Aerofoil

Aerofoil (Airfoil):

An airfoil or aerofoil is a streamlined body or a a structure with curved surfaces designed to give the most favourable ratio of lift to drag in flight, used as the basic form of the wings, fins, and tailplanes of most aircraft.

Some terms related to airfoil are:

• Aerofoil Section

• Chord Line

• Camber line

• Angle of attack

• Stall

• Aspect ratio

Aerofoil Section:

The shape of a wing section is called aerofoil section.

Chord Line:

The line joining the leading and trailing edge of an airfoil is called chord and its length is called chord length.

Camber line:

The line connecting the mid-points of the upper and lower faces of the aerofoil section is called....Show More

Aerofoil (Airfoil):

An airfoil or aerofoil is a streamlined body or a a structure with curved surfaces designed to give the most favourable ratio of lift to drag in flight, used as the basic form of the wings, fins, and tailplanes of most aircraft.

Some terms related to airfoil are:

• Aerofoil Section

• Chord Line

• Camber line

• Angle of attack

• Stall

• Aspect ratio

Aerofoil Section:

The shape of a wing section is called aerofoil section.

Chord Line:

The line joining the leading and trailing edge of an airfoil is called chord and its length is called chord length.

Camber line:

The line connecting the mid-points of the upper and lower faces of the aerofoil section is called Camber line.

Angle of Attack:

It is the angle between the chord line and direction of the fluid stream.

Stall:

An airfoil is said to be in stall condition when the angle of attack of an airfoil is greater than the angle of attack at the minimum lift.

At stall, the air separates from the airfoil or wings and eddies are formed as a consequence of which there is a considerable increase in the drag coefficient.

Aspect Ratio:

The ratio of span of wing to its mean chord is called the aspect ratio.

I.e, A.R=`L/c`

Where, `L=`span of the wing

`C=` mean chord.

Since, `A=L*C`

`C=A/L`

Thus, A.R`=L^2/A`

We know the lift force is,

`F_l=rho*L*U*gamma`

Where, `gamma=` circulation developed on airfoil

`gamma=pi c*U*sinalpha`

where,`c=` chord length

`alpha=`angle of attack

`U=` free stream velocity of an air foil.

Thus, `F_l=rhoL*U*picUsinalpha`

`F_l=pi rho cLU^2sinalpha`...(i)

The lift force is also given by,

`F_l=C_L*(rho*U^2)/2*A`

`F_l=C_L*(rho*U^2)/2*C*L`..(ii)

From (i) and (ii),

`C_L=2pisinalpha`

NOTE: when a flying object such as aeroplane is in steady state, then

• Weight of aeroplane = lift force, 

• `W=C_l*(rho*U^2)/2*A`

• Thrust developed by the engine = the drag force