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When speed increases in straight and level flight on a positively cambered aerofoil, what happens to the:
I. centre of pressure
II. magnitude of the total lift?
I. C.P moves aft
II. remains constant
Assuming no flow effect and no compressibility effects the location of centre of pressure of a positively cambered aerofoil section:
Centre of pressure moves backwards when AoA decreases.
Centre of pressure moves backwards when AoA decreases.
Assuming no flow seperation and no compressibility effects the location of the aerodynamic centre of an aerofoil section:
is independent of angle of attack.
aerodynamic centre is independent of AoA.
Assuming no flow seperation and no compressibility effects the location of the centre of pressure of a symmetrical aerofoil section:
is at approximately 25% chord irrespective* of angle of attack.
*rücksichtslos
centre of pressure of a symmetrical = 25% chord irrespective of AoA.
Assuming no flow seperation and no compressibility effects the location of the
centre of pressure of a symmetrical aerofoil section:
is independent of angle of attack.
centre of pressure of a symmetrical aerofoil section :
is either at approximately 25% chord irrespective of AoA or is independent of AoA!
The lift force acting on aerofoil:
(no flow seperation)
is mainly caused by suction* on the upperside of the aerofoil.
*Absaugung
The lift force acts upperside of the aerofoil because of suction.
Lift is generated when:
the flow dircetion of a certrain mass of air is changed.
see above.
correct statement about the CI and angle of attack:
for a symmetric aerofoil, if angle of attack = 0, CI = 0!
symmetric aerofoil: AoA = 0, CI = 0.
The CI-alpha curve of a positive cambered aerofoil intersects with the vertical axis of the
CI-alpha graph:
above the origin.
CI alpha curve - positive cambered aerofoil... above the origin.
Considering a positive cambered aerofoil, the pitch moment when CI=0 is:
negative (pitch-down).
CI = 0, positive cambered aerofoil = negative.
The critical angle of attack:
remains unchanged regardless of gross weight.
Critical AoA bleibt unverändert ohne Rücksicht auf das grobe Gewicht.
When the lift coefficient CI of a symmetrical aerofoil section is zero, the pitching moment is:
zero.
CI of a symmetrical aerofoil section 0 = pitching moment is also 0.
When lift coefficient CI of a negatively cambered aerofoil section is zero, the pitching moment is:
nose up (positive)
negatively cambered aerofoil = pitching moment positive!
The LIFT FORMULA can be written as:
L = CL * 1/2 rho * V² * S
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.
Zur Berechnung des Auftriebformels benötigen wir:
Dynamic Pressure, lift coeffiecient 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.
Unter Berücksichtigung der Standart Atmosphäre, zwecks Erzeugung der selben Summe des Auftriebs wobei die Höhe steigt, muss das Flugzeug mit
einem höheren TAS zu gegebenen AoA fliegen.
Regarding the lift formula, density doubles, lift will
also double.
density ~ lift
Bsp: Dichte verdoppelt = Auftrieb verdeoppelt sich auch.
Regarding the lift formula, if airspeed doubles, lift will:
be 4 times greater.
Airspeed double - lift 4times greater.
Airespeed verdoppelt sich, Auftrieb vervierfacht sich.
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.
Wing area doubles, lift doubles.
If the wing area is increased, lift will
increase because it is dircetly proportional to wing area.
Lift ~ wing area.
(proportional)
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.
Airspeed 1 : Aerodynamic force 4.
The lift coefficient CI versus angle of attack curve of a symmetrical aerofoil section intersects the vertical axis of the graph:
at the origin.
curve of a symmetrical aerofoil section = 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.
curve of a negatively cambered = 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.
curve of a positively cambered = 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.
positive cambered aerofoil section:
NEGATIVE value AoA
A nose down pitching moment.
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.
symmetric aerofoil section:
AoA is zero, when CI = 0
Pitching moment is zero, when CI = 0.
The correct drag formula can be written as:
D = CD 1/2 RHO V²* S
Widerstandsformel:
D = CD 1/2 RHO V²* S
The polar curve of an aerofoil is a graphic realtion between:
CL and CD
CL and CD
"Lift and Drag"
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.
airstream 3 : drag increase 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!!
airspeed 4 : aerodynamic drag 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!
density airspeed halbiert = aerodynamic drag verdoppelt.
The aerofoil polar is:
a graph of the realtion between lift coefficient and drag coefficient.
aerofoil polar = graph lift 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.
low speed, high AoA = normal axis nearly vertical
comparing the lift and drag coefficient for convential aeroplanes:
CL is much greater than CD.
CL > 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 dynamic pressure at speeds above 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.
increasing density = increasing drag.
Minimum Drag of an aeroplane in straight and level flight occurs at the:
maximum CL-CD ratio.
Minimum Drag : 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.
relevatn variables for drag formula:
dynamic pressure, CD, Awing
A positive cambered aerofoil generates zero lift:
at a negative angle of attack.
Positiv cambered airfoil + negative AoA = zero lift!
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.
different altitudes, same AoA: higher TAS at higher altitude.
