of the airfoil is called a chord line.
The line which is drawn equidistant from the upper and lower surfaces of the airfoil and connecting the leading and trailing edge is called a mean camber line.
The distance between the leading and trailing edge is called a chord of an airfoil.
The maximum distance between the mean camber line and the chord line is referred to as the maximum camber. This is one of the variables that determine the aerodynamic characteristics of the wings.
The maximum thickness to chord ratio is expressed in terms of percentage. For subsonic wings, the ratio is normally 12-14%.
The suction surface (a.k.a. upper surface) is generally associated with higher velocity and lower static pressure.
The pressure surface (a.k.a. lower surface) has a comparatively higher static pressure than the suction surface. The pressure gradient between these two surfaces contributes to the lift force generated for a given airfoil.
Depending on the camber (curvature) of the airfoil it is defined into two categories:
- Symmetrical airfoil
- Asymmetrical airfoil;
Symmetrical airfoil has identical upper and lower surfaces and the mean camber line is coincident with the camber line i.e. wing camber is reduced to zero.
Asymmetrical airfoil has wing camber and two different types of wings are used normally. A thick well-cambered wing will produce high lift at slow speeds whereas a thin long wing will produce good high-speed characteristics.
However, the symmetrical airfoil produces less lift than a non-symmetrical airfoil and also has relatively undesirable stall characteristics.
National Advisory Committee for Aeronautics (NACA) has developed numerous airfoils that are used in various aircraft around the world. They are usually of four or five digits. If the first two digits of the airfoil are zero then it is a symmetrical airfoil.
NACA 4 digit Airfoil Specification
This NACA airfoil series is controlled by 4 digits e.g. NACA 2412, which designate the camber, position of the maximum camber, and thickness. If an airfoil number is
NACA MPXX
e.g.
NACA 2412
then:
- M is the maximum camber divided by 100. In the example M=2 so the camber is 0.02 or 2% of the chord
- P is the position of the maximum camber divided by 10. In the example P=4 so the maximum camber is at 0.4 or 40% of the chord.
- XX is the thickness divided by 100. In the example XX=12 so the thickness is 0.12 or 12% of the chord.
NACA 4 digit Airfoil Calculation
The NACA airfoil section is created from a camber line and a thickness distribution plotted perpendicular to the camber line.
The equation for the camber line is split into sections either side of the point of maximum camber position (P). In order to calculate the position of the final airfoil envelope later the gradient of the camber line is also required. The equations are:
The thickness distribution is given by the equation:
- The constants a0 to a4 are for a 20% thick airfoil. The expression T/0.2 adjusts the constants to the required thickness.
- At the trailing edge (x=1) there is a finite thickness of 0.0021 chord width for a 20% airfoil. If a closed trailing edge is required the value of a4 can be adjusted.
- The value of yt is a half thickness and needs to be applied both sides of the camber line.
Using the equations above, for a given value of x it is possible to calculate the camber line position Yc, the gradient of the camber line, and the thickness. The position of the upper and lower surface can then be calculated perpendicular to the camber line.
The most obvious way to plot the airfoil is to iterate through equally spaced values of x calculating the upper and lower surface coordinates. While this works, the points are more widely spaced around the leading edge where the curvature is greatest and flat sections can be seen on the plots. To group the points at the ends of the airfoil sections a cosine spacing is used with uniform increments of β
Procedure:-
- Initially download the sample NACA file. Then create a new part file and insert the curve by clicking at insert command.
- 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) that is the angles 0,2,4,6,8,10 degrees with respect to the origin.
- After setting the angle of attack (AOA), the sketch of this curve is extruded by 0.5meters.
- Then by setting the Flow simulation wizard we set up an external flow and Air is selected as the fluid which flows across X-direction over the airfoil at 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) of at least 6-8 times of chord length.
- Now add Global Goals, Surface Goals, and Run the study.
- Generate Velocity and Pressure Cut Plots along the XY plane at 0.25meters (center of the wing).
- Generate Drag (surface goal in X-direction) and Lift (surface goal in Y-direction) plots and export in excel sheets for comparison.
Simulation Results:-
Study 1- Air velocity of 600 m/s and 0-degree angle of attack
Velocity Plot:-
The Velocity Cut-plot is the lowest at the front side of the airfoil and is maximum around the tail side. This is due to the boundary layer formation across the airfoil. As the angle of attack is zero (AOA = 0), velocity profiles will be identical for the lower and upper surface of the airfoil.
