Flow over an airfoil Using SOLIDWORKS

Aim: Running simulations on NACA air foils 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:

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. Since a wing can have twist, a chord line of the whole wing may not be definable, so an alternate reference line is simply defined.

Procedure:

1. Import the required air foil for the data base provided in the solidworks flow simulation module
2. Model the air foil as required and save it as new file.
3. Now open the new wizard and set the required boundary conditions like external flow, fluid, etc etc
4. Now define the outputs required in the global goals and surface goals.
5. After completion of simulation save the velocity, drag force and lift force plots.
6. Repeat the same procedure for different angles of attack and compare the results.

The objective is to simulate the flow over an airfoil for different angle of attacks & compare drag force & lift force for each angle of attacks.

• The airfoil used is NACA0018.
• Angle of attacks are 0,2,4,6,8,10

1. For Angle of Attack 0:-

Goal Name	Unit	Value	Averaged Value	Minimum Value	Maximum Value	Progress [%]	Use In Convergence	Delta	Criteria
Drag Force	[N]	3286.371	3299.512	3275.659	3316.766	100	Yes	1.824112	80.9092
Lift Force	[N]	-1.12128	6.317244	-23.9505	44.62295	100	Yes	4.896537	2.492815

Pressure Distribution for aoa 0

Velocity Distribution for aoa 0

2. For Angle of Attack 2:-

Goal Name	Unit	Value	Averaged Value	Minimum Value	Maximum Value	Progress [%]	Use In Convergence	Delta	Criteria
Drag Force	[N]	3344.851	3349.688	3334.951	3370.839	100	Yes	2.138	79.551
Lift Force	[N]	1412.293	1431.389	1396.74	1471.711	100	Yes	6.251	28.264

Pressure Distribution for aoa 2

Velocity Distribution for aoa 2

3. For Angle of Attack 4:-

Goal Name	Unit	Value	Averaged Value	Minimum Value	Maximum Value	Progress [%]	Use In Convergence	Delta	Criteria
Drag Force	[N]	3464.634	3479.597	3458.613	3499.508	100	Yes	24.41269	83.10572
Lift Force	[N]	2841.503	2814.484	2755.455	2865.372	100	Yes	44.00535	51.51521

Pressure Distribution for aoa 4

Velocity Distribution for aoa 4

4. For Angle of Attack 6:-

Goal Name	Unit	Value	Averaged Value	Minimum Value	Maximum Value	Progress [%]	Use In Convergence	Delta	Criteria
Drag Force	[N]	3713.524	3702.089	3675.247	3721.587	100	Yes	4.309542	85.35508
Lift Force	[N]	4347.18	4319.094	4279.668	4348.548	100	Yes	68.8794	78.25289

Pressure Distribution for aoa 6

Velocity Distribution for aoa 6

5. For Angle of Attack 8:-

Goal Name	Unit	Value	Averaged Value	Minimum Value	Maximum Value	Progress [%]	Use In Convergence	Delta	Criteria
Drag Force	[N]	4005.415	4010.482	3995.032	4026.885	100	Yes	3.459469	93.33758
Lift Force	[N]	5817.398	5801.511	5784.122	5822.16	100	Yes	26.82007	102.4178

Pressure Distribution for aoa 8

Velocity Distribution for aoa 8

6. For Angle of Attack 10:-

Goal Name	Unit	Value	Averaged Value	Minimum Value	Maximum Value	Progress [%]	Use In Convergence	Delta	Criteria
Drag Force	[N]	4399.4	4414.281	4386.895	4435.148	100	Yes	3.005275	102.0666
Lift Force	[N]	7220.753	7206.045	7183.092	7234.538	100	Yes	19.98759	125.9684

Pressure Distribution for aoa 10

Velocity Distribution for aoa 10

Conclusions:

With the help of SOLIDWORKS Flow Simulation, successful analysis of the NACA-0017 airfoil has been carried at various angles of attack (0, 2, 4, 6, 8, 10 degrees) and following observation has been made.

• With the increase in angle of attack, the number of iterations required to give the result get reduced.
• With the increase in angle of attack, Drag force increases slowly and lift force increases rapidly.
• With the increase in angle of attack, the velocity of air over the airfoil increased

The greater the angle of attack, the more lift is generated by the wing. However, this is only true to a point. At some point the wing reaches its critical angle of attack, and at higher angles of attack the amount of lift generated drops dramatically. This is because the flow of air separates from the wing, forming turbulent eddies that destroy the ability of the wing to generate lift.Beyond the critical angle of attack is the stalled region, where the wing is no longer effectively generating lift and the airplane is stalled.