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The lift coefficient cl versus angle of attack curve of a negatively cambered ? [ Formation assignment ]

Question 237-1 : To the right of the origin at the origin to the left of the origin nowhere

. /com en/com080 984 jpg.a negatively cambered aerofoil section produces lift at positive angle of attack only in our example at approximately 3° angle of attack . stanley .hi there was no graph with the question best regards ..this is normal there is no graph with this question at the exam you need to know the shape of the cl/aoa curve exemple 337 To the right of the origin.To the right of the origin.

Regarding the lift formula if density doubles lift will ?

Question 237-2 : Also double be 4 times greater halve remain the same

Lift formula .lift = 1/2 rho cl x v² x s..where .'1/2 rho' is dynamic pressure .'cl' is coefficient of lift .'v' is speed in m/s .'s' is wing area ..if density is double lift will also double exemple 341 Also double.Also double.

Assuming all bodies have the same cross sectional area and are in motion which ?

Question 237-3 : Body b body a body c body d

. /com en/com080 986 jpg. exemple 345 Body b.Body b.

When a wing spoiler is extended at constant angle of attack ?

Question 237-4 : Drag increases but lift decreases both drag and lift increase drag increases but lift remains constant both drag and lift decrease

.when you deploy the spoilers you will need to increase angle of attack to restore the coefficient of lift to that required at the lower speed the aircraft will be traveling .however if the angle of attack is unchanged the aircraft will no longer produce sufficient lift to oppose the weight and therefore the aircraft will descend .right answer drag increases but lift decreases exemple 349 Drag increases but lift decreases.Drag increases but lift decreases.

The point where the single resultant aerodynamic force acts on an aerofoil is ?

Question 237-5 : Centre of pressure centre of gravity neutral point aerodynamic centre

.the pressure created by an aerofoil at any point may be represented by a vector at right angles to its surface whose length is proportional to the difference between absolute pressure at the point and the free stream static pressure .all of them can be represented by a single vector acting at a particular point called the centre of pressure . 669.the centre of pressure is a theoretical point on the chord line through which the resultant of all forces the total reaction is said to act .its position is usually around 25% of the way from the leading edge simply because more lift is generated there but it moves steadily forward as the angle of attack is increased until just before the stalling angle when it moves rapidly backwards the centre of pressure's most forward point is just before the stalling angle this is why an aeroplane's nose drops when the wings stall and the centre of pressure moves behind the cg .thus when speed is increased in straight and level flight on a positively cambered aerofoil you have to decrease the angle of attack to keep the the total lift force constant and the point where the resultant of all forces are acting the centre of pressure moves aft exemple 353 Centre of pressure.Centre of pressure.

Wing loading is the ratio between ?

Question 237-6 : Aeroplane weight and wing area chord and aeroplane weight aeroplane weight and lift coefficient aeroplane weight and wing span

. /com en/com080 992 jpg.wing loading is the number of kilos of aircraft weight supported by each square foot of wing area wing loading is usually expressed as the value when the airplane is at its maximum certified weight so the actual wing loading will vary depending on fuel and payload onboard the airplane at any point .wing loading = 'aircraft weight' divided by 'wing area' exemple 357 Aeroplane weight and wing area.Aeroplane weight and wing area.

Increasing the aspect ratio of a wing ?

Question 237-7 : Decreases induced drag decreases gust load increases stall speed increases induced drag

exemple 361 Decreases induced drag.Decreases induced drag.

What may happen if the 'ultimate load factor' is exceeded ?

Question 237-8 : Structural failure elastic or temporary deformation only no structural failure only plastic or permanent deformation flutter

exemple 365 Structural failure.Structural failure.

The increase in stall speed ias with increasing altitude is due to ?

Question 237-9 : Compressibility effects exceedance of mcrit an increase in tas the larger angle of attack necessary in lower density air to obtain the same lift as at sea level

.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 amount .however at higuer altitudes we consider another factor the compresibility that affects the anemometer and hence increases the reading of ias .resuming the ias stall speed is considered constant at lower altitudes below 10000 ft and increases with altitude at higuer altitudes due to compresibility effects .higher 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 exemple 369 Compressibility effects.Compressibility effects.

