Flow over an airfoil

Aim: To Perform Transient Flow analysis over NACA 0017 airfoil using Solidworks Flow Simulation.

Theory: Airfoil geometry can be characterized by the coordinates of the upper and lower surface. It is often summarized by a few parameters such as: maximum thickness, maximum camber, position of max thickness, position of max camber, and nose radius. One can generate a reasonable airfoil section given these parameters. This was done by Eastman Jacobs in the early 1930's to create a family of airfoils known as the NACA Sections.

• Mean Chamber Line: Set of points halfway between upper and lower surfaces
  • Measured perpendicular to mean chamber line itself
• Leading Edge: Most forward point of mean chamber line
• Trailing Edge: Most reward point of mean chamber line
• Chord Line: Straight line connecting the leading and trailing edges
• Chord, c: Distance along the chord line from leading to trailing edge
• Chamber: Maximum distance between mean chamber line and chord line
  • Measured perpendicular to chord line

NACA (National Advisory Committee for Aeronautics) or NASA (National Aeronautics and Space administration) Identified different airfoil shapes with a logical numbering system. Before NACA series, airfoil design was rather arbitrary with nothing to guide designers except experience with known shapes and experimentation with modified shapes.

• NACA 4-digit series
• NACA 5-digit series
• NACA 1-series or 16-series
• NACA 6- series
• NACA 7- series
• NACA 8- series

NACA Four Digit Series

• First digit specifies maximum camber in percentage of chord
• Second digit indicates position of maximum camber in tenths of chord
• Last two digits provide maximum thickness of airfoil in percentage of chord

Reason for using Transient Simulation instead of Steady State Simulation

The reason for using Transient Simulation is that Lift Coefficient is not necessarily a transient changing quantity, e.g. If we analyse the Laminar Flow over an airfoil under 0-degree angle of attack than the result will not be changeable with respect to time.

But for Turbulent Flow and higher angle of attack the flow around airfoil is subjected to flow separation on the upper side of the airfoil and separating flow will generate transient vortices, which are convected by the main airstream to downstream. This is causing the pressure field around the airfoil to transiently change with the movement of vortices and consequently lift coefficient is becoming a changeable quantity over time.

Procedure:

1. In this we will first download NACA airfoil coordinates from airfoil tool which will be downloaded as csv file and then we will open notepad and then we will copy the text of X, Y and Z coordinates and the save it.
2. Now we will open Solidworks and then go for part modelling and then in feature we will go to curve and then we will browse the file and click on open and then ok to open the file.
3. Now our airfoil is produced, We will go to Front Plane and then we will click on convert entities to convert it into sketch.

Note: If we have open airfoil then we have to add a line at the end of the airfoil and if it is a closed airfoil then there is no need to add anything.

4. Now since at first we are designing for 0-degree angle of attack we are not going to rotate the airfoil sketch.
5. Now we will go to features and then Extrude boss/base and then we will extrude it to 2mm using the mid plane and then we will save our file.
6. Now we will go to Solidworks Flow Simulation and then we will go to wizard and then we will define it as external flow problem, switching on the time dependency the fluid will be air and X- velocity to be 600 m/s.
7. We will get an automatically generated computation domain and then we will edit the domain and we will keep a distance of 10-15 times from the leading edge and 20-25 times from the trailing edge to get a better plot of the contours.
8. We will now go to mesh and then insert local mesh and then select the top face of the airfoil and then refinement to 3.
9. Now we will add the goals and then we will go to Global goals and then we will add X Force, Y Force, X Frictional Force and Y Frictional Force and then we will go to Local goals and then we will add X Force and Y Force.
10. Now we will run the simulation for our results. Now after getting the results we will plot Pressure and Velocity Contour.
11. We will repeat the steps just the changes we will do that after converting the sketch we will rotate it in which the entities to rotate would be the sketch and then rotation about would be the front axis and degree would be -2 degree and simultaneously for all the cases.

Computation Domain

Local Mesh

Results and Plots:

0-Degree Angle of Attack

Model

Velocity Contour

Pressure Contour

Global Goal Force

Global Goal Friction Force

Surface Goal Force

Goal Report

2-Degree Angle of Attack

Model

Velocity Contour

Pressure Contour

Global Goal Force

Global Goal Friction Force

Surface Goal Force

Goal Report

4-Degree Angle of Attack

Model

Velocity Contour

Pressure Contour

Global Goal Force

Global Goal Friction Force

Surface Goal Force

Goal Report

6-Degree Angle of Attack

Model

Velocity Contour

Pressure Contour

Global Goal Force

Global Goal Friction Force

Surface Goal Force

Goal Report

8-Degree Angle of Attack

Model

Velocity Contour

Pressure Contour

Global Goal Force

Global Goal Friction Force

Surface Goal Force

Goal Report

10-Degree Angle of Attack

Model

Velocity Contour

Pressure Contour

Global Goal Force

Global Goal Friction Force

Surface Goal Force

Goal Report

Drag Force

Lift Force

Drag Vs. Lift

Observation:

• It has been observed that with increase in angle of attack drag force increases and hence vice versa.
• It has been also observed that with increase in angle of attack lift force also increases and hence vice versa.
• We can also find drag and lift coefficient of different airfoil and can choose the airfoil which suits in its application.
• If we do Transient analysis animation of different quantities, we will get behaviour changes with respect to time.