Question 260-1 : What may happen if the 'ultimate load factor' is exceeded ? [ Theoretical lift off ]

Question 260-2 : The increase in stall speed ias with increasing altitude is due to ?

Compressibility effects.

The ias stall speed is considered constant at low altitudes because of the tas stall speed increases due to the decreasing altitude but the ias is almost constant this is because the factor increasing the tas stall speed density also affects the ias speed in the same amounthowever at higuer altitudes we consider another factor the compresibility that affects the anemometer and hence increases the reading of iasresuming the ias stall speed is considered constant at lower altitudes below 10000 ft and increases with altitude at higuer altitudes due to compresibility effectshigher altitude > lower air density + lower air viscosity because of lower temperatures > lower reynolds number > less kinematic energy in the boundary layer to oppose adverse pressure gradients > earlier flow separation > lower clmax and aoacrit > higher stall speed

Question 260-3 : An large jet transport aeroplane has the following four flap positions up take off approach and landing and two slat positions retracted and extended generally speaking the selection that provides the highest positive contribution to clmax is ?

Slats from retracted to extended.

When trailing edge flaps are deployed from up to take off there will be a comparatively large increase in cl and a small increase in cd but with each successive increase in flap angle the increase in cl coefficient of lift will be less whereas the increase in cd coefficient of drag will be greaterwhen considering just trailing edge flaps in isolation it could be said that the greatest increase in clmax will be when the flaps move from up to take offdata shows that the most efficient trailing edge flap triple slotted fowler flap increases clmax by about 100% but only when fully deployed we can only speculate on the increment in clmax from up to take offhowever when considering just slats in isolation the same data shows the increase in clmax for the slat from retracted to extended is about 60%of the four answers offered it can be said that the slats moving from retracted to extended would be the selection that will provide the highest positive contribution to clmax

Question 260-4 : In a straight steady climb the thrust must be ?

Greater than the drag because it must also balance a component of weight.

Question 260-5 : Regarding a positively cambered aerofoil section which statement is correct i the angle of attack has a negative value when the lift coefficient equals zeroii a nose down pitching moment exists when the lift coefficient equals zero ?

I is correct and ii is correct.

com encom080 1030ajpg com encom080 1030bjpgfor a positively cambered aerofoil section the angle of attack has a negative value when the lift coefficient equals zeroexample a positively cambered aerofoil section com encom080 909jpgin this example lift coefficient = 0 when angle of attack = 4° the angle of attack is negative and the pitching moment is nose down

Question 260-6 : Regarding a symmetric aerofoil section which statement is correct i the angle of attack is zero when the lift coefficient equals zeroii the pitching moment is zero when the lift coefficient equals zero ?

I is correct and ii is correct.

com encom080 687ajpga symmetrical aerofoil needs to have a positive pitch to produce lift if the pitching moment is zero the lift is zero

Question 260-7 : Regarding a symmetric aerofoil section which statement is correct i the angle of attack has a positive value when the lift coefficient equals zeroii a nose down pitching moment exists when the lift coefficient equals zero ?

I is incorrect and ii is incorrect.

com encom080 687ajpga symmetrical aerofoil needs to have a positive pitch to produce lift if the pitching moment is zero the lift is zero

Question 260-8 : Which of these statements about wing sweepback are correct or incorrect i increasing wing sweepback increases mcritii increasing wing sweepback increases the drag divergence mach number ?

I is correct ii is correct.

What is mcrit critical mach number com encom080 1049ajpgthe critical mach number of an aeroplane is the highest speed possible without supersonic flow over the wing what is sweepback most of the difficulties of transonic flight are associated with shock wave induced flow separation therefore any means of delaying or alleviating the shock induced separation improves aerodynamic performance one method is wing sweepback sweepback theory is based upon the concept that it is only the component of the airflow perpendicular to the leading edge of the wing that affects pressure distribution and formation of shock waveson a straight wing aircraft the airflow strikes the wing leading edge at 90° and its full impact produces pressure and lift a wing with sweepback is struck by the same airflow at an angle smaller than 90° this airflow on the swept wing has the effect of persuading the wing into believing that it is flying slower than it really is thus the formation of shock waves is delayed advantages of wing sweep include an increase in critical mach number force divergence mach number and the mach number at which drag rises peaks in other words sweep delays the onset of compressibility effects what is drag divergence mach number the drag divergence mach number not to be confused with critical mach number is the mach number at which the aerodynamic drag on an airfoil or airframe begins to increase rapidly as the mach number continues to increase this increase can cause the drag coefficient to rise to more than ten times its low speed valueincreasing wing sweepback increases the drag divergence mach number com encom080 1049bjpgthe effect of increasing angle of sweep is an increase in the critical mach number and an increase in the drag divergence mach number

Question 260-9 : Considering subsonic incompressible airflow through a venturi which statement is correct i the dynamic pressure in the throat is higher than in the undisturbed airflowii the total pressure in the throat is the same as in the undisturbed airflow ?

