Flow over an airfoil

Flow Over NACA0017 Aerofoil

Aim:- To study Airflow over the NACA0017 aerofoil for different Angle of Attack and observe respective Drag and Lift Force in corresponding cases.

Theory:- Aerofoil is the cross-sectional shape of a wing or a blade. 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 a symmetric curvature of upper and lower surfaces.

Angle of Attack (α)- In aerodynamics, angle of attack specifies the angle between the chord line of the wing of a fixed-wing aircraft and the vector representing the relative motion between the aircraft and the atmosphere.

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 thiickness is 0.12 or 12% of the chord.

So NACA0017 stands for an aerofoil with no camber and maximum thickness at 17% of the Chord Length. The aerofoil studied had a chord length of 1 metre, which means maximum thickness is 170 mm at 30% = 300 mm from front

Study Objectives: - Run Air Flow Simulation over suggested aerofoil with different Angle of Attack as mentioned below and observe the variation of Drag and Lift Force for each case

Sl.No	Angle of Attack (degree)
Case-1	0
Case-2	2
Case-3	4
Case-4	6
Case-5	8
Case-6	10

Boundary Conditions -

1. Inlet Velocity – Inlet Velocity for all six cases is set to 600 m/s.
2. Computational Domain (CD) – Since the effect of an aerofoil is dominant in the 2-dimensional plane, therefore computational domain is restricted to 2-D. Aerofoil is kept facing the negative X direction, and limits of CD is given below.

Axis	Maximum (m) (+ direction)	Minimum (m) (- direction)
X	13	5
Y	6	6
Z	0.2	0.2

Mesh Size –

1. It is set at “Automatic” and Level “3”
2. Additional Local Mesh of Level “3” is provided near the surface of the aerofoil, for better results.

Methodology:-

1. For different cases, I have created different Part files with same CFD configuration, that is, same Wind speed and Simulation Time of 2 seconds.
2. The results are obtained as graph of required quantity against the number of iterations.

Plots of different cases is shown below

Case 1 – α = 0 deg

Observation – Since the aerofoil is symmetrical, it doesn’t provide any lift at zero angle of attack, but do provide a significant drag force. This can be seen from the solver box, where even after many iterations, the solution for Lift Force (Force Y) doesn’t converge to any particular value.

Case 2 – α = 2 deg

Observation – A slight increase in the Angle of Attack, though negligible, produces a Lift Force. While Maximum, 6.5 kN Lift is generated, but it still less than the Max Drag Force, 14 kN, nearly half.

Case 3 – α = 4 deg

Observation – In this case, the Lift Force, max of 13.4kN reaches up to 89% of the Drag Force, max of 14.9kN

Case 4 – α = 6 deg

Observation – It is for this Angle of Attack 6 degree, that the NACA0017 Aerofoil’s Lift Force (Max of 20.36 kN) surpasses the Drag Force (Max 1.9 kN)

Case 5 – α = 8 deg

Observation – At this scenario drag (17.4 kN) drops to mere 65% of the Lift Force (26.6 kN)

Case 6 – α = 10 deg

Observation – In this case, Lift Force (33.7kN) is much higher than the Drag Force (19.2 kN)

Over all Summary

Sl.No	Angle of Attack (α) (degree)	Drag Force (X) (kN)		Lift Force (Y) (kN)
		Min	Max	Min	Max
Case-1	0	13.8	13.9	-0.15	0.1
Case-2	2	14.06	14.26	6.56	6.97
Case-3	4	14.73	14.92	13.07	13.47
Case-4	6	15.68	15.90	20.12	20.36
Case-5	8	17.21	17.41	26.67	26.89
Case-6	10	19.07	19.27	33.5	33.78

Conclusion –

1. It is clearly visible that the Symmetric Aerofoil like NACA0017 are ineffective at zero angle of attack (α, AOA) The Lift produced by the Wing increases as the AOA increases and correspondingly, the Drag decreases.
2. Also, the time taken by the Processor to reach the steady-state result for the Lift also improved as the AOA was increased.
3. The Stall Angle for most of the Aerofoil is round about 15-20 degree. So, no Stalling of the Aerofoil was observed.

Folder Link -Aishwary Mahajan - NACA0017 Aerofoil_GDrive

This folder contains

1. Six cad files with all settings for the CFD of this study.
2. A text file containing coordinates of the NACA0017-REFINED Aerofoil.
3. A sub-folder having excel files into which data from Solidworks is exported.
4. Result folders of all the six cases.

References:-

• Engineer's Edge - Fluid Property of Air
• NACA0017
• Wikipedia-Airfoil
• Wikipedia-Angle of Attack