Maximum velocity = 652.91 m/s
Minimum velocity = 46.79 m/s
Pressure Plot:-
Pressure plot across the airfoil when it is in (AOA = 0) shows that the airfoil is symmetric and the pressure pattern is identical for the upper and the lower surface of the airfoil.
Maximum pressure = 445745.63 pascal
Minimum pressure = 56626.65 pascal
Drag Force Plot:-
Drag force is the force occurring on the airfoil in the X-direction and it is opposite to the flow of the air.
Average drag force = 3453.75 N
Lift Force Plot:-
Lift force is the force occurring on the airfoil in the Y-direction (vertically upwards) acting on the lower surface of the airfoil and it acts opposite to the force acting like the weight of the airplane.
Average lift force = -0.47 N
Study 2- Air velocity of 600 m/s and 2-degree angle of attack
Velocity Plot:-
The Velocity Cut-plot is the lowest at the front side of the airfoil and is maximum around the tail side. This is due to the boundary layer formation across the airfoil. As the angle of attack is zero (AOA = 0), velocity profiles will be identical for the lower and upper surface of the airfoil.
Maximum velocity = 717.20 m/s
Minimum velocity = 70.00 m/s
Pressure Plot:-
Pressure plot across the airfoil when it is in (AOA = 0) shows that the airfoil is symmetric and the pressure pattern is identical for the upper and the lower surface of the airfoil.
Maximum pressure = 451198.17 pascal
Minimum pressure = 31193.76 pascal
Drag Force Plot:-
Drag force is the force occurring on the airfoil in the X-direction and it is opposite to the flow of the air.
Average drag force = 3583.60 N
Lift Force Plot:-
Lift force is the force occurring on the airfoil in the Y-direction (vertically upwards) acting on the lower surface of the airfoil and it acts opposite to the force acting like the weight of the airplane.
Average lift force = 1596.00 N
Study 3- Air velocity of 600 m/s and 4-degree angle of attack
Velocity Plot:-
The Velocity Cut-plot is the lowest at the front side of the airfoil and is maximum around the tail side. This is due to the boundary layer formation across the airfoil. As the angle of attack is zero (AOA = 0), velocity profiles will be identical for the lower and upper surface of the airfoil.
Maximum velocity = 679.82 m/s
Minimum velocity = 51.30 m/s
Pressure Plot:-
Pressure plot across the airfoil when it is in (AOA = 0) shows that the airfoil is symmetric and the pressure pattern is identical for the upper and the lower surface of the airfoil.
Maximum pressure = 445656.94 pascal
Minimum pressure = 46280.94 pascal
Drag Force Plot:-
Drag force is the force occurring on the airfoil in the X-direction and it is opposite to the flow of the air.
Average drag force = 3667.27 N
Lift Force Plot:-
Lift force is the force occurring on the airfoil in the Y-direction (vertically upwards) acting on the lower surface of the airfoil and it acts opposite to the force acting like the weight of the airplane.
Average lift force = 3350.80 N
Study 4- Air velocity of 600 m/s and 6-degree angle of attack
Velocity Plot:-
The Velocity Cut-plot is the lowest at the front side of the airfoil and is maximum around the tail side. This is due to the boundary layer formation across the airfoil. As the angle of attack is zero (AOA = 0), velocity profiles will be identical for the lower and upper surface of the airfoil.
Maximum velocity = 689.70 m/s
Minimum velocity = 52.58 m/s
Pressure Plot:-
Pressure plot across the airfoil when it is in (AOA = 0) shows that the airfoil is symmetric and the pressure pattern is identical for the upper and the lower surface of the airfoil.
Maximum pressure = 442347.55 pascal
Minimum pressure = 41599.41 pascal
Drag Force Plot:-
Drag force is the force occurring on the airfoil in the X-direction and it is opposite to the flow of the air.
Average drag force = 3938.44 N
Lift Force Plot:-
Lift force is the force occurring on the airfoil in the Y-direction (vertically upwards) acting on the lower surface of the airfoil and it acts opposite to the force acting like the weight of the airplane.
Average lift force = 5038.45 N
Study 5- Air velocity of 600 m/s and 8-degree angle of attack
Velocity Plot:-
The Velocity Cut-plot is the lowest at the front side of the airfoil and is maximum around the tail side. This is due to the boundary layer formation across the airfoil. As the angle of attack is zero (AOA = 0), velocity profiles will be identical for the lower and upper surface of the airfoil.
