Fragen - Principles of Flight

Number: 12989 Question: The speed for minimum glide angle occurs at an angle of attack that corresponds to: (assume zero thrust) A B C D

CLmax. (CL/CD)max. (CL³/CD²)max. (CL/CD²)max.

Number: 4551 Question: The speed for minimum glide angle occurs at an angle of attack that corresponds to: (assume zero thrust; ^ … denotes power of …) A CLmax. B (CL/CD^2)max C (CL^3/CD^2)max D (CL/CD)max.

12989 - B

4551 - D

3. From a polar curve of the entire aeroplane one can read: A B C D

the minimum CL/CD ratio and the minimum drag. the maximum CL/CD ratio and maximum lift coefficient. the minimum drag coefficient and the maximum lift. the minimum drag and the maximum lift.

Number: 12963 Question: Which of the following parameters can be read from the parabolic polar diagram of an aeroplane? A The minimum glide angle and the parasite drag coefficient. B The aspect ratio of the wing and the induced drag coefficient. C The minimum rate of descent and the induced drag. D The induced drag and the parasite drag. The polar diagram's axes are Cl and Cd. Where the line crosses the Cl=0 axis there is no lift, and therefore no induced drag, so Cd at that point is from parasite drag only. Angle of glide is Cl/Cd, the lift/drag ratio. If the lift/drag ratio is 20:1 then the glide angle is 1 in 20. The tangent to the curve from the Cl/Cd origin shows the best Cl/Cd ratio, and thus the best glide angle you can get. Rate of descent would depend on what speed you were doing down the hill and that is not shown on the polar diagram. The polar diagram shows coefficients only. To get real world data you have to add in all the other bits of the lift and drag equations like S and rho and V.

xxx - B

12963 - A

Number: 3415 Question: The point in the figure corresponding to CL for minimum horizontal flight speed is: A B C D

Point c Point b Point a Point d

3415 - D

Number: 3416 Question: Assuming zero thrust, the point on the diagram corresponding to the value for minimum sink rate is: A B C D

Point a Point b Point d Point c

Minimum sink rate occurs at Vmp. Although the drag is more than at Vmd the fact that it is a lot slower means that your rate of descent is low.

Maximum range occurs at Vmd where you get the best CL for the least CD, so that will be at the point of tangency. On the polar diagram the only points that can be read directly are the top of the curve, which is CL max or critical alpha, CD min where the graph is furtherest to the left, and the best L/D ratio which is found at the tangent to the curve from the origin. You can assume two other points with some accuracy. The best EAS/drag speed occurs at a speed faster than Vmd so that must be a smaller angle of attack, or smaller CL so that must be below the point of tangency to the graph. Vmp is always slower than Vmd so must be at a higher angle of attack or CL so it must be between the point of tangency and CL max.

Number: 3417 Question: Assuming zero thrust, the point on the diagram corresponding to the minimum glide angle is: A Point A It is at the best L/D ratio or where Cl/Cd is max, but B Point D that occurs at point (B) the tangent to the curve from C Point C the origin. D Point B

3416 - D

3417 - B

Number: 3418 Question: Which point in the diagram gives the lowest speed in horizontal flight? A Point c B Point b C Point a D Point d

Number: 16574 Question: The point in the diagram giving the lowest speed in unaccelerated flight is: A point 4. B point 1. C point 3. D point 2.

3418 - D

16574 - A

Number: 14963 Question: Given: theta = pitch angle. gamma = flight path angle. alpha = angle of attack. no wind, bank or sideslip. The relationship between these three parameters is: A B C D

alpha = gamma - theta. gamma = alpha - theta. theta = gamma - alpha. theta = gamma + alpha.

14963 - D

Number: 144 Question: The Cl - alpha curve of a positive cambered aerofoil intersects with the vertical axis of the Cl - alpha graph: A in the origin. B below the origin. C nowhere. D above the origin.

Number: 15691 Question: A positively cambered aerofoil will generate zero lift: A at zero angle of attack. B it can never generate zero lift. C at a negative angle of attack. D at a positive angle of attack. Number: 15983 Question: Regarding a positively cambered aerofoil section, which statement is+L7901 correct? I. The angle of attack has a negative value when the lift coefficient equals zero. II. A nose down pitching moment exists when the lift coefficient equals zero. A I is incorrect and II is incorrect. B I is correct and II is incorrect. C I is correct and II is correct. D I is incorrect and II is correct. Number: 14588 Question: The lift coefficient Cl versus angle of attack curve of a symmetrical aerofoil section intersects the vertical axis of the graph: A at the origin. B below the origin. C nowhere. D above the origin. Number: 14589 Question: The lift coefficient Cl versus angle of attack curve of a negatively cambered aerofoil section intersects the vertical axis of the graph: A nowhere. B below the origin. C at the origin. D above the origin.

144 - D

15691 - C

15983 - C

14588 - A

14589 - B

Number: 14594 Question: Regarding a positively cambered aerofoil section, which statement is correct? I. The angle of attack has a positive value when the lift coefficient equals zero. II. A nose up pitching moment exists when the lift coefficient equals zero. A B B C

I is correct and II is incorrect. I is incorrect and II is incorrect. I is correct and II is correct. I is incorrect and II is correct.

I. The angle of attack has a positive value when the lift coefficient equals zero.

II. A nose up pitching moment exists when the lift coefficient equals zero.

Number: 14595 Question: Regarding a symmetric aerofoil section, which statement is correct? I. The angle of attack has a positive value when the lift coefficient equals zero. II. The pitching moment is zero when the lift coefficient equals zero. A B C D

I is incorrect and II is incorrect. I is correct and II is correct. I is correct and II is incorrect. I is incorrect and II is correct.

14594 - B

14595 - D

Number: 14923 Question: An aeroplane in straight and level flight at 300 kt is subjected to a sudden disturbance in speed. Assuming the angle of attack remains constant initially and the load factor reaches a value of approximately 1.2: This question is based on load factors and speed with the A the speed will have decreased by 60 kt. question asking you to determine the change in speed. B the speed will have increased by 30 kt. C the speed will have increased by 60 kt. Note: The aircraft is in steady level flight so the original D the speed will have decreased by 30 kt. load factor will be 1g Furthermore Lift is proportional V^2 and lift is also proportional to load factor It follows that Load factor is proportional to V^2 New LF / Original Load Factor = (new speed / original speed)^2 New speed = 300 x SQRT (1.2 / 1) = 329 kt The speed therefore increases by 30 kt

14923 - B

Number: 16325 Question: An aeroplane flying at 100 kt in straight and level flight is subjected to a disturbance that suddenly increases the speed by 20 kt. Assuming the angle of attack remains constant, the load factor will initially: A increase to 1.21. B increase to 1.44. C increase to 1.10. D remain unchanged, since the angle of attack does not change. Load Factor is proportional to lift which is in turn proportional to V^2 LFnew / LForiginal = (Vnew / Voriginal)^2 LFnew = LForiginal x (120 / 100)^2 = 1 x 1.2^2 = 1.44 The LF will therefore initially increase to 1.44 if the angle of attack remains constant.

16325 - B

Number: xxx Question: Which statement is correct? A The flow on the upper surface of the wing has a component in wing root direction. B Tip vortices can be diminished by vortex generators. C Tip vortices and induced drag decrease with increasing angle of attack. D The flows on the upper and lower surfaces of the wing are both in wing tip direction. Number: 10130 Question: Which of the following wing planforms gives the highest local lift coefficient at the wing root? A Positive angle of sweep B Rectangular C Tapered D Elliptical Number: 2357 Question: High aspect ratio, as compared with low aspect ratio, has the effect of: A increasing lift and critical angle of attack. B increasing lift and drag. C increasing induced drag and decreasing critical angle of attack. D decreasing induced drag and critical angle of attack.

Number: 13215 Question: If the aspect ratio of a wing increases whilst all other relevant factors remain constant, the critical angle of attack will: A increase. B remain constant only for a wing consisting of symmetrical aerofoils. C decrease. D remain constant. Number: 12943 Question: What is the effect on induced drag of an increase in aspect ratio? A Induced drag increases, because a larger aspect ratio increases the frontal area. B Induced drag decreases, because a larger aspect ratio causes more downwash. C Induced drag increases, because the effect of tip vortices increases. D Induced drag decreases, because the effect of tip vortices decreases.

xxxxx - A

10130 - B

2357 - D

13215 - C

12943 - D

Number: 16805 Question: When an aeroplane enters ground effect: A the effective angle of attack is decreased. B the induced angle of attack is increased. C drag and lift are reduced. D the lift is increased and the drag is decreased.

