Flow over an airfoil

FLOW OVER AIRFOIL

AIM: 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:

The NACA airfoils are airfoil shapes for aircraft wings developed by the National Advisory Committee of Aeronautics(NACA). The shape of the NACA airfoils is described using a series of digits following the word "NACA". The parameters in the numerical code can be entered into equations to precisely generate the cross-section of the airfoil and calculate its properties.

The NACA four-digit wing sections define the profile by:

1. First digit describing maximum camber as percentage of the chord
2. Second digit describing the distance of maximum camber from the airfoil leading edge in tenths of the chord.
3. Last two digits describing maximum thickness of the airfoil as percent of the chord.

For example, the NACA 2412 airfoil has a maximum camber of 2% located 40% (0.4 chords) from the leading edge with a maximum thickness of 12% of the chord.

The NACA 0015 airfoil is symmetrical, the 00 indicating that it has no camber. The 15 indicates that the airfoil has a 15% thickness to chord length ratio: it is 15% as thick as it is long.

angle of attack is the angle between the sail's chord line and the direction of the relative airflow.

PROCEDURE:

STEP 1: Open a new document .Click on insert -> Curves -> Curves through XYZ Planes.Then browse for NACA Sample and click OK.There will be a curve.View the curve in front plane and click on sketch.Select the curve and Click on convert entities to convert the curve into sketch.Close the loop at the other end of curve by line tool.Click on move entities ->rotate entities and select line and spline.Select the rotation point.And give angle of attack(in negative w.r.t origin).Extrude it to 0.200m and save the file.

STEP 2: After saving the file go to flow simulation wizard.give the project name .For Analysis Type click External ->Time Dependent ->Total Analysis time-5sec ,Output time step-1sec.Select Air as Fluid.Change the Velocity in X-direction as 600 m/s.Click OK.

STEP 3: Click on Computational Domain ->Edit Definations ->2D Simulation.Again click on Computational Domain ->Edit Definations ->Adjust the airfoil i.e 5 times the chord length at the left side of airfoil and 10-15 time the chord length at the right side of air foil.

Air foil after the conditions given in solver: Airfoil when given 2D Simulation:

After editting Computational Domain:

STEP 4: Click on Goals ->Insert Global Goals ->Select Force(X),Force(Y),Friction Force(X) and Friction Force(Y).Again click on Goals -> Insert Surface Goals ->Select top surface of foil and select Force(X) and Force(Y).Click on Mesh ->Show Basic Mesh ->Insert Local Mesh ->Select top surface and give 3 in Refine Cells.Run the Simulation.

STEP 5: Stop the Simulation solver and save the results.come back to fileand click on Results->Cut plot ->Contours -> Mesh(if needed) ->velocity(in countours) ->click ok.There you will get velocity profile of air foil.For Pressure profile slect pressure in countours.

STEP 6: Now click on Goals ->Insert Goals ->Select SG Force(X) ans SG Force(Y) ->Show data ->OK.

RESULTS:

SG Force(X)1 is Surface Goal Force at X-Axis called as Drag Force.

SG Force(Y)1 is Surface Goal Force at Y-Axis called as Lift Force.

Case 1:Angle of Attack=0 Degree.

Velocity profile:

Pressure profile:

Goal Plots:

Drag Force and Lift Force:

Case 2:Angle of Attack=2 Degree

Velocity Profile:

Pressure Profile:

Goal Plots:

Drag Force and Lift Force:

Case 3:Angle of Attack=4 Degree

Velocity Profile:

Pressure Profile:

Goal Plots:

Drag Force and Lift Force:

Case 4:Angle of Attack=6 Degree

Velocity Profile:

Pressure Profile:

Goal Plots:

Drag Force and Lift Force:

Case 5:Angle of Attack=8 Degree

Velocity Profile:

Pressure profile:

Goal Plots:

Drag Force and Lift Force:

Case 6:Angle of Attack=10 Degree

Velocity profile:

Pressure Profile:

Goal plot:

Drag Force and Lift Force:

Comaprision of Drag Force

Comparision of Lift Force

Lift Force v/s Drag Force

CONCLUSION:

• Generally increase in Angle of Attack results in increase of both Drag Force And Lift Force.
• As per observation Drag Force is Inversly proportion to Angle of Attack and Lift Force is Directly Proportion to Angle of Attack.
• Symmetry in velocity and pressure profile is seen in Case 1 i.e AOA=0.