An large jet transport aeroplane has the following four flap positions up take ?

Question 237-10 : Slats from retracted to extended flaps from approach to landing flaps from up to take off flaps from take off to approach

.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 greater ..when 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 off ..data 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 off ..however 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 exemple 373 Slats from retracted to extended.Slats from retracted to extended.

In a straight steady climb the thrust must be ?

Question 237-11 : Greater than the drag because it must also balance a component of weight greater than drag because more lift has to be produced lower than the drag because it is assisted by a component of weight equal to the drag

exemple 377 Greater than the drag because it must also balance a component of weight.Greater than the drag because it must also balance a component of weight.

Regarding a positively cambered aerofoil section which statement is correct .i ?

Question 237-12 : I is correct and ii is correct i is correct and ii is incorrect i is incorrect and ii is correct i is incorrect and ii is incorrect

. /com en/com080 1030a jpg. /com en/com080 1030b jpg.for a positively cambered aerofoil section the angle of attack has a negative value when the lift coefficient equals zero .example a positively cambered aerofoil section . /com en/com080 909 jpg.in this example lift coeffient = 0 when angle of attack = 4° the angle of attack has a negative and the pitching moment is nose down exemple 381 I is correct and ii is correct.I is correct and ii is correct.

Regarding a symmetric aerofoil section which statement is correct .i the angle ?

Question 237-13 : I is correct and ii is correct i is correct and ii is incorrect i is incorrect and ii is incorrect i is incorrect and ii is correct

. /com en/com080 687a jpg.a symmetrical aerofoil needs to have a positive pitch to produce lift if the pitching moment is zero the lift is zero exemple 385 I is correct and ii is correct.I is correct and ii is correct.

Regarding a symmetric aerofoil section which statement is correct .i the angle ?

Question 237-14 : I is incorrect and ii is incorrect i is correct and ii is incorrect i is incorrect and ii is correct i is correct and ii is correct

. /com en/com080 687a jpg.a symmetrical aerofoil needs to have a positive pitch to produce lift if the pitching moment is zero the lift is zero exemple 389 I is incorrect and ii is incorrect.I is incorrect and ii is incorrect.

Which of these statements about wing sweepback are correct or incorrect .i ?

Question 237-15 : I is correct ii is correct i is correct ii is incorrect i is incorrect ii is correct i is incorrect ii is incorrect

. what is mcrit critical mach number . /com en/com080 1049a jpg.the 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 waves .on 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 value .increasing wing sweepback increases the drag divergence mach number . /com en/com080 1049b jpg.the effect of increasing angle of sweep is an increase in the critical mach number and an increase in the drag divergence mach number exemple 393 I is correct, ii is correct.I is correct, ii is correct.

Considering subsonic incompressible airflow through a venturi which statement ?

Question 237-16 : I is correct ii is correct i is correct ii is incorrect i is incorrect ii is correct i is incorrect ii is incorrect

Img /com en/com080 815 jpg.static pressure decreases in a venturi and airflow speed increases exemple 397 I is correct, ii is correct.I is correct, ii is correct.

Which of these statements about the strength of wing tip vortices are correct ?

Question 237-17 : I is correct ii is correct i is correct ii is incorrect i is incorrect ii is correct i is incorrect ii is incorrect

.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 drag .the 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 be .the 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 drag .the 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 aircraft .good answers for this kind of questions .assuming no flow separation the strength of wing tip vortices increases as the angle of attack increases .assuming no flow separation the strength of wing tip vortices decreases as the angle of attack decreases .the strength of wing tip vortices decreases as the aspect ratio increases .the strength of wing tip vortices increases as the aspect ratio decreases exemple 401 I is correct, ii is correct.I is correct, ii is correct.

Considering subsonic incompressible airflow through a venturi which statement ?

Question 237-18 : I is incorrect ii is incorrect i is correct ii is incorrect i is correct ii is correct i is incorrect ii is correct

Img /com en/com080 815 jpg.static pressure decreases in a venturi and airflow speed increases exemple 405 I is incorrect, ii is incorrect.I is incorrect, ii is incorrect.

Assuming no flow separation which of these statements about the flow around an ?