Number: 16838 Question: Assuming constant IAS, when an aeroplane enters ground effect: A the induced angle of attack increases. B the effective angle of attack decreases. C induced drag increases. D the effective angle of attack increases.

Number: 2065 Question: Floating due to ground effect during an approach to land will occur: A when the height is less than halve of the length of the wing span above the surface. B at a speed approaching the stall. C when a higher than normal angle of attack is used. D when the height is less than twice the length of the wing span above the surface.

Number: 2064 Question: What will happen in ground effect ? A a significant increase in thrust required B the wing downwash on the tail surfaces increases C the induced angle of attack and induced drag decreases D an increase in strength of the wing tip vortices

16805 - D

16838 - D

2065 - A

2064 - C

Number: xxx Question: Which statement about an aeroplane entering ground effect is correct? I. The downwash angle increases. II. The induced angle of attack decreases. Number: 14637 A I is incorrect, II is incorrect. Refer to fig 11-3-8. The fig shows three airfoils. The B I is correct, II is correct. top one shows 2 dimensional flow, ie no induced C I is correct, II is incorrect. flow. D I is incorrect, II is correct. The centre airfoil shows how the induced flow now deflects the relative air flow and we get what is called the effective airflow. 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 down wash angle. On entering ground effect induce flow reduces and the induced angle of attack decreases, with an increase in the effective angle of attack. Lift increases and induced drag and down wash reduce.

Number: 14662 Question: Assuming constant IAS, when an aeroplane enters ground effect: A downwash does not change. B induced drag increases. C the effective angle of attack decreases. D downwash reduces. A

downwash does not change.

B

induced drag increases.

C

the effective angle of attack decreases.

D

downwash reduces.

5-D

14662 - D

Number: 14660 Question: Assuming constant IAS, when an aeroplane enters ground effect: A the effective angle of attack does not change. B induced drag increases. C the effective angle of attack decreases. D the induced angle of attack reduces.

A

the effective angle of attack does not change.

B

induced drag increases.

C

the effective angle of attack decreases.

D

the induced angle of attack reduces

14660 - D

Number: xxx Question: An aeroplane accelerates from 80 kt to 160 kt at a load factor equal to 1. The induced drag coefficient (i) and the induced drag (ii) alter with the following factors: A (i) 1/16 (ii) 1/4 Load Factor - The ratio of the weight of an aircraft to the load imposed by lift. B (i) 1/4 (ii) 2 The correct symbol for load factor is (n), but is colloquially known as (g). C (i) 4 (ii) 1/2 D (i) 1/2 (ii) 1/16

Watch the difference between coefficients of drag (or lift) and the actual values for drag and lift. The coefficient of induced drag is proportional to the square of the coeficient of lift. This is because the more lift you need the greater the downwash and all the other factors that generate induced drag. Now write angle of attack in place of the coeficient of lift, and consider level flight. In level flight, from the lift formula, as you speed up alpha goes down, proportionally as the square of increasing IAS. But the “coefficient” of induced drag is proportional to alpha squared, and as alpha is decreasing as the square of IAS the coefficent of induced drag is decreasing as the fourth power of IAS. But the actual induced drag is the coefficient times the square of increasing IAS, and this now makes the actual induced drag decrease as the square of IAS So with the IAS doubling in level flight the coefficient of induced drag, Cdi, is down 1/16, but the actual induced drag is down 1/4.

Number: xxx Question: Increasing dynamic pressure will have the following effect on the total drag of an aeroplane: A total drag increases across the whole speed range. B at speeds below the minimum drag speed, total drag decreases. C at speeds above the minimum drag speed, total drag increases. D total drag decreases across the whole speed range. Number: xxx Question: Increasing dynamic pressure will have the following effect on the drag of an aeroplane (all other factors of importance remaining constant): A at speeds greater than the minimum drag speed, drag increases. B drag decreases across the whole speed range. C drag increases across the whole speed range. D none.

Number: xxx Question: The frontal area of a body, placed in a certain airstream is increased by a factor 3. The shape will not alter. The aerodynamic drag will increase with a factor: A 6. B 9. C 1.5 . D 3. Number: xxx Question: The aerodynamic drag of a body, placed in a certain airstream depends amongst others on: A The cg location of the body. B The weight of the body. C The specific mass of the body. D The airstream velocity.

Number: xxx Question: A body is placed in a certain airstream. The airstream velocity increases by a factor 4. The aerodynamic drag will increase with a factor: A 12 . B 4. C 8. D 16 . Number: xxx Question: A body is placed in a certain airstream. The density of the airstream decreases to half of the original value. The aerodynamic drag will decrease with a factor: A 4. B 8. C 1.4 . D 2.

Number: xxx Question: Assuming no compressibility effects, induced drag at constant IAS is affected by: A aeroplane mass. B engine thrust. C altitude. D outside air temperature.

Number: xxx Question: The induced angle of attack is: A the angle between the local flow at the wing and the horizontal tail. B the angle by which the flow over the wing is deflected when landing flaps are set. C caused by the fuselage and is greatest at the wing root. D the angle by which the relative airflow is deflected due to downwash. The effect of these vortices is to create more down wash behind the wing than up wash in front ofthe w ing. This results in the effective air Flow (EAF) being inclined to the relative airftow (RAF) by an angle called the induced angle of attack (ai).

Number: xxx Question: What is the effect of high aspect ratio of an aeroplane's wing on induced drag? A It is reduced because the effect of wing-tip vortices is reduced. B It is increased because high aspect ratio produces greater downwash. C It is unaffected because there is no relation between aspect ratio and induced drag. D It is increased because high aspect ratio has greater frontal area.

S = WING AREA, sq. m (b x c) b = SPAN, m c = AVERAGE CHORD, m

AR = ASPECT RATIO AR = b/c AR = b²/S

Aspect ratio is the ratio of wing span to wing chord.

Number: xxx Question: If flaps are deployed at constant IAS in straight and level flight, the magnitude of tip vortices will eventually : (flap span less than wing span) A decrease. The clue is maintaining level B increase. C increase or decrease, depending on the initial angle of attack. flight.............if you deploy flap to stay level you will have to REDUCE A of A D remain the same. and decrease induced drag ......................... decreasing the tip vortices.

Number: xxx Question: Excluding constants, the coefficient of induced drag (CDi) is the ratio of : A CL and b (wing span) B CL and CD C CL²and S (wing surface) D CL² and AR (aspect ratio)

Number: xxx Question: Which statement concerning the local flow pattern around a wing is correct?

A

By fitting winglets to the wing tip, the strength of the wing tip vortices is reduced which in turn reduces induced drag. B Sweepback reduces drag since, compared with a straight wing of equal area, the span increases. C Vortex generators on the wing partially block the spanwise flow over the wing leading to a reduction in induced drag. D Slat extension, at a constant angle of attack and nomal extension speeds, will increase the lift coefficient, which will also increase the induced drag coefficient.

Number: xxx Question: An aeroplane transitions from steady straight and level flight into a horizontal co-ordinated turn with a load factor of 2, the speed remains constant and the: A angle of attack increases by a factor of 1/4. B induced drag increases by a factor of 4. To expand a little, total drag has increased, but total drag C total drag increases by a factor of 4. is the sum of profile drag and induced drag, so the D lift increases by a factor of 4. increase in total drag is due to the increase in induced drag not to any increase in profile drag. Profile drag depends on speed squared, and the speed has not changed. Induced drag depends on the square of the lift co-efficient, which has changed. That is why on take-off, when the lift co-efficient is a maximum, induced drag is by far the greater part of total drag

Number: xxx Question: The span-wise flow on an unswept wing is from the: A upper surface via the trailing edge to the lower wing surface. B lower to the upper surface via the wing tip. C upper surface via the leading edge to the lower wing surface. D lower surface via the trailing edge to the upper wing surface.

Number: xxx Question: Which one of the following statements about the lift-to-drag ratio in straight and level flight is correct? A At the highest value of the lift/drag ratio the total drag is lowest. B The lift/drag ratio always increases as the lift decreases. C The highest value of the lift/drag ratio is reached when the lift is equal to the aircraft weight. D The highest value of the lift/drag ratio is reached when the lift is zero. The speed at which total drag is a minimum (V md) occurs when the induced and parasite drag are equal

17. At a load factor of 1 and the aeroplane's minimum drag speed, what is the ratio between induced drag Di and parasite drag Dp?

