ADMIX

The CI-alpha curve of a positive cambered aerofoil intersects with the vertical axis of the 
CI-alpha graph:

above the origin.

Considering a positive cambered aerofoil, the pitch moment when CI=0 is:

negative (pitch-down).

The critical angle of attack:

remains unchanged regardless of gross weight.

When the lift coefficient CI of a symmetrical aerofoil section is zero, the pitching moment is:

zero.

When lift coefficient CI of a negatively cambered aerofoil section is zero, the pitching moment is:

nose up (positive)

The LIFT FORMULA can be written as:

L = CL * 1/2 rho * V² * S

Which of the following variables are required to calculate lift from the lift formula?

Dynamic pressure, lift coefficient and wing area.

Assuming standart atmospheric conditions, in order to generate the same amount of lift as altitude is increased, an aeroplane must be flown at:

a higher TAS for any given angle of attack.

Regarding the lift formula, density doubles, lift will 

also double.

Regarding the lift formula, if airspeed doubles, lift will:

be 4 times greater.

If the lift generated by a given wing is 1000 kN, what will be the lift if the wing area is doubled?

-> lift: 2000 kN.

If the wing area is increased, lift will 

increase because it is dircetly proportional to wing area.

AIRSPEED : AERODYNAMIC FORCE:

If the airspeed doubled, whilst maintaining the same control surface deflcetion the aerodynamic force on this control surface will:

becomes 4 times greater.

The lift coefficient CI versus angle of attack curve of a symmetrical aerofoil section intersects the vertical axis of the graph:

at the origin.

The lift coefficient CI versus angle of attack curve of a negatively cambered aerofoil section intersects the vertical axis of the graph:

below the origin.

The lift coefficient CI versus angle of attack curve of a positively cambered aerofoil intersects the horizontal axis of the graph:

to the left of the origin.

Regarding a positively cambered aerofoil section, whicht statement is correct?

The angle of attack has a negative value when the lift coefficient equals zero.
A nose down pitching moment exists when the lift coefficient equals zero.

Regarding a symmetric aerofoíl section.

The angle of attack is zero when the lift coefficient eqauls zero.

The pitching moment is zero when lift coefficient equals zero.

The correct drag formula can be written as:

D = CD 1/2 RHO V²* S

The polar curve of an aerofoil is a graphic realtion between:

CL and CD

The frontal area of a body, placed in a certrain airstream is increased by a factor 3.
The shape will not alter. The aerodynamic drag will increase with a factor 3.

Aerodynamic drag:

A body is placed in a certrain airstram. The airstream velocity increases by a factor 4. The aerodynamic drag will increase with a factor : 16!!

density of airsream decreases to half:

A body is placed in a certrain airstream. The density of airstream decreases to half of the original value. The aerodynamic drag will increase with a factor: 2!

The aerofoil polar is:

a graph of the realtion between lift coefficient and drag coefficient.

When an aircraft with a typical aerofoil is in level flight at low speed and high angle of attack, the normal axis is:

nearly vertical.

comparing the lift and drag coefficient for convential aeroplanes:

CL is much greater than CD.

Increasing dynamic pressure will have the following effect on the total drag of an aeroplane:

at speeds above the minimum drag speed, total drag increases.

Increasing air destiny will have the following effect on the drag of a body in an airstram
(angel of attack and TAS are constant):

Drag increases.

Minimum Drag of an aeroplane in straight and level flight occurs at the:

maximum CL-CD ratio.

Variables Drag Fromula:

Which variables are requirede to calculate drag from the drag fromula?

Dynamic pressure, drag coefficient and wing area.

A positive cambered aerofoil generates zero lift:

at a negative angle of attack.

Assuming ISA Conditions, and no compressibility effects, if an aeroplane maintains straight and level flight at the same anlge of attack at two different altitudes, the:

TAS is higher at the higher altitude.

Assuming ISA Conditions, and no compressibility effects, if an aeroplane maintains straight and level flight at the same anlge of attack at two different altitudes, the:

IAS is the same at both altitudes.

Assuming ISA Conditions, and no compressibility effects, if an aeroplane maintains straight and level flight at the same anlge of attack at two different altitudes, the:

TAS is lower at the lower altitude.

An Aeroplanes AoA is the angle between:

 speed vector and lonitudinal axis.

Which of the following planforms gives the HIGHEST local profile lift coefficient at the wingroot?

Rectangular

The fundemental difference between the areodynamic charecteristics of two and three-dimensional flow is that, in a three-dimensional flow about a wing:

a spanwise component exists in addition to the chordwise speed component.

flow on the upper surface:

The flow on the upper surface of the wing has a component in wing root dircetion.

Airplane accerlerates from 80 to 160 with load factor 1.

The induced drag coefficient alters with factor 1:16 

and induced drag with factor 1/4.

increasing AR on induced Drag:

it is reduced because the effect of wing-tip vortices is reduced.