Question 237-19 : I is correct ii is correct i is incorrect ii is incorrect i is incorrect ii is correct i is correct ii is incorrect

exemple 409 I is correct, ii is correct.I is correct, ii is correct.

Considering subsonic incompressible airflow through a venturi which statement ?

Question 237-20 : I is correct ii is incorrect i is incorrect ii is incorrect i is correct ii is correct i is incorrect ii is correct

Img /com en/com080 815 jpg.static pressure decreases in a venturi and airflow speed increases exemple 413 I is correct, ii is incorrect.I is correct, ii is incorrect.

Assuming isa conditions and no compressibility effects if an aeroplane ?

Question 237-21 : Tas is lower at the lower altitude ias is lower at the lower altitude ias is higher at the lower altitude tas is higher at the lower altitude

.in straight and level flight lift does not change as it is only balancing against weight ..aircraft is maintaining level flight and therefore lift cannot change ..lift = cl 1/2rho v² s..cl = lift coefficient.rho = density.v = tas in m/s .s = surface..at 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 exemple 417 Tas is lower at the lower altitude.Tas is lower at the lower altitude.

The main purpose of a boundary layer fence on a swept wing is to ?

Question 237-22 : Improve the low speed handling characteristics improve the high speed handling characteristics increase the critical mach number improve the lift coefficient of the trailing edge flaps

.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 en/com080 1207 jpg.in this manner a strong boundary layer buildup over the ailerons is prevented .notice recent aircrafts do not use boundary layer fence anymore exemple 421 Improve the low speed handling characteristics.Improve the low speed handling characteristics.

The stall speed decreases . all other relevant factors are constant ?

Question 237-23 : When during a manoeuvre the aeroplane nose is suddenly pushed firmly downwards e g as in a push over in a horizontal turn when flaps are retracted when the cg is moved forward

.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 decreases .stall speed changes at the square root of the load factor .vs at 1g = 100 kt .if load factor is 0 5g now stall speed is square root 0 5 = 0 7 x 100 kt = 70 kt exemple 425 When, during a manoeuvre, the aeroplane nose is suddenly pushed firmly downwards (e.g. as in a push over).When, during a manoeuvre, the aeroplane nose is suddenly pushed firmly downwards (e.g. as in a push over).

The load factor is greater than 1 one ?

Question 237-24 : When lift is greater than weight in steady wings level horizontal flight during a wings level stall before recovery when lift is less than weight

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

Which of these statements about weight or mass is correct ?

Question 237-25 : In the si system the unit of measurement for weight is the newton in the si system the unit of measurement for weight is the kilogram the weight of an object is independent of the acceleration due to gravity the mass of an object depends on the acceleration due to gravity

The si unit of weight is the newton n = kg m/s² exemple 433 In the si system the unit of measurement for weight is the newton.In the si system the unit of measurement for weight is the newton.

What is the effect of winglets on the drag of the wing ?

Question 237-26 : Increase parasite drag decrease induced drag increase induced drag decrease friction drag increase friction drag decrease form drag increase induced drag decrease interference drag

exemple 437 Increase parasite drag, decrease induced drag.Increase parasite drag, decrease induced drag.

If the wing area is increased lift will ?

Question 237-27 : Increase because it is directly proportional to wing area increase with the square of the wing area remain constant not change because the lift coefficient is constant

Which of these statements about stall speed is correct ?

Question 237-28 : Increasing forward sweep increases stall speed increasing forward sweep decreases stall speed increasing wing anhedral decreases stall speed decreasing wing anhedral decreases stall speed

.the effect of increasing wing sweep angle is to decrease the lift curve slope and clmax as depicted below . /com en/com080 999 jpg.while 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 en/com080 1226 jpg.spanwise 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

The lift formula can be written as . rho = density ?

Question 237-29 : L = cl * 1/2rho * v² * s l = w l = cl * 2rho * v² * s l = n * w

Where .'1/2 rho' is dynamic pressure .'cl' is coefficient of lift .'v' is speed in m/s .'s' is wing area exemple 449 L = cl * 1/2rho * v² * s.L = cl * 1/2rho * v² * s.

During a straight steady climb and with the thrust force parallel to the flight ?