A B C D

Di/Dp = 1. Di/Dp = 1/2. Di/Dp = 2. It varies between aeroplane types.

18. The value of the parasite drag in straight and level flight at constant weight varies linearly with the:

A B C D

square of the angle of attack. speed. square of the speed. angle of attack.

The speed at which total drag is a minimum (V md) occurs when the induced and parasite drag are equal

19. How does the total drag vary as speed is increased from stalling speed (VS) to maximum IAS (VNE) in a straight and level flight at constant weight?

A B C D

Increasing, then decreasing. Increasing. Decreasing. Decreasing, then increasing.

20. Which one of the bodies in motion (all bodies have the same cross section area) will have lowest drag?

A B C D

Body b Body c Body d Body a

21. Which line represents the total drag line of an aeroplane?

A B C D

Line b Line d Line c Line a

22. What is the effect on induced drag of mass and speed changes? (all other factors of importance remaining constant)

A B C D

Decreases with increasing speed and decreasing mass. Increases with increasing speed and increasing mass. Decreases with decreasing speed and decreasing mass. Increases with increasing speed and decreasing mass.

23. The diagram shows the parameter X versus TAS. If a horizontal flight is considered the axis X shows

A B C D

the total drag. the induced drag. the lift force. the parasite drag.

RELATION LIFT - SPEED 1. An aeroplane maintains straight and level flight while the IAS is doubled. The change in lift coefficient will be:

x 4.0 x 0.5 x 0.25 x 2.0

Remember, if an aircraft is in level flight then Lift = Weight. In the scenario given this is the case. So if we just put some figures in as an example. Just for ease, shall we say that S = 2, Rho = 10, Cl = 1 and V = 100. So to start with we have: Lift = 1 x 0.5 x 10 x (100 x 100) x 2 = 100000 Now, if we double the speed to 200 and leave everything else the same we have: Lift = 1 x 0.5 x 10 x (200 x 200) x2 = 400000 But we know that the aircraft is maintaining level flight and therefore Lift cannot change. Hence we need to modify something on the right hand side of the equation so that lift remains at the value of 100000. We can't modify S (only engineers do that) and Rho remains the same as there is nothing to say that the air density has changed. Therefore the only factor we can change is Cl. If we multiply Cl on the right by 0.25 we can ensure that Lift remains at 100000. You MUST understand that in straight and level flight, lift does not change as it is only balancing against weight (which for this example you can assume to be constant). So if density remains the same, if IAS doubles TAS doubles too.

2. Whilst maintaining straight and level flight with a lift coefficient CL = 1 what will be the new approximate value of CL after the speed is increased by 30%?

A B C D

0.50. 0.30. 0.25. 0.60.

3. In straight and level flight at a speed of 1.3 VS, the lift coefficient, expressed as a percentage of its maximum (CLmax), would be:

A B C D

59%. 77%. 130%. 169%.

If you visualise the CL curve then CLmax (or Vs) will be at critical alpha or the very peak of the graph. 1.3Vs is faster than Vs so it must be at a smaller CL. To determine the percentage you need to rearrange the lift formula. L = CL x Vsquared (all other factors remaining constant) So CL = L/Vsquared We dont know the lift but at Vs we can assume unity so CL = 1/(1.3)squared and then x100 to get a percentage =59%

4. Whilst maintaining straight and level flight with a lift coefficient CL=1, what will be the new value of CL after the speed has doubled?

A B C D

1.00. 0.25. 2.00. 0.50.

5. When an aeroplane is flying at an airspeed which is 1.3 times its basic stalling speed, the coefficient of lift as a percentage of the maximum lift coefficient (CLmax) would be:

A B C D

77%. 59%. 130%. 169%.

6. Whilst maintaining straight and level flight with a lift coefficient CL = 1, what will be the new approximate value of CL after the speed is increased by 41%?

A B C D

0.30. 0.60. 0.25. 0.50.

1. Given an initial condition in straight and level flight with a speed of 1.4 VS. The maximum bank angle attainable without stalling in a steady co-ordinated turn, whilst maintaining speed and altitude, is approximately:

A B C D

30°. 60°. 32°. 44°.

2. An aeroplane has a stall speed of 78 KCAS at its gross weight of 6850 Ibs. What is the stall speed when the weight is 5000 Ibs?

A B C D

91 KCAS. 78 KCAS. 57 KCAS. 67 KCAS.

The speed at which total drag is a minimum (V md) occurs when the induced and parasite drag are equal

3. An aeroplane has a stall speed of 100 kt at a mass of 1000 kg. If the mass is increased to 2000 kg, the new value of the stall speed will be:

A B C D

150 kt. 123 kt. 200 kt. 141 kt.

4. The stall speed in a 60° banked turn increases by the following factor:

A B C D

1.30. 1.07. 2.00. 1.41.

5. Two identical aircraft A and B, with the same mass, are flying steady level co-ordinated 20 degree bank turns. If the TAS of A is 130 kt and that of B is 200 kt:

A B C D

the rate of turn of A is greater than that of B. the lift coefficient of A is less than that of B. the load factor of A is greater than that of B. the turn radius of A is greater than that of B.

6. By what approximate percentage will the stall speed increase in a horizontal co-ordinated turn with a bank angle of 45°?

A B C D

41%. 31%. 52%. 19%.

7. An aeroplane has a stall speed of 100 kt. When the aeroplane is flying a level co-ordinated turn with a load factor of 1.5, the stall speed is:

A B C

141 kt. 150 kt. 82 kt.

Number: 3878 Question: An aeroplane has a stall speed of 100 kt at a load factor n=1. In a turn with a load factor of n= 2, the stall speed is:

A B C D

141 kt. 200 kt. 282 kt. 70 kt.

3878 - A

9. If the stall speed of an aeroplane is 60 kt, at what speed will the aeroplane stall if the load factor is 2?

A B C D

72 kt. 85 kt. 66 kt. 120 kt.

10. An increase in wing loading will:

A B C D

increase sensitivity to turbulence. increase the stall speed. decrease the minimum glide angle. increase CLmax.

11. The stall speed:

A B C D

does not depend on weight. decreases with an increased weight. increases with the length of the wingspan. increases with an increased weight.

12. Stall speed (IAS) varies with:

A B C D

altitude. (below approximately 10000 ft) density. weight. temperature.

11. Which of the following statements about the stall of a straight wing aeroplane is correct?

A B C D

The horizontal tail will stall at a higher speed than the wing. Just before the stall the aeroplane will be have an increased nose down tendency. Buffeting is the result of tailplane flow separation.. The nose down effect is the result of increasing downwash, due to flow separation.

12. Entering the stall the centre of pressure - of a straight (1) wing and - of a strongly swept back wing (2) will:

A B C D

(1) move aft, (2) move aft. (1) move aft, (2) not move. (1) move aft, (2) move forward. (1) not move (2) move forward.

13. Which of these statements about the effect of wing sweep on centre of pressure location are correct or incorrect? I. The centre of pressure on a straight wing moves aft as the angle of attack approaches and exceeds the critical angle of attack. II. The centre of pressure on a strongly swept back wing moves forward as the angle of attack approaches and exceeds the critical angle of attack.

A B C D

I is correct, II is incorrect. I is incorrect, II is correct. I is incorrect, II is incorrect. I is correct, II is correct.

14. A jet aeroplane cruises buffet free at constant high altitude. Which type of stall is most likely to occur if this aeroplane decelerates during an inadvertant increase in load factor?

A B C D

Accelerated stall. Deep stall. Shock stall. Low speed stall.

15. Two identical aircraft A and B, with the same mass, are flying steady level co-ordinated 20 degree bank turns. If the TAS of A is 130 kt and that of B is 200 kt:

A B C D

the turn radius of A is greater than that of B. the lift coefficient of A is less than that of B. the load factor of A is greater than that of B. the rate of turn of A is greater than that of B.

16. When a strongly swept-back wing stalls and the wake of the wing contacts the horizontal tail, the effect on the stall behaviour can be a(n):

A B C D

nose down tendency. increase sensitivity of elevator inputs. tendency to increase speed after initial stall. nose up tendency and/or lack of elevator response.

