تقرير عن التحليل الديناميكي لتدفق الهواء حول جناح.pdf
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تقرير عن التحليل الديناميكي لتدفق الهواء حول جناح مروحه ريحيه
تقرير عن التحليل الديناميكي لتدفق الهواء حول جناح.pdf
- 1. الرحيم الرحمن هللا بسم University of Khartoum Faculty of engineering Mechanical Department Ms.c program renewable energy RE02 Numerical techniques Aerodynamic of a NACA 1410 Airfoil Prepare :- 1 Mohammed Abdelgadir Hassan Ibrahim 2 ALDOW Osman Ahmed
- 2. Abstract:- Through CFD we find the results of lift and drag coefficient (alpha from 0 to 20) and we make comparison with experimental data of aerofoil NACA1410. The comparisons Lift coefficient agrees within 2% of NACA published data Problem definition: An aerofoil NASA1410 moving within a velocity of 0.6774 Mach number Assumptions: • The fluid is ideal-gas. • Turbulence is taken into account. • The air is moving instead of aerofoil. • Velocity changes according to attack angle Select Aerofoil: We are selected aerofoil from naca 4digit program . number of aerofoil naca#1410 at 121 point . We save the coordinates of the points and add the coordinates of the axis value (z) of zero
- 3. Import points:- We import the points that have been saved in the desktop as file . txt
- 4. After that was drawn consisting of points for aerofoil in the window gampit By drawing geometry we connecting the points from the list of edges selected it nurbs
- 5. Flow Domain:- We drew Flow domain , which consists of a rectangular 20 * 20 and a semi-circle of radius 20 on the form of points connected edges and semi-circle by drawing a arc
- 6. Edges was converted into the back of the flow domain rectangular to the two faces and the transfer of semi-circle in front of the flow domain to a third face by convert the edges to faces
- 7. Mesh edges:- We have distributed the nodes consisting Mesh on each edge and chose the successive ratio, spacing 50 for width and then chose first lingth, spacing 60 for two lengths and then draw mesh for each face
- 8. Specific Boundary Type :- To determine the far field and the walls of aero foil from the list of zone selected specific boundary type, and we add edges contained far field and then edges of the aerofoil as walls
- 9. Export Mesh:- Select a solution from the list of solve selected fluent 5 / 6 and then we exported the mesh to the program fluent after that select the type of drawing 2D
- 10. Fluent Program :- We opened the fluent program, we have chosen 2 D version , and mode full simulations and click on Run , after that the program was open From file list we chose read case automatically program read our mesh Steps Analyses by using fluent program :- 1 We chose from a list of define model-solver
- 11. 2 From define list we selected solve_ viscous selected K- epsilon from viscous Model .
- 12. 3 from the define list we chose material we change density as Ideal gas 4 from define list select operating condition enter operating pressure 0 (Pascal)
- 13. 5 from define list select operating condition selected far field and set the far field (pressure , mach number , enter X-component of flow direction (cos α) and Y- component of flow direction(sin α) ) Calculation Mach number :- At 300 k kinematics viscosity of air = 15.68*10-6 D U D u * * * Re = = D=camber *2= 1mm *2= 2mm =0.002m Re from experimental data = 3*104 U= 235.2 m/s s m T R a / 19 . 347 4 . 1 * 287 * 300 * * = = = 6774 . 0 19 . 347 2 . 235 . = = = a U number Mach
- 14. 6 solve list we selected control – solution we set ( pressure 0.3, Density 1,Momentum 0.7 and discretization type (second order upwind ) also pressure –velocity coupling (simple )
- 15. 7 from solve list selected initialize – initialize enter gauge pressure = 101325 Pascal 8 from solve list selected monitor residual
- 16. 9 from solve list selected monitor – force 10 from solve list selected Iterate 11 from report list selected force – report drown left and drag force
- 17. Aerodynamic Curves:- number Angle CL CD 0 0 7613.329 1243.611 1 2 22329.61 851.0866 2 4 36026.47 -48.692 3 6 47708.99 -806.731 4 8 54434.18 -1202.63 5 10 57397.05 -1442.64 6 12 50247.85 -188.741 7 14 44632.18 552.1642 8 16 46456.29 926.9454 9 18 48906.65 1381.966 10 20 49785.22 1598.204 CL& Angle 0 10000 20000 30000 40000 50000 60000 70000 1 2 3 4 5 6 7 8 9 1 11 Angle Cl Figure Lift curve
- 18. CL , CD & angle -10000 0 10000 20000 30000 40000 50000 60000 70000 1 2 3 4 5 6 7 8 9 1 11 Angle CD Cl CD & angle -2000 -1500 -1000 -500 0 500 1000 1500 2000 1 2 3 4 5 6 7 8 9 1 11 Angle CD Figure Drag curve Figure Drag Lift curve
- 19. Results:- Stall angle 10 degrees for 30000 Re from simulation 12 degree for 30000 Re from NACA data Lift coefficient agrees within 2% of NACA published data Noticeable inaccuracies in drag coefficient data from the pressure ported airfoil Drag coefficient is Re dependent Lift Pressure Distribution Figure of Static pressure
- 20. Figure velocity magnitude Figure of static Temperature
- 21. Summary Objectives Study airflow over an airfoil Compare Results With NACA published data Stall angle is a function of the Reynolds number Lift coefficient relates closely to published data Drag coefficient highly dependent on Reynolds number