Airfoil Terminology and Definitions

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AIRFOIL Nomenclature, Classification, Flow and Pressure distribution

Airfoil Terminology and Definitions

Now, this is the best post to start with if you wanna be an Aeronautical Engineer. It's about the cross-sectional shape of the aircraft wing. Cross section of the aircraft wing is called as an Airfoil. Airfoils are streamlined shapes which are used in the aircraft to generate the lift and reduce the drag.

Nomenclature of the Airfoil

In the airfoil profile, the forward point is called the leading edge and the rearward point is called the trailing edge. The straight line connecting the leading and trailing edges is called the chord line of the airfoil. The distance from the leading edge to the trailing edge measured along the chord line is designated as a chord (c). The mean camber line is the locus of points midway between the lower surface and upper surface when measured normal to the mean camber line itself. The camber is the maximum distance between the mean camber line and the chord line, measured normal to the chord line. The thickness is the distance between the upper and lower surfaces also measured normal to the chord line. The shape of the airfoil at the leading edge is usually circular, with a leading edge radius of 0.02c, where c is the chord length. The upper and lower surfaces are also known as suction and pressure surfaces respectively.

Airfoil nomenclature


AIRFOIL CLASSIFICATION

The Network of Aquaculture Centres in Asia-Pacific, airfoil series, the 4-digit, 5-digit, and the updated 4-/5- digit, were generated using analytical equations and analogies that described the curvature of the airfoil's mean-line (geometric centerline) as well as the section's thickness distribution along the length. Also, the families, which included the 6-Series, were more complex shapes which were derived using theoretical methods.

(I) NACA Four-Digit Series:
The family of airfoils which was curated by utilizing this approach was called the NACA Four-Digit Series. Here in, the maximum camber in the percentage of the chord (airfoil length) is given by the first digit, the second indicates the position of the maximum camber and lastly, the maximum thickness of the airfoil in the percentage of the chord is provided by the last two numbers. For example, the NACA 2415 airfoil has a maximum thickness of 15% with a camber of 2% located at 40% chord from the airfoil leading edge (or 0.4c). Using these values, one can compute the coordinates of the entire airfoil using specific equations,

(II)NACA Five-Digit Series:
The NACA Five-Digit Series and the Four-Digit Series are quite similar as they use the same thickness forms, but the mean camber line is defined differently and the naming convention is a bit more complex. The design lift coefficient (cL) is given by the first digit, when multiplied by 3/2, yields it in tenths. The next two digits, when divided by 2, give the position of the maximum camber in tenths of the chord. The final two digits again indicate the maximum thickness in a percentage of chord. Taking an example, the NACA 24013 has a peak thickness of 13%, a design lift coefficient of 0.3, and the maximum camber located 20% behind the leading edge.

At present, the resources available for computation allow the designers to design and optimize the airfoils specifically tailored to a particular application.

FLOW OVER AIRFOIL

flow over airfoil


The air approaching the leading edge of an airfoil is first slowed down. It then speeds up again as it passes over or beneath the airfoil. As the velocity changes, the dynamic pressure changes and, according to Bernoulli's principle, the static pressure also changes. Air that is passing above and below the airfoil has speeded up to a value higher than the flight path velocity and will produce static pressures that are lower than ambient static pressure. The maximum velocity and minimum static pressure will occur at a point near-maximum thickness. The shape of the wing directly impacts the airflow. The differential pressure so produced when multiplied by the plan area of the airfoil generates an upward resultant force normal to chord line. Component of this resultant force normal to the relative wind direction is called “lift” and the component in the direction of the relative wind is called “drag”. The angle of attack (also called angle of incidence) (α) is the angle made between the chord line in the direction of airflow. For a given airfoil and at a given airflow velocity, the lift force increases with angle of attack up to a limit and then decreases. Reason for the decrease in lift beyond an angle of attack is “separation of flow” on the suction surface. The lift progressively decreases with an increase in angle of attack beyond the angle of attack corresponding to maximum lift. This is the principle behind the operation of spoilers and canards.

PRESSURE DISTRIBUTION

The pressure on the surface of an airfoil in flight is not uniform. The coefficient of pressure (Cp) can be defined with reference to free stream pressure far ahead of the airfoil as follows:

When α is zero, there are small regions near leading and trailing edges where Cp is positive but over most of the section, Cp is negative. The reduced pressure on the upper surface is tending to draw the section upwards while that on the lower surface is tending to draw the section downwards. With the pressure distribution as shown, the effect on the upper surface is larger, and hence there is a resultant force acting upwards on the airfoil which is the lift.


As the angle of attack is increased from zero we note that Pressure reduction on the upper surface increases both in intensity and extent until, at large angle of attack, it actually spreads to a small part of the front lower surface. The stagnation point advances backward on the lower surface and the pressure increase on the lower surface covers a greater part of the surface. The pressure reduction on the lower surface simultaneously decreases in both intensity and extent.

From the foregoing, following conclusions may be drawn:

At lower angles of attack, the lift arises from the difference between the pressure reduction on the upper and lower surfaces. At higher angles of attack, the lift is partly due to pressure reduction on the upper surface and partly due to pressure increase on the lower surface. At very high angles of incidence, the pressure reduction on the upper surface suddenly collapses due to flow separation and whatever lift that remains is due principally to the pressure increase on the lower surface. The increase and decrease of pressure distribution are greatest near the leading edge of the airfoil. If the distributed forces are replaced by a single resultant force, the point at which this resultant force would act is called the Centre of Pressure.

Since vertical component of resultant force is lift, we can assume that lift acts at the Centre of Pressure
As the angle of attack is varied the lift and drag change very rapidly due to changes in pressure distribution over the airfoil
Consequent to changes in pressure distribution due to change of angle of attack, there will be a movement of the center of pressure
As the angle of attack increases, the center of pressure tends to move towards the leading edge of the airfoil (wing) until the stalling angle is reached.
For subsonic flows, the center of pressure moves forward as the angle of attack increases until the stalling angle is reached. Consequently, the center of pressure tends to move backward along the chord.
For supersonic flows, The center of pressure tends to move backward as the aircraft's speed increases from low subsonic to high subsonic. It moves forward rapidly as soon as the airflow over the aircraft becomes sonic.
When the free-stream Mach number increases beyond M = 1, the center of pressure moves rearward and settles at around 50% of the chord at supersonic speeds.

Thanks for reading!

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