Abstract
In this
experiment we used the pressure distribution to calculate the center of
pressure, the lift coefficient, drag coefficient and the moment coefficient for
different angles of attack for symmetrical 0020 NACA airfoil using an air flow
bench and an airfoil with tappings, and their plots with the angles of attack
and compared them with different data such as XFlR5 and standard data from
NACA.
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.
Unlike the
cambered airfoil, a symmetrical airfoil produces no lift on an angle of attack
of zero degree.
“The center of
pressure of an aircraft is the point where all of the aerodynamic pressure field
may be represented by a single force vector with no moment. A similar idea
is the aerodynamic center which is the point on an airfoil where the pitching moment produced
by the aerodynamic forces is constant with angle of attack “[Wikipedia]
For symmetrical
airfoil, when angle of attack varies, the pressure distribution will also vary.
But for any angle of attack if we consider the resultant lift force position or
center of pressure it will lie on the aerodynamic center. It won’t vary with
angle of attack. Because aerodynamic center is at constant location and
usually lies at 0.25 of the chord.
Aerodynamic
center and Center of Pressure are same in symmetrical airfoil.
Discussion
The moment coefficient increases as
the angle of attack increases
The center of pressure should not be
changed due to the symmetry of the airfoil, and it should be located at 25% of
the chord
Minimum drag coef. was between -5
and 5 degrees (AoA)
Maximum lift coef. was between 5 and
10 degrees for Re= 50,000 / 100,000
Conclusion
The data from the XFlR5 and NACA
tools are identical stall angle 10 at cl max 1.1
The data taken showed that the angle
of stall was at 20 at cl max 2.7
Error is large 145%, 100%
The data taken from the experiment
was hard to solve to get the coefficient, using the xflr is more convenient.
Also human error should be taken
into consideration in this experiment.
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