I is correct ii is correct.

Img com encom080 815jpgstatic pressure decreases in a venturi and airflow speed increases

Question 260-10 : Which of these statements about the strength of wing tip vortices are correct or incorrect i assuming no flow separation the strength of wing tip vortices decreases as the angle of attack decreasesii the strength of wing tip vortices increases as the aspect ratio decreases ?

I is correct ii is correct.

Formation of vortices the action of the airfoil that gives an aircraft lift also causes induced drag when an airfoil is flown at a positive aoa a pressure differential exists between the upper and lower surfaces of the airfoil the pressure above the wing is less than atmospheric pressure and the pressure below the wing is equal to or greater than atmospheric pressure since air always moves from high pressure toward low pressure and the path of least resistance is toward the airfoil's tips there is a spanwise movement of air from the bottom of the airfoil outward from the fuselage around the tips this flow of air results in 'spillage' over the tips thereby setting up a whirlpool of air called a 'vortex'just as lift increases with an increase in aoa induced drag also increases this occurs because as the aoa is increased there is a greater pressure difference between the top and bottom of the airfoil and a greater lateral flow of air consequently this causes more violent vortices to be set up resulting in more turbulence and more induced dragthe strength of the wingtip vortices is determined by the magnitude of the pressure difference between the upper and lower wing surfaces and by the time during which this pressure difference acts to drive the air into the vortices the greater the pressure difference or the longer the time available the stronger the vortices will bethe time during which the pressure difference acts upon the air is proportional to the time it takes the air to pass from the leading edge to the trailing edge of the wing so for any given airspeed the length of the wing chord determines the time taken this means that decreasing the chord length will decrease the vortex strength which will decrease the stalling angle and the dragthe intensity or strength of the vortices is directly proportional to the weight of the aircraft and inversely proportional to the wingspan and speed of the aircraftgood answers for this kind of questions assuming no flow separation the strength of wing tip vortices increases as the angle of attack increasesassuming no flow separation the strength of wing tip vortices decreases as the angle of attack decreasesthe strength of wing tip vortices decreases as the aspect ratio increasesthe strength of wing tip vortices increases as the aspect ratio decreases

Question 260-11 : Considering subsonic incompressible airflow through a venturi which statement is correct i the dynamic pressure in the undisturbed airflow is higher than in the throatii the total pressure in the undisturbed airflow is lower than in the throat ?

I is incorrect ii is incorrect.

Img com encom080 815jpgstatic pressure decreases in a venturi and airflow speed increases

Question 260-12 : Assuming no flow separation which of these statements about the flow around an aerofoil as the angle of attack decreases are correct or incorrect i the stagnation point moves upii the point of lowest static pressure moves aft ?

I is correct ii is correct.

Question 260-13 : Considering subsonic incompressible airflow through a venturi which statement is correct i the dynamic pressure in the undisturbed airflow is lower than in the throatii the total pressure in the undisturbed airflow is higher than in the throat ?

I is correct ii is incorrect.

Img com encom080 815jpgstatic pressure decreases in a venturi and airflow speed increases

Question 260-14 : Assuming isa conditions and no compressibility effects if an aeroplane maintains straight and level flight at the same angle of attack at two different altitudes the ?

Tas is lower at the lower altitude.

In straight and level flight lift does not change as it is only balancing against weightaircraft is maintaining level flight and therefore lift cannot change lift = cl 12rho v² scl = lift coefficientrho = densityv = tas in ms s = surfaceat the same angle of attack same cl with 'rho' increasing with decreasing altitude since surface 's' does not change only tas can be must be reduced

Question 260-15 : The main purpose of a boundary layer fence on a swept wing is to ?

Improve the low speed handling characteristics.

The spanwise flow encountered with swept wings may be reduced by the use of stall fences which are thin plates parallel to the axis of symmetry of the airplane com encom080 1207jpgin this manner a strong boundary layer buildup over the ailerons is preventednotice recent aircrafts do not use boundary layer fence anymore

Question 260-16 : The stall speed decreases all other relevant factors are constant ?

When during a manoeuvre the aeroplane nose is suddenly pushed firmly downwards eg as in a push over .

When flaps are retracted the stall speed increases you will stall at a higher speed than the speed with flaps deployed stall speed is based on 1g flight when you do a push over load factor decreasesstall speed changes at the square root of the load factorvs at 1g = 100 ktif load factor is 05g now stall speed is square root 05 = 07 x 100 kt = 70 kt

Question 260-17 : The load factor is greater than 1 one ?

When lift is greater than weight.

Dividing lift by weight gives load factorload factor increases in a turn and in a pull up manoeuvre those actions require more lift

Question 260-18 : Which of these statements about weight or mass is correct ?