17. Dangerous stall characteristics, in large transport aeroplanes that require stick pushers to be installed, include:

A B C D

pitch down and increase in speed. distinct aerodynamic buffet. excessive wing drop and deep stall. pitch down and yaw.

18. During a climbing turn to the right the:

A B C D

angle of attack of the left wing is smaller than the angle of attack of the right wing. stall angle of attack of the left wing will be larger than the corresponding angle for the right wing. angle of attack of both wings is the same. angle of attack of the left wing is larger than the angle of attack of the right wing.

Climbing Turns During climbing turns, an aircraft describes an upward spiral path. Thus, the relative airflow comes downward to meet the wings, thereby reducing their angles of attack and hence their coefficients of lift. The faster moving outer wing is subject to a smaller reduction in angle of attack. This makes the net coefficient of lift higher than on the inner wing, thereby producing greater lift. Its increased velocity further enhances the lifting capability of the outer wing. An aircraft in a climbing turn therefore tends to overbank more than in a steady level turn. If necessary, utilise the ailerons to maintain the desired angle of bank.

19. Which of these statements about boundary layers is correct?

A B C D

Compared with a laminar boundary layer, a turbulent boundary layer is better able to resist a positive pressure gradient before it separates. A laminar boundary layer is thicker than a turbulent one. A turbulent boundary layer becomes laminar at the transition point. A turbulent boundary layer produces less friction drag than a laminar one.

20. Which statement is correct about the laminar and turbulent boundary layer :

A B C D

friction drag will be equal in both types of layers separation point will occur earlier in the turbulent layer friction drag is lower in the laminar layer friction drag is lower in the turbulent layer

The boundary layer is the layer of air between the surface and the free stream velocity in which local retardation takes place. Like the main airflow, the boundary layer flow can be either laminar or turbulent in nature. The laminar boundary layer is a very thin layer of smooth airflow. It consists of a series of laminations or smooth regular streamlines, in which the air particles do not intermingle. The turbulent boundary layer is a layer of disturbed or turbulent airflow, in which the streamlines break up. The air particles become intermingled and move in a random, irregular pattern. Notably, the turbulent boundary layer creates greater drag than the laminar boundary layer.

21. Behind the transition point in a boundary layer:

A B C D

the boundary layer gets thinner and the speed increases. the boundary layer gets thicker and the speed decreases. the mean speed increases and the friction drag decreases. the mean speed and friction drag increases.

Number: 987 Question: Which of the following statements about the difference between Krueger flaps and slats is correct?

A B C D

Deploying a Krueger flap will increase critical angle of attack, deploying a slat does not. Deploying a slat will increase critical angle of attack, deploying a Krueger flap does not. Deploying a slat will form a slot, deploying a Krueger flap does not. Deploying a Krueger flap will form a slot, deploying a slat does not.

987 - C

2. After take-off the slats (when installed) are always retracted later than the flaps. Why?

A Because VMCA with SLATS EXTENDED is more favourable compared with the FLAPS EXTENDED situation. B Because SLATS EXTENDED gives a large decrease in stall speed with relatively less drag. C Because FLAPS EXTENDED gives a large decrease in stall speed with relatively less drag.

3. The high lift device shown in the figure is a

A B C D

Slotted flap Krueger flap Slat Fowler flap

4. The high lift device shown in the figure below is a

A B C D

Slotted flap Fowler flap Slot or slat Krueger flap

2-B

3-C

4-D

5. What increases the stalling angle of attack? Use of:

A B C D

spoilers. flaps. fuselage mounted speed-brakes. slats.

6. Which statement is correct?

A Flap extension reduces the stallspeed, which increases the maximum glide distance. B Flap extension has no effect on the minimum rate of descent as this is only affected by TAS. C Spoiler extension increases the stallspeed, the minimum rate of descent and the minimum angle of descent. D Flap extension reduces the maximum lift/drag ratio thus reducing the minimum rate of descent. 7. Given the following aeroplane configurations: 1. Clean wing. 2. Slats only extended. 3. Flaps only extended. Place these configurations in order of increasing critical angle of attack:

A B C D

1, 3, 2. 2, 3, 1. 2, 1, 3. 3, 1, 2.

8. A slat will

A B C D

increase the lift by increasing the wing area and the camber of the aft portion of the wing. prolongs the stall to a higher angle of attack. provide a boundary layer suction on the upper surface of the wing. increase the camber of the aerofoil and divert the flow around the sharp leading edge.

9. A plain flap will increase CLmax by

A B C D

increasing the camber of the aerofoil. centre of lift movement. increasing angle of attack. boundary layer control.

10. For most jet transport aeroplanes, slat extension has:

A B C D

the same minor effect on stall speed as flap extension. a minor effect on stall speed whereas flap extension has a significant effect. a greater effect on stall speed than flap extension. the same significant effect on stall speed as flap extension.

1. Ignoring downwash effects on the tailplane, extension of Fowler flaps, will produce:

A B C D

no pitching moment. a nose-up pitching moment. a nose-down pitching moment. a force which reduces drag.

2. On a wing fitted with a "fowler" type trailing edge flap, the "Full extended" position will produce:

A B C D

an increase in wing area only. an increase in wing area and camber. an unaffected wing area and increase in camber. an unaffected CD, at a given angle of attack.

3. When flaps are extended whilst maintaining straight and level flight at constant IAS, the lift coefficient will eventually:

A B C D

increase. remain the same. first increase and then decrease. decrease.

Remember CL can be changed either by changing angle of attack or by changing the camber. (a couple of others as well, but as a pilot you can't directly control them) As the question specifies straight and level flight and at a constant IAS then you have to generate the same lift, so CL must stay constant. Lowering the flaps will increase the camber (and in some aircraft the wing area as well) so the CL WILL increase. You have to keep the CL constant, so you will have to reduce the angle of attack to achieve that. The Cl increases when you extend flaps, BUT the question states that you are maintaining straight and level flight.. what do you have to do when you extend the flaps? You have to lower the nose to maintain level flight.. thus Cl remains the same..

4. When flaps are deployed at constant angle of attack the lift coefficient will:

A B C D

vary as the square of IAS. remain the same. decrease. increase.

5. Trailing edge flap extension will:

A B C D

decrease the critical angle of attack and decrease the value of CLmax. increase the critical angle of attack and increase the value of CLmax. increase the critical angle of attack and decrease the value of CLmax. decrease the critical angle of attack and increase the value of CLmax.

6. What is the most effective flap system?

A B C D

Fowler flap. Split flap. Single slotted flap. Plain flap.

7. Deploying a Fowler flap, the flap will:

A B C D

just move aft. move aft, then turn down. just turn down. turn down, then move aft.

8. A slotted flap will increase the CLmax by:

A B C D

decreasing the skin friction. increasing the critical angle of attack. increasing only the camber of the aerofoil. increasing the camber of the aerofoil and re-energising the airflow. The slotted flap is similar to the plain flap except that when deflected, a slot forms between the flap and main wing. This allows high pressure air below the wing to flow through the slot and re-energise the boundary layer over the upper surface of the flap. The combination of variable geometry and boundary layer control thus increases the wing's lift performance beyond that of the plain flap at all angles of attack.

9. In order to maintain straight and level flight at a constant airspeed, whilst the flaps are being retracted, the angle of attack must be:

A B C D

increased. increased or decreased depending on type of flap. decreased. held constant.

10. An aeroplane has the following flap settings: 0°, 15°, 30° and 45°. Slats can also be selected. Which of the following selections will most adversely affect the CL/CD ratio?

A B C D

Flaps from 15° to 30°. Flaps from 30° to 45°. The slats. Flaps from 0° to 15°.

11. An aeroplane has the following flap positions: 0°, 15°, 30°, 45°. Slats can also be selected. Generally speaking, which selection provides the highest positive contribution to the CLMAX?

The flaps from 15° to 30°. The slats from the retracted to the take-off position. The flaps from 30° to 45°. The flaps from 0° to 15°. 12. Compared with the clean configuration, the angle of attack at CLmax with trailing edge flaps extended is:

A B C D

smaller. unchanged. smaller or larger depending on the degree of flap extension. larger.

13. Flap extension at constant IAS whilst maintaining straight and level flight will increase the:

A B C D

Slightly tricky. From 1/2CL S rho Vsquared, if you maintain IAS in

lift and the drag. level flight and S does not change then CL doesn't change either. lift coefficient and the drag. Fowler flaps increase S, but this does not increase CL for level flight. Strictly, with Fowler flaps CL required for level flight goes stall speed. maximum lift coefficient (CLmax) and the drag. down.

