Abstract
In this
experiment we used the pressure distribution to calculate the lift coefficient and
drag coefficient for different angles of attack for 0015 NACA airfoil and their
plots with the angles of attack and compared them with standard data.
Introduction
A finite wing
is the sum of an infinite airfoil cross sections in a two dimension. A pressure
imbalance in produced over the wing, it’s responsible of generating lift. This
pressure distribution is simply the pressure at all points around an airfoil.
The main forces
applied over a wing are pressure and shear. Normal and shear force are the
resultant of the forces over the wing. The normal force in perpendicular to the
chord line and the shear force in parallel to the chord line.
We can derive
these forces to obtain the total aerodynamic forces; the lift and the drag. The
lift component is perpendicular to the relative wind that makes an angle of
attack with the chord line. The drag component is parallel to the relative
wind.
By using a
micro-manometer and the pressure distribution over the airfoil we can obtain
the pressure coefficient which is used to calculate the lift characteristics of
out NACA 0015 airfoil.
Discussion
The values for
the lift coefficient for the small angles of attack are not accurate but
acceptable as experimental data for higher angles of attack
Our data show a
stall angle of 10 degrees at Cl max around 0.22, however the standard data
shows about 14 degrees at cl max around 1.2
The distance
between the curves is small for small angles (must be zero for angle zero) and
increase as AoA increase.
Cd max is at Cl
max as seen in the graphs
Conclusion
Large errors
are found in this experiment mainly human errors in reading the data from the
manometer and the lack of accuracy.
The pressure
distribution is very important to examine the airfoil characteristics and how
lift is
generated by the pressure imbalance.
generated by the pressure imbalance.
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