In the si system the unit of measurement for weight is the newton.

The si unit of weight is the newton n = kgms²

Question 260-19 : What is the effect of winglets on the drag of the wing ?

Increase parasite drag decrease induced drag.

Question 260-20 : If the wing area is increased lift will ?

Increase because it is directly proportional to wing area.

Question 260-21 : Which of these statements about stall speed is correct ?

Increasing forward sweep increases stall speed.

The effect of increasing wing sweep angle is to decrease the lift curve slope and clmax as depicted below com encom080 999jpgwhile the swept wing offers a strong transonic drag reduction the lift penalty at high angle of attack is substantial as a wing is swept clmax decreases com encom080 1226jpgspanwise airflow over a forward swept wing is the reverse of a conventional swept wing but the stall speed increases or decreases in a same way

Question 260-22 : The lift formula can be written as rho = density ?

L = cl * 12rho * v² * s.

Where 12 rho is dynamic pressure cl is coefficient of lift v is speed in ms s is wing area

Question 260-23 : During a straight steady climb and with the thrust force parallel to the flight path ?

Lift is the same as during a descent at the same angle and mass.

During climb and descent lift is lower than weightlift is a vertical vectorlift in climb or descent = weight x cos climb angle

Question 260-24 : Assuming no compressibility effects induced drag at constant ias is affected by ?

Aeroplane mass.

Question 260-25 : Given that pstat = static pressurerho = densitypdyn = dynamic pressureptot = total pressurebernoulli's equation reads as follows ?

Pstat + 12rho * tas² = constant.

Dynamic pressure is '12 rho tas²'bernoulli's theorem is pt = ps + pdthen if pd inscreases ps decreases as pt always remains constant

Question 260-26 : The point in the diagram giving the lowest speed in unaccelerated flight is err a 081 1265 ?

Point 4.

Img com encom080 1264jpg

Question 260-27 : The point in the annex showing zero lift is err a 081 1269 ?

Point a.

Question 260-28 : If the airspeed reduces in level flight below the speed for maximum ld the total drag of an aeroplane will ?

Question 260-29 : Slat or flap asymmetry occurring after either extension or retraction may have an effect on controllability since ?

Slat asymmetry causes a yawing moment whereas flap asymmetry causes a large rolling moment.

At high speed the angle of attack will be small and with an asymmetric leading edge slat deployment there will be little change in cl but cd would increase on the deployed slat turbulent boundary layer this causes yaw which can be controlled by rudderhowever at lower speeds and high angle of attack a large change in clmax will occur between the wings with a flap asymmetry this would then lead to a strong rolling moment

Question 260-30 : Which of the following variables are required to calculate lift from the lift formula ?

Dynamic pressure lift coefficient and wing area.

Lift formula = 12 rho cl x v² x swhere '12 rho v²' is dynamic pressure rho = density v = speed in ms 'cl' is coefficient of lift's' is wing area

Question 260-31 : Which of these statements about weight or mass is correct ?

Weight is a force.

Question 260-32 : The induced angle of attack is ?

The angle by which the relative airflow is deflected due to downwash.

The angle between the chord line and effective airflow is the effective angle of attack the angle between effective airflow and the horizontal flow that was the relative airflow is called the induced angle of attack which is equal to the downwash angle

Question 260-33 : When an aeroplane enters ground effect ?

The lift is increased and the drag is decreased.

Question 260-34 : Which of these statements about weight or mass is correct ?

The mass of a body can be determined by dividing its weight by the acceleration due to gravity.

Question 260-35 : Which of these statements about weight or mass is correct ?

The weight of a body can be determined by multiplying its mass by the acceleration due to gravity.

Question 260-36 : From a polar curve of the entire aeroplane one can read ?

The maximum clcd ratio and maximum lift coefficient.

Question 260-37 : The diagram shows the parameter x versus tas if a horizontal flight is considered the axis x shows err a 081 1319 ?

The coefficient of lift.

New question cqb 15 august 2011this question already exists in the database with answer 'induced drag''lift formula lift = 12 rho cl x v² x swhere '12 rho' is dynamic pressure'cl' is coefficient of lift'v' is speed in ms's' is wing areaif the speed is increase angle of attack has to be reduced to maintain lift contant coefficient of lift will reduce as the speed increases

Question 260-38 : The additional increase in drag at mach numbers above the critical mach number is due to ?

Wave drag.

Question 260-39 : Air passes a normal shock wave which of the following statements is correct ?

The static temperature increases.

Img com ennormal and oblic shock wavejpg

Question 260-40 : The bow wave will first appear at ?

A mach number just above m = 1.

A bow wave is a shock wave that forms when the aircraft is flying at a speed just faster than the speed of soundthe bow wave is so named because it is similar to a wave inwater off the bow of a boat com encom080 41jpgthe wave is in front of the body