Question 237-30 : Lift is the same as during a descent at the same angle and mass lift is greater than weight lift is equal to weight drag is equal to thrust

.during climb and descent lift is lower than weight .lift is a vertical vector .lift in climb or descent = weight x cos climb angle exemple 453 Lift is the same as during a descent at the same angle and mass.Lift is the same as during a descent at the same angle and mass.

Assuming no compressibility effects induced drag at constant ias is affected by ?

Question 237-31 : Aeroplane mass outside air temperature altitude engine thrust

Given that .pstat = static pressure .rho = density .pdyn = dynamic pressure ?

Question 237-32 : Pstat + 1/2rho * tas² = constant pdyn + pstat = 1/2rho * tas² pstat + 1/2rho * tas² = pdyn ptot = 1/2rho * ias² + pstat

.dynamic pressure is '1/2 rho tas²' .bernoulli's theorem is .pt = ps + pd.then if pd inscreases ps decreases as pt always remains constant exemple 461 Pstat + 1/2rho * tas² = constant.Pstat + 1/2rho * tas² = constant.

The point in the diagram giving the lowest speed in unaccelerated flight is . ?

Question 237-33 : Point 4 point 1 point 2 point 3

Img /com en/com080 1264 jpg. exemple 465 Point 4.Point 4.

The point in the annex showing zero lift is . err a 081 1269 ?

Question 237-34 : Point a point b point c point d

If the airspeed reduces in level flight below the speed for maximum l/d the ?

Question 237-35 : Increase because of increased induced drag reduce because of reduced induced drag increase because of increased parasite drag reduce because of reduced friction drag

.total drag is lowest when parasite drag is equal to the induced drag . /com en/com032 209 jpg.if the airspeed reduces below the speed for maximum lift/drag ratio then induced drag increases thus total drag increases .notice if the airspeed increases above the speed for maximum l/d total drag also increases due to increase of the parasite drag .parasite drag only varies with speed and is directly proportional to v² .induced drag varies with lift speed and aspect ratio is inversely proportional to aspect ratio and v² so multiply by 1/v² exemple 473 Increase because of increased induced drag.Increase because of increased induced drag.

Slat or flap asymmetry occurring after either extension or retraction may have ?

Question 237-36 : Slat asymmetry causes a yawing moment whereas flap asymmetry causes a large rolling moment slat and flap asymmetry both cause a large rolling moment slat asymmetry causes a large rolling moment whereas flap asymmetry causes a large yawing moment slat and flap asymmetry both cause a large yawing 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 rudder .however 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 exemple 477 Slat asymmetry causes a yawing moment, whereas flap asymmetry causes a large rolling moment.Slat asymmetry causes a yawing moment, whereas flap asymmetry causes a large rolling moment.

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

Question 237-37 : Dynamic pressure lift coefficient and wing area square root of wing area density and wing loading total pressure and wing area only angle of attack aspect ratio and dynamic pressure

.lift formula = 1/2 rho cl x v² x s..where .'1/2 rho v²' is dynamic pressure rho = density v = speed in m/s .'cl' is coefficient of lift .'s' is wing area exemple 481 Dynamic pressure, lift coefficient and wing area.Dynamic pressure, lift coefficient and wing area.

Which of these statements about weight or mass is correct ?

Question 237-38 : Weight is a force the weight of an object is independent of the acceleration due to gravity the mass of an object depends on the acceleration due to gravity mass = weight * volume

exemple 485 Weight is a force.Weight is a force.

The induced angle of attack is ?

Question 237-39 : The angle by which the relative airflow is deflected due to downwash the angle between the local flow at the wing and the horizontal tail the angle by which the flow over the wing is deflected when landing flaps are set caused by the fuselage and is greatest at the wing root

.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 exemple 489 The angle by which the relative airflow is deflected due to downwash.The angle by which the relative airflow is deflected due to downwash.

When an aeroplane enters ground effect ?

Question 237-40 : The lift is increased and the drag is decreased the effective angle of attack is decreased the induced angle of attack is increased drag and lift are reduced

exemple 493 The lift is increased and the drag is decreased.The lift is increased and the drag is decreased.

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