CL max available increases with all types of flap - after all, flap is called a "high lift device". So will drag. The pilot adjusts attitude and power to maintain speed and S&L

14. Which type of flap is shown in the picture?

A B C D

Plain flap Fowler flap Double slotted flap Split flap

15 Which type of flap is shown in the picture?

A B C D

Single slotted flap Fowler flap Plain flap Split flap

16. From an initial condition of level flight the flaps are extended at a constant pitch attitude. The aeroplane will subsequently:

A B C D

maintain level flight. start to bank. start to sink. start to climb.

17. From an initial condition of level flight the flaps are retracted at a constant pitch attitude. The aeroplane will subsequently:

A B C D

start to climb. start to bank. maintain level flight. start to sink.

1 Question: When an aeroplane with the centre of gravity forward of the centre of pressure of the combined wing / fuselage is in straight and level flight, the vertical load on the tailplane will be: A B C D

downwards because it is always negative regardless of the position of the centre of gravity. downwards. zero because in steady flight all loads are in equilibrium. upwards.

2 Question: The pitching moment versus angle of attack line in the diagram, which corresponds to a CG located at the neutral point of of a given aeroplane at low and moderate angles of attack is: A B C D

line 4. line 1. line 3. line 2.

A couple of basics you must first grasp... First off, what is the Neutral Point? The Neutral Point is the position of the CG that gives the aircraft neutral longitudinal static stability. Now do you know about CM/alpha diagrams? If not, you need to make sure you go back to your notes and get to grip with this... Now the answer to this particular question is in Line 2, the horizontal line that shows neutral longitudinal static stability.

3 Question: Which line in the graphic of Cm versus angle of attack graph shows a statically stable aeroplane? A B C D

Line 2. Line 1. Line 4. Line 3.

4 Question: Which line in the diagram illustrates an aeroplane which is statically longitudinally stable at all angles of attack? A B C D

Line 3. Line 2. Line 4. Line 1.

Number: 5 Question: Where on the curve in the diagram does the aeroplane exhibit neutral static longitudinal stability?

A B C D

to exhibit - aufweisen, zeigen

The whole curve. Part 3. Part 1. Point 2.

Number: 16575 Question: Where on the curve in the diagram does the aeroplane exhibit static longitudinal stability? A The whole curve. B Part 3. C Point 2. D Part 1.

4-C

5-D

16575 - D

7 Question: The CG of an aeroplane is in a fixed position forward of the neutral point. Which of these statements about the stick force stability is correct? A

Stick force stability is not affected by trim.

B

An increase of 10kt from the trimmed position at low speed has more effect on the stick force than an increase of 10kt from the trimmed position at high speed.

C

Aeroplane nose up trim decreases the stick force stability.

D

Maintaining a steady speed above the trim speed requires a pull force.

Answer” Increase of speed generates pull forces.” Is incorrect. If the CG is ahead of the neutral point we will have a stable configuration, so an increase in speed would require a push Answer “Aeroplane nose up trim decreases the stick force stability.” Is incorrect. If you trim nose up, you would have to keep a forward pressure on the stick to maintain straight and level. If you increase speed, the forward pressure would increase as the elevator become more effective, and that is in the same sense you would expect with a speed increase, so stick force stability increases. Answer “Stick force stability is not affected by trim. “ Is incorrect. If you trimmed nose down you would require a pull force to keep the nose up, and fly straight and level. As you increased speed the pull force would increase, which is not what you would expect, so that is an unstable stick force. So trim does affect stick force stability.

You can calculate the difference of a 10 kt change at low speed and high speed to prove the answer. It goes about the percentage change in lift. If you assume a CL of 0.3 in straight and level and substitute changes of speed maintaining the same angle of attack, then you will see that the percentage change in lift, and therefore stick forces, will be greater at lower speeds. L = CL x V2 (no change in the ½ ρ or S so we can leave them out, in this calculation) (I also left it in Kts, as it will not affect this specific calculation, but we should use M/sec) Change in speed: Squared: X

60 3600 0.3

Change: Percentage:

750 17%

390 36%

70 4900 1080

120 14400 1470

130 16900 4320

5070

8 Question: The value of the manoeuvre stability of an aeroplane is 150 N/g. The stick force required to achieve a load factor of 2,5 from steady level flight is: A B C D

375 N. 225 N. 150 N. 450 N.

Adding 1.5G to get from 1G to 2.5G, at a rate of 150N/G, answer 1.5 x 150 = 225N

9 Question: The stick force per g of a heavy transport aeroplane is 300 N/g. What stick force is required, if the aeroplane in the clean configuration is pulled to the limit manoeuvring load factor from a trimmed horizontal straight and steady flight? A B C D

450 N. 1125 N. 825 N. 750 N.

1. When a jet transport aeroplane takes off with the CG at the forward limit and the trimmable horizontal stabiliser (THS) is positioned at the maximum allowable nose down position for take-off:

A B C D

there will be a tendency to over-rotate. rotation will be normal using the normal rotation technique. rotation will require a higher than normal stick force. early nose wheel raising will take place.

2. When a jet transport aeroplane takes off with the CG at the aft limit and the trimmable horizontal stabiliser (THS) is positioned at the maximum allowable nose up position for take-off:

A B C D

there will be a tendency to under-rotate. rotation will require higher than normal stick force. rotation will be normal using the normal rotation technique. early nose wheel raising will take place.

3. When a jet transport aeroplane takes off with the CG at the forward limit and the trimmable horizontal stabiliser (THS) is positioned at the maximum allowable nose up position for take-off:

A B C D

early nose wheel raising will take place. rotation will be normal using the normal rotation technique. rotation will require a higher than normal stick force. there will be a tendency to over-rotate.

4. Comparing the differences between a horizontal trimmable stabiliser and an elevator trim tab, which of these statements are correct or incorrect? I. The effects of a trim tab runaway are more serious. II. A jammed trim tab causes less control difficulty.

A B C D

I is correct, II is incorrect. I is correct, II is correct. I is incorrect, II is incorrect. I is incorrect, II is correct.

5. Which of these statements about a trimmable horizontal stabiliser is correct?

A

Because take-off speeds do not vary with CG position, the need for stabiliser adjustment is dependent on flap position only.

B

At the aft CG limit, stabiliser trim is adjusted fully nose up to obtain maximum elevator authority at rotation during take-off.

C

A trimmed aeroplane with an aft CG requires the stabiliser leading edge to be lower than in the case of a forward CG in the same condition.

D

A trimmed aeroplane with an aft CG requires the stabiliser leading edge to be higher than in the case of a forward CG in the same condition.

6. When comparing an elevator trim system with a stabiliser trim system, which of these statements is correct?

A B C D

an elevator trim produces lower trim drag an elevator trim is more suitable for aeroplanes with a large CG range. an elevator trim is able to compensate larger changes in pitching moments. an elevator trim is more sensitive to flutter.

2. Which three aerodynamic means decrease manoeuvring stick forces?

A B C D

Spring tab - trim tab - mass balancing weight. Servo tab - horn balance - spring tab. Spring tab - horn balance - bobweight. Servo tab - trim tab - balance tab.

3. The tab in the figure represents:

A B C D

a balance tab . an anti-balance tab. a servo tab. a trim tab.

4. The tab in the figure represents:

A B C D

a control tab. an antibalance tab. a balance tab that also functions as a trim tab. a trim tab.

jams - klemmt 5. An aeroplane has a servo tab controlled elevator. What will happen if the elevator jams during flight?

A B C D

Only the tab is left functioning. It is not connected to the The pitch control forces double. elevator but works in the unnatural sense to drive the Pitch control is lost. elevator in the natural sense, so with the elevator locked Pitch control sense is reversed. you only have very limited control from the tab only and in The servo-tab now works as a negative trim-tab. the unnatural sense - and, of course, you have no trim or balance function left.

6. Which statement about a primary control surface controlled by a servo tab, is correct?

A B C D

Due to the effectiveness of the servo tab the control surface area can be smaller. The servo tab can also be used as a balance tab. The position is undetermined during taxiing, in particular with tailwind. The control effectiveness of the primary surface is increased by servo tab deflection. The function of a servo tab is very different from a balance tab. With a servo tab control system movement of the pilot’s flight controls moves the servo tab. The servo tab at the trailing edge of the main flying control surface produces a aerodynamic force to move the control surface. The servo tab is displaced in the opposite direction in which the flight control surface moves. ie if you wish to pitch the aircraft nose up, servo tab is deflected down and moves the elevator up. The system requires airflow from leading edge to trailing edge, when taxiing in a tailwind the effectiveness of this type of control is reduced.