Maximum velocity = 701.97 m/s
Minimum velocity = 47.54 m/s
Pressure Plot:-
Pressure plot across the airfoil when it is in (AOA = 0) shows that the airfoil is symmetric and the pressure pattern is identical for the upper and the lower surface of the airfoil.
Maximum pressure = 440311.20 pascal
Minimum pressure = 37332.53 pascal
Drag Force Plot:-
Drag force is the force occurring on the airfoil in the X-direction and it is opposite to the flow of the air.
Average drag force = 4317.88 N
Lift Force Plot:-
Lift force is the force occurring on the airfoil in the Y-direction (vertically upwards) acting on the lower surface of the airfoil and it acts opposite to the force acting like the weight of the airplane.
Average lift force = 6570.53 N
Study 6- Air velocity of 600 m/s and 10-degree angle of attack
Velocity Plot:-
The Velocity Cut-plot is the lowest at the front side of the airfoil and is maximum around the tail side. This is due to the boundary layer formation across the airfoil. As the angle of attack is zero (AOA = 0), velocity profiles will be identical for the lower and upper surface of the airfoil.
Maximum velocity = 710.58 m/s
Minimum velocity = 60.41 m/s
Pressure Plot:-
Pressure plot across the airfoil when it is in (AOA = 0) shows that the airfoil is symmetric and the pressure pattern is identical for the upper and the lower surface of the airfoil.
Maximum pressure = 445276.00 pascal
Minimum pressure = 33935.84 pascal
Drag Force Plot:-
Drag force is the force occurring on the airfoil in the X-direction and it is opposite to the flow of the air.
Average drag force = 4790.21 N
Lift Force Plot:-
Lift force is the force occurring on the airfoil in the Y-direction (vertically upwards) acting on the lower surface of the airfoil and it acts opposite to the force acting like the weight of the airplane.
Average lift force = 8340.72 N
Comparison of Data:-
It is quite visible that the average drag force and average lift force increases with the increase in the angle of attack. The graph below represents that the drag force is greater than the lift force up to a 4-degree angle of attack and till this condition the airfoil or the wing is not suitable for flying. After 4 degrees angle of attack, the lift force gradually increases and the drag force comparatively decreases which is the perfect condition for an airfoil to fly. But if we study more iterations, a stall condition occurs at which the wing becomes unsuitable for flying. The stall condition can be represented as a point where the drag force is greater than the lift force.
| S.no. | The Angle of Attack (degrees) | Average Drag Force (newton) | Average Lift Force (newton) |
|---|---|---|---|
| 1 | 0 | 3453.753 | -0.475 |
| 2 | 2 | 3583.604 | 1596.004 |
| 3 | 4 | 3667.272 | 3350.801 |
| 4 | 6 | 3938.447 | 5038.453 |
| 5 | 8 | 4317.885 | 6570.530 |
| 6 | 10 | 4790.210 | 8340.720 |
Conclusion:-
-
As the angle of attack (AOA) increases, the lift force & drag forces increase gradually. Thus, this continues to increase until the critical or stalling condition is achieved which is mostly around 15° - 20° for airfoils generally. Although the critical stalling angle depends on various factors such as aircraft weight, load factor, the centre of gravity of the aircraft, etc., the analysis should be done in consideration of these parameters.
- In this study, the flying condition of the wing or the airfoil is between 4-degree to 10-degree angle of attack because the average lift force is fairly greater than the drag force.
-
The magnitude of the drag generated by an object depends on the shape of the object and how it moves through the air. For thin airfoils, the drag is nearly constant at small angles (+/- 5 degrees). As the angle increases above 5 degrees, the drag quickly rises because of increased frontal area and increased boundary layer thickness. When the boundary layer separates, the wing is said to be stalled and both drag and lift become unsteady. Determining the drag is very difficult under stalled conditions.
- The angle of attack is directly proportional to the drag and lift force until the separation of the boundary layer because after that drag increases exponentially and the stall condition occurs.
Google drive link -
https://drive.google.com/drive/folders/1xz3TGmtmy9BIPc8YBO3WSDlOkeHgTJCt?usp=sharing
Reference:-
http://airfoiltools.com/search/airfoils
http://airfoiltools.com/airfoil/naca4digit
http://wwwaeroforum.blogspot.com/2015/03/air-foilterminology-air-foil-isprimary.html