CFD Analysis of Flow over NACA 0017 Airfoil

OBJECTIVE :

• To perform flow simulation over a NACA 0017 airfoil.
• To compare lift and drag forces at different values of angle of attack.

ABSTRACT :

The purpose of this project is to perform flow analysis over a NACA 0017 airfoil having air as a fluid. The flow simulation is performed in total six cases for different angles of attack, 0, 2, 4, 6, 8, 10. Velocity and pressure contours are obtained using cut plots for each cases. Subsequently, the lift and drag forces are calculated for each case. Further, for all the angles of attack, lift and drag forces are compared and a conclusion is made from the analysis of the result.

SOLVER USED : Solidworks Flow Simulation

INTRODUCTION :

Computational Fluid Dynamics (CFD) :

Computational fluid dynamics (CFD) is a branch of fluid mechanics that uses numerical analysis and data structures to analyse and solve problems that involve fluid flows. Computers are used to perform the calculations required to simulate the free-stream flow of the fluid, and the interaction of the fluid (liquids and gases) with surfaces defined by boundary conditions.

CFD is applied to a wide range of research and engineering problems in many fields of study and industries, including aerodynamics and aerospace analysis, weather simulation, natural science and environmental engineering, industrial system design and analysis, biological engineering, fluid flows and heat transfer, and engine and combustion analysis.

NACA 0017 Airfoil :

An airfoil (American English) or aerofoil (British English) 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 NACA airfoils are developed by the America's National Advisory Committee for Aeronautics (NACA). The shape of the NACA airfoils is described by 4 digits followig "NACA" e.g. NACA 0017, which designate the camber, position of the maximum camber and thickness. If an airfoil number is NACA MPXX 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.

All NACA airfoils can be accessed here : AirFoil Data Base

Here, in this project, NACA 0017 Airfoil is used whose geometry looks as follows :

Angle of Attack :

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. Angle of attack is the angle between the body's reference line and the oncoming flow.

Lift Force :

A fluid flowing around the surface of an object exerts a force on it. Lift is the component of this force that is perpendicular to the oncoming flow direction. Lift conventionally acts in an upward direction in order to counter the force of gravity, but it can act in any direction at right angles to the flow.

Drag Force :

A drag force is the resistance force caused by the motion of a body through a fluid. It acts opposite to the direction of the oncoming flow velocity.

MODEL :

The Curve is inserted in solidworks using its XYZ coordinates and then extruded.

CASES :

Cases :	Case 1	Case 2	Case 3	Case 4	Case 5	Case 6
Angle Of Attack :	0 °	2 °	4 °	6 °	8 °	10 °

FLOW SIMULATION :

Setup :

Project name: angle_of_attack_0

Analysis type: External; Time-Dependent; Simulation Time (5s); Time Step (1s)

Type of fluid: Air

Type of flow: Laminar & Turbulent

Wall Condition: Adiabatic Wall with 0 micro-meter roughness

Initial Condition: Pressure (101325 Pa), Temperature (293.2K), Velocity x-direction (600 m/s)

Gravity : -9.81 in Y direction

Computational Domain Type :2D Simulation

Global Mesh :

Surface Mesh :

Global Goals : Force(x), Force(y), Frictional Force(x) &Frictional Force(y).

Surface Goals : Force(x) refers to the drag force and Force(y) refers to the lift force.

For each cases, the calculation is run and stopped when solutions seems to converge.

RESULT :

In order to generate a contour, a cut plot was provided parallel to the cross-section of the airfoil and at the mid length.

Note : Integrated Force in X direction represents drag force while integrated force in Y direction represents lift force.

Case 1 :

Case 2 :

CASE 3 :

CASE 4 :

CASE 5 :

CASE 6 :

OBSERVATION :

The following table shows the useful results obtained in different cases.

Angle of attack	Average Drag Force (N)	Average Lift Force (N)	Max. Velocity (m/s)	Min. Velocity (m/s)	Max. Pressure (Pa)	Min. Pressure (Pa)
0°	1490.69	-310.40	679.69	0	460859.71	53185.70
2°	1503.43	910.02	682.12	0	456734.07	49889.24
4°	1592.20	2098.04	694.64	0	492032.85	44790.34
6°	1759.15	3321.24	697.39	0	496847.09	40470.14
8°	2015.19	4502.41	712.72	0	436829.74	34854.02
10°	2337.08	5774.18	717.00	0	422058.58	32513.95

1. On increasing the angle of attack (AOA) from 0⁰ to 10⁰, the drag & lift force increases.
2. Since, up to 10⁰ of AOA, the lift force is increasing. Therefore, the stall angle has not achieved yet.
3. In first case, the lift forces are negative as the lift due to pressure difference above and below the airfoil is insufficient to overcome the gravity force.
4. As the AOA increases, the maximum velocity of air over the airfoil increases with increase in AOA.
5. It is observed that at 0⁰ AOA, the pressure distribution of air over the upper & lower surface of the airfoil is almost symmetrical but as the AOA increases, the air pressure over the upper surface increases and the air pressure decreases over the lower surface of the airfoil.
6. Similar is the case for velocity distribution.

CONCLUSION :

As the angle of attack increases, the lift & drag forces also increases but as the AOA reaches a critical value called stall angle, then the lift force starts to reduce.

REFERENCES :

1. http://www.airfoiltools.com/
2. http://airfoiltools.com/airfoil/naca4digit
3. https://www.nasa.gov/ames/the-national-advisory-committee-for-aeronautics

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