7. What is the fundamental difference between a trim tab and a servo tab?

A The purpose of a trim tab is to reduce continuous stick force to zero, a servo tab only reduces stick force. B The functioning of a trim tab is based on aerodynamic balancing, whereas a servo tab is usually adjusted via a screwjack. C A trim tab is automatically adjusted when its particular control surface moves, whereas a servo tab is moved The trim tab is used to trim the aircraft, which is reducing stick force required to hold an attitude to zero. The spring tab is a system used to reduce control forces over the whole operating range

2. When the cg position is moved forward, the elevator deflection for a manoeuvre with a given load factor greater than 1 will be:

A B C D

smaller. unchanged. dependent on trim position. larger.

3. What is the effect of an aft shift of the centre of gravity on (1) static longitudinal stability and (2) the required control deflection for a given pitch change?

A B C D

(1) reduces (2) increases. (1) increases (2) increases. (1) increases (2) reduces. (1) reduces (2) reduces.

4. Which statement in respect of a trimmable horizontal stabiliser is correct?

A

An aeroplane with a forward cg requires the stabiliser leading edge to be higher than for one with an aft cg in the same trimmed condition.

B

Because take-off speeds do not vary with centre of gravity location, the need for stabiliser adjustment is dependent on flap position only.

C at

At the forward C.G. limit , stabiliser trim is adjusted fully nose down to obtain maximum elevator authority

D

rotation during take-off. An aeroplane with a forward cg requires the stabiliser leading edge to be lower than for one with an aft cg If the C.G is moved aft it results in a larger nose down pitching moment which has to be compensated for by placing a download on the tail plane. This can be achieved by deflecting the elevator up, although this results in high trim drag at higher airspeeds. On large aircraft small movements of the variable incidence stabilizer allow the trimming to be carried out, but with reduced trim drag. An upward deflection of the elevator is required to trim the aircraft, so think of the stabilizer as a large elevator with its trailing edge moving in the same direction as the elevator. To achieve this you must therefore lower the leading edge of the stabilizer to produce the necessary download. Note: Once in a trimmed condition on fully powered flying controls the elevator will then align itself with the stabilizer to reduce the tail load and also the loads acting on the hinges and servo actuator. On power assisted controls there is however no guarantee that this occurs due to downwash and configuration effects that can alter the flow over the surface. It is also fair to say that trim drag also

5. An example of differential aileron deflection during initiation of left turn is:

A B C D

Left aileron: 2° down. Right aileron: 5° up. Left aileron: 5° down. Right aileron: 2° up. Left aileron: 5° up. Right aileron: 2° down. Left aileron: 2° up. Right aileron: 5° down.

6. When a turn is initiated, adverse yaw is:

A

a momentary yawing motion opposite to the turn due to an incorrect differential aileron movement.

B

the tendency of an aeroplane to yaw in the opposite direction of turn mainly due to the difference in aileron form drag.

C

the tendency of an aeroplane to yaw in the same direction of turn due to the different wing speeds.

D

the tendency of an aeroplane to yaw in the opposite direction of turn mainly due to the difference in induced drag on each wing.

7. In what phase of flight are the outboard ailerons (if fitted) not active?

A B C D

Cruise. Landing with a strong and gusty crosswind, to avoid over-controlling the aeroplane. Take-off, until lift-off. Approach.

8. When roll spoilers are extended, the part of the wing on which they are mounted:

A

experiences a reduction in lift, which generates the desired rolling moment. In addition there is a local increase in drag, which suppresses adverse yaw.

B

is forced downwards as a reaction to the increased drag.

C

experiences extra drag, which generates a yawing moment. The speed difference between both wings generates the desired rolling moment.

D

stalls. This causes a difference in lift between both wings, which generates the desired rolling moment.

9. When are outboard ailerons (if present) de-activated?

A B C D

Flaps (and slats) retracted or speed above a certain value. Landing gear extended. Flaps (and/or slats) extended or speed below a certain value. Landing gear retracted.

10. Yaw is followed by roll because the:

A

rudder is located above the longitudinal axis and when it is deflected, it causes a rolling moment in the same direction as the yaw.

B

rolling motion generated by rudder deflection causes a speed increase of the outer wing which increases the lift on that wing so the aeroplane starts to roll in the direction of the turn.

C

yawing motion generated by rudder deflection causes a speed increase of the inner wing, which increase the lift on that wing so that the aeroplane starts to roll in the same direction as the yaw.

D

yawing motion generated by rudder deflection causes a speed increase of the outer wing, which increases the lift on that wing so that the aeroplane starts to roll in the same direction as the yaw. Yawing motion generated by: -> rudder deflection -> speed increase of the outer wing, -> increase the lift on that wing -> aeroplane starts to roll in the same direction as the yaw.

11 If the nose of an aeroplane yaws left, this causes:

A B C D

a roll to the left. an increase in lift on the left wing. a roll to the right. a decrease in relative airspeed on the right wing.

1. For this question use the reference. The sequence which correctly represents blade twist at the given sections is:

A B C D

Sequence 4 Sequence 1 Sequence 3 Sequence 2 This graphic is a propeller blade. Blades are twisted along their length so the angle they meet the air is the same from root to tip. This makes sequence 4 correct.

2. For this question use the reference. The diagram that correctly represents the propeller in the feathered position is:

A B C D

Diagram 4 Diagram 1 Diagram 3 Diagram 2

3. Which statement is correct? I. A propeller with little blade twist is referred to as being in fine pitch. II. A propeller with significant blade twist is referred to as being in coarse pitch.

A B C D

I is incorrect, II is incorrect. I is correct, II is incorrect. I is correct, II is correct I is incorrect, II is correct.

4. Which statement is correct? I. A propeller with little blade twist is referred to as being in fine pitch. II. A propeller with a large blade angle is referred to as being in coarse pitch.

A B C D

I is incorrect, II is incorrect I is correct, II is incorrect. I is correct, II is correct. I is incorrect, II is correct.

to twist - verdrehen

2. The variation of propeller efficiency of a fixed pitch propeller with TAS at a given RPM is shown in:

A B C D

figure 4. figure 2. figure 3. figure 1.

3. For this question use the reference. The diagram that correctly represents the aerodynamic forces acting on a propeller in normal flight is:

A B C D

Diagram 3 Diagram 4 Diagram 2 Diagram 1

thrust is the component of the total aerodynamic force on the propeller parallel to the rotational axis.

4. Use graphic at reference. A rotating propeller blade element produces an aerodynamic force F that may be resolved into two components: - a force T perpendicular to the plane of rotation (thrust). - a force R generating a torque absorbed by engine power. The diagram representing a windmilling propeller is:

A B C D

diagram 3. diagram 2. diagram 1. diagram 4.

5. A windmilling propeller:

A B C D

improves the glide performance of an aeroplane. has a greater blade angle than a feathered propeller. produces neither thrust nor drag. produces drag instead of thrust.

6. For this question use the reference. A rotating propeller blade element produces an aerodynamic force F that may be resolved into two components: - a force T perpendicular to the plane of rotation (thrust). - a force R generating a torque absorbed by engine power. The diagram representing a rotating propeller blade element during reverse operation is:

A B C D

Diagram 4 Diagram 3 Diagram 2 Diagram 1

Reverse Pitch An aerodynamic brake position used for braking and sometimes ground manoeuvring. It is achieved by accelerating air forward by the blade going into a negative angle. Feathered When the chord line of the blade is parallel to the airflow, therefore preventing wind milling. Coarse Pitch The maximum cruising pitch in normal operation. Flight Fine Pitch The minimum pitch obtainable in flight. Ground Fine Pitch The minimum torque position for ground operation and is sometimes referred to as superfine pitch.

Number: 15800 Question: If the propeller pitch of a windmilling propeller is decreased during a glide at constant IAS the propeller drag in the direction of flight will:

A B C D

increase and the rate of descent will increase. decrease and the rate of descent will increase. decrease and the rate of descent will decrease. increase and the rate of descent will decrease.

Number: 16318 Question: If the propeller pitch of a windmilling propeller is increased during a glide at constant IAS the propeller drag in the direction of flight will:

A B C D

decrease and the rate of descent will increase. decrease and the rate of descent will decrease. increase and the rate of descent will decrease. increase and the rate of descent will increase.

Number: 16319 Question: During a glide with idle power and constant IAS, if the RPM lever of a constant speed propeller is pulled back from its normal cruise position, the propeller pitch will:

A B C D

decrease and the rate of descent will increase. increase and the rate of descent will decrease. decrease and the rate of descent will decrease. increase and the rate of descent will increase.

In this question the aircraft is in the cruise with a coarse pitch. If the rpm lever is pulled rearwards the pitch will increase further. This will reduce the drag and thus the L/D ratio, reducing the glide angle and decreasing the rate of descent. This is what the question asks, but it is useful to go a stage further and identify the effect on the A of A. As the aircraft is descending at a constant IAS in the glide the TAS will decrease and the A of A will increase.

15800 - A

16318 - B

16319 - B

1. If the propeller pitch of a windmilling propeller is increased during a glide at constant IAS the propeller drag in the direction of flight will:

A B C D

increase and the rate of descent will decrease. increase and the rate of descent will increase. decrease and the rate of descent will increase. decrease and the rate of descent will decrease.

2. If the propeller pitch of a windmilling propeller is decreased during a glide at constant IAS the propeller drag in the direction of flight will:

A B C D

decrease and the rate of descent will increase. decrease and the rate of descent will decrease. increase and the rate of descent will increase. increase and the rate of descent will decrease.

3. During a glide with idle power and constant IAS, if the RPM lever of a constant speed propeller is pulled back from its normal cruise position, the propeller pitch will:

A B C D

decrease and the rate of descent will decrease. decrease and the rate of descent will increase. increase and the rate of descent will decrease. increase and the rate of descent will increase.

In this question the aircraft is in the cruise with a coarse pitch. If the rpm lever is pulled rearwards the pitch will increase further. This will reduce the drag and thus the L/D ratio, reducing the glide angle and decreasing the rate of descent. As the aircraft is descending at a constant IAS in the glide the TAS will decrease and the A of A will increase.

Coarse Pitch = The maximum cruising pitch in normal operation.

4. Assuming that the RPM remains constant throughout, the angle of attack of a fixed pitch propeller will:

A B C D

remain constant at a fixed value only if the airspeed decreases. remain constant at a fixed value irrespective of any airspeed changes. increase with increasing airspeed. decrease with increasing airspeed.

5. The difference between a propeller's blade angle and its angle of attack is called:

A B C D

the effective pitch. propeller slip. the propeller angle. the helix angle.

Blade Angle = Angle of Attack + Helix Angle Helix Angle

= Blade Angle

-

Angle of Attack

6. If the RPM lever of a constant speed propeller is moved forward during a glide with idle power and whilst maintaining constant airspeed, the propeller pitch will:

A B C D

decrease and the rate of descent will decrease. increase and the rate of descent will increase. decrease and the rate of descent will increase. increase and the rate of descent will decrease. The RPM lever is the prop lever. The throttle only controls the butterfly in the carb. So with the throttle at idle, if you push the RPM lever (prop lever) forward you are forcing the blades to go to the full fine pitch position - decreasing the blade angle (pitch) At idle that presents a big flat face to the air so a lot of drag, so rate of descent will increase. With the throttle open then it would have minimum drag therefore maximum RPM.

7. If you decrease the propeller pitch during a glide with idle-power at constant IAS the lift to drag ratio will

A B C D

decrease and the rate of descent will increase. decrease and the rate of descent will decrease. increase and the rate of descent will decrease. increase and the rate of descent will increase.

8. If you increase the propeller pitch during a glide with idle-power at constant IAS the lift to drag ratio will

A B C D

increase and the rate of descent will increase. decrease and the rate of descent will increase. increase and the rate of descent will decrease. decrease and the rate of descent will decrease.

9. For this question use the reference. The diagram the letter which correctly represents the Angle of Advance (Helix Angle) is:

A B C D

B C A D

Blade Angle = Angle of Attack + Helix Angle Helix Angle

= Blade Angle

-

Angle of Attack

10. Use graphic at reference. The angle of attack of a rotating propeller blade element shown in the annex is:

A B C D

angle 2. angle 3. not correctly indicated in the diagram. angle 1.

11. Use graphic at reference. The blade angle of a rotating propeller blade element shown in the annex is:

A B C D

angle 3. angle 2. not correctly indicated in the diagram. angle 1.

12. If you pull back the RPM lever of a constant speed propeller during a glide with idle power and constant speed, the propeller pitch will:

A B C D

increase and the rate of descent will decrease. decrease and the rate of descent will increase. decrease and the rate of descent will decrease. increase and the rate of descent will increase.

If your engine is out, gliding with the prop at the cruise setting would keep engine rpm up, the prop is windmilling, keeping the hydraulics and electrics going. But there is much straight drag from the prop, and pulling back to coarse will give less straight drag and a reduced rate of descent. What happens, though, is that the windmilling effect lessens and the prop slows down. So, if you need every last yard to reach your chosen field pull the prop lever right back. If this is a practice forced landing DO NOT FORGET to run the lever forward for the overshoot!

Number: 14823 Question: Which of these statements about propellers is correct or incorrect? I. A cruise propeller has a greater geometric pitch when compared with a climb propeller. II. A coarse pitch propeller is less efficient during take-off and in the climb, but more efficient in the cruise, when compared with a fine pitch propeller.

A B C D

I is correct, II is incorrect. I is correct, II is correct. I is incorrect, II is correct. I is incorrect, II is incorrect.

Slip = Geometric Pitch - Effective Pitch

14823 - B

PROPELLER SOLIDITY

1. Increasing the number of propeller blades will:

A B C D

decrease the torque in the propeller shaft at maximum power. increase the maximum absorption of power. increase the noise level at maximum power. increase the propeller efficiency. Efficiency is power out compared to power in, so you need an aerofoil design that has a good lift/drag ratio. A high aspect blade (long and narrow) will give you low induced drag and so it needs low engine power to overcome drag. As propellers operate at big angles of attack to produce thrust, a low induced drag will give you a good efficiency. If you look at your light aircraft that is exactly what they have. Fairly long narrow blades (high aspect ratio) and good efficiency, so you don’t need a big engine to drive them. Unfortunately they don't move a lot of air backwards, so although they are fine if the aircraft weight is low, you need more disc solidity, if you need more thrust. The minute you increase disc solidity (with wider blades) you are reducing aspect ratio and losing efficiency. You will need a stronger engine to overcome the higher drag. Power absorption is the same as disc solidity in that it reduces efficiency. If you have a powerful engine you are wasting it if you put a high aspect ratio (low disc solidity) propeller on it, because you could drive it at idle. You might as well use the power available by increasing the disc solidity and produce more thrust. Increasing disc solidity or absorbing more available power will always be contrary to

2. Which is one of the disadvantages of increasing the number of propeller blades ?

A B C D

decrease propeller efficiency Higher tip-speed Less power can be absorbed by the propeller Increased noise

3. If S is the frontal area of the propeller disc, propeller solidity is the ratio of:

A B C D

the total frontal area of all the blades to S. the mean chord of one blade to S. S to the frontal area of one blade. the frontal area of one blade to S.

4. Which statement is correct for a propeller of given diameter and at constant RPM? I. Assuming blade shape does not change power absorption increases if the number of blades increases. II. Power absorption increases if the mean chord of the blades increases

A B C D

I is incorrect, II is incorrect. I is correct, II is incorrect. I is correct, II is correct. I is incorrect, II is correct.

Asymmetric Blade Effect The thrust asymmetry between the two blades causes the aircraft to yaw to the left on take-off.

Conversely, if an aircraft is flying yawed, then the asymmetry of the thrust causes a pitching moment.

1. Given an aeroplane with a propeller turning clockwise as seen from behind, the torque effect during the take off run will tend to:

A B C D

pitch the aeroplane nose down. pitch the aeroplane nose up. roll the aeroplane to the left. roll the aeroplane to the right.

2. Gyroscopic precession of the propeller is induced by:

A B C D

pitching and rolling. increasing RPM and yawing. pitching and yawing. increasing RPM and rolling.

3. Asymmetric propeller blade effect is mainly induced by:

A B C D

large angles of yaw. the inclination of the propeller axis to the relative airflow. large angles of climb. high speed.

4. Which statement about a propeller is correct? I. Asymmetric blade effect increases when engine power is increased. II. Asymmetric blade effect increases when the angle between the propeller axis and airflow through the propeller disc increases.

A B C D

I is incorrect, II is correct. I is correct, II is correct. I is incorrect, II is incorrect I is correct, II is incorrect.

5. With respect to the gyroscopic effects of precession acting upon the clockwise rotating propeller of a single acting aeroplane (when viewed from behind): i. The effect of pitch up is right yaw ii. The effect of right yaw is pitch down.

A B C D

i. is correct ii. is incorrect i. is incorrect ii. is correct i. is incorrect ii. is incorrect i. is correct ii. is correct

6. Which statement is correct regarding the gyroscopic effect of a clockwise propeller on a single engine aeroplane? I.Pitch down produces left yaw. II.Left yaw produces pitch up.

A B C D

I is correct, II is incorrect I is correct, II is correct I is incorrect, II is incorrect I is incorrect, II is correct

7. Which statement is correct? I. Propeller gyroscopic effect occurs during aeroplane pitch changes. II. Propeller gyroscopic effect is most noticeable during low speed flight at high propeller RPM.

A B C D

I is incorrect, II is correct. I is incorrect, II is incorrect. I is correct, II is incorrect. I is correct, II is correct.

8. Which statement is correct? I. Propeller gyroscopic effect occurs during flight at constant aeroplane attitude. II. Propeller gyroscopic effect is most noticeable during low speed flight at low propeller RPM.

A B C D

I is correct, II is correct I is incorrect, II is correct. I is incorrect, II is incorrect. I is correct, II is incorrect.

9. Which statement is correct? I. Propeller gyroscopic effect occurs during aeroplane yaw changes. II. Propeller gyroscopic effect is most noticeable during low speed flight at high propeller RPM.

A B C D

I is correct, II is correct. I is incorrect, II is incorrect. I is correct, II is incorrect. I is incorrect, II is correct.

1. Which statement is correct about an aeroplane, that has experienced a left engine failure and continues afterwards in straight and level cruise flight with wings level?

A B C D

turn indicator neutral, slip indicator left of neutral. turn indicator left of neutral, slip indicator left of neutral. turn indicator left of neutral, slip indicator neutral. turn indicator neutral, slip indicator neutral. It is certainly possible to achieve straight asymetric flight with level wings by use of rudder. As you say, this would involve the aircraft slipping. But why would this deflect the ball? The wings (and therefore the instrument and the central portion of the tube) are level. The aircraft is not accelerating laterally, so the tube isn't being snatched away sidewayds, leaving the ball deflected by lagging behind. The balls weight is acting towards the Earth, as per usual, and is opposed by the ball being held up by the bottom edge of the central part of the tube. Again, no unbalanced forces. So there is no mechanism to cause the ball to be deflected, so it isn't.

2. During the take-off roll with a strong crosswind from the left, a four engine jet aeroplane with wing mounted engines experiences an engine failure. The greatest control problem is caused by the loss of which engine?

A B C D

The right inboard engine. The right outboard engine. The left inboard engine. The left outboard engine.

Yes the worst scenario is an outboard engine failure on the into wind wing. So outboard left in this case. The wind will already cause a yaw into wind, so if outboard goes then the thrust on the right will aggravate the yaw to the left. It can get to the point where you have insufficient rudder authority to keep straight.

3. In general, directional controllability with one engine inoperative on a multi-engine aeroplane is adversely affected by: adversely {adv} - ungünstig 1. high temperature. 2. low temperature. 3. aft CG location. 4. forward CG location. 5. high altitude. 6. low altitude. The combination that regroups all of the correct statements is

A B C D

1, 4, 5. 2, 3, 5. 2, 3, 6. 1, 4, 6.

9. In general, directional controllability with one engine inoperative on a multi-engine aeroplane is favourably affected by: 1. high temperature. 2. low temperature. 3. aft CG location. 4. forward CG location. 5. high altitude. 6. low altitude. The combination that regroups all of the correct statements is:

A B C D

2, 4, 6. 2, 3, 5. 1, 4, 6. 1, 4, 5.

With one engine inoperative on a multi-engine aeroplane, the loss of directional control is due to asymmetric thrust i.e. the turning-moment at high thrust settings. Hence, if the temperature is high, the thrust will be less - reducing the turning-moment at high power. Likewise, at high altitude, the air density is reduced, once again resulting in less thrust being produced from the remaining live engine(s), therefore resulting in a lesser turning-moment.

Number: 13127 Question: For a given aeroplane which two main variables determine the value of VMCG? VMCG Minimum Control Speed – Ground minimale Kontrollgeschwindigkeit beim Startlauf am Boden, wenn das Flugzeug den Start fortsetzten soll – gilt nicht für leichte Zweimots

A Engine thrust and rudder deflection. B Airport elevation and temperature. C Engine thrust and gear position. D Air density and runway length. The VMCG is calculated at full rudder deflection so that is not a variable. Yes thrust will reduce with altitude if not supercharged. The VMCG speed is calculated at the maximum thrust available. Remember the question does say which 2 variables will affect it?

The definition of VMCL is the minimum control speed at which it is possible to maintain directional control in the landing configuration with an angle of bank of not more than 5'.

Number: 4552 Question: How does VMCG change with increasing field elevation and temperature?

A B C D

increases, because VMCG is related to V1 and VR and those speeds increase if the density decreases. decreases, because the engine thrust decreases. decreases, because VMCG is expressed in IAS which decreases with constant TAS and decreasing density. increases, because at a lower density a larger IAS is necessary to generate the required rudder force.

Asymmetric thrust is balanced by rudder forces. At full rudder deflection if you fly faster you get more correcting force: Fly slower and you get less. Minimum control speeds are the speeds at which the rudder forces just balance the asymmetric thrust. Less thrust, and you can balance that with full rudder at a lower speed. Obviously Vmcg has to be down to, at most, V1, or you couldn't control the continuing takeoff after engine out. No point in having Vmcg much below that, for before V1 you are going to shut down the live engine(s) and there will be no asymmetric thrust. Vmcg is certified in the worst case, and the real benefit of reduced max thrust when you are hot and high is that less rudder will be needed to keep straight.

VMCG Minimum Control Speed – Ground minimale Kontrollgeschwindigkeit beim Startlauf am Boden, wenn das Flugzeug den Start fortsetzten soll – gilt nicht für leichte Zweimots V1 Critical Engine Failure Speed; Decision Speed for Take-off Engine Failure; Takeoff Decision Speed. Startabbruchgeschwindigkeit; Entscheidungsgeschwindigkeit, bis zu der ein Startabbruch möglich ist (z. B. B737 ca. 150 kt); Entscheidungsgeschwindigkeit bei Triebwerksausfall beim Start; Geschwindigkeit, bis zu der bei Ausfall eines kritischen Motors der Start abgebrochen werden kann; ist das Flugzeug schneller, als V1, dann muss es auch bei einem Triebwerksausfall starten; V1 gilt für größere Flugzeuge, die auch nach einem Triebwerksausfall eine ausreichende und genau definierte Steigleistung haben.

13127 - B

4552 - B

Number: 14681 Question: Which of these statements about VMCG determination are correct or incorrect? I. In order to simulate a wet runway, nose wheel steering may not be used during VMCG determination. II. During VMCG determination, the CG should be on the aft limit.

A B C D

I is incorrect, II is correct. I is correct, II is incorrect. I is incorrect, II is incorrect. I is correct, II is correct.

Number: 3854 Question: 6. Why is VMCG determined with the nosewheel steering disconnected?

A B C D

Because the nosewheel steering could become inoperative after an engine has failed. Because the value of VMCG must also be applicable on wet and/or slippery runways. Because nosewheel steering has no effect on the value of VMCG. Because it must be possible to abort the take-off even after the nosewheel has already been lifted off the ground.

Number: 15314 Question: Consider the following statements about VMCG: 1. VMCG is determined with the gear down. 2. VMCG is determined with the flaps in the landing position. 3. VMCG is determined by using rudder and nosewheel steering 4. During VMCG determination the aeroplane may not deviate from the straight-line path by more than 30 ft. The combination that regroups all of the correct statements is:

A B C D

1, 4. 1, 2, 3, 4. 3. 2, 3.

VMCG Minimum Control Speed – Ground minimale Kontrollgeschwindigkeit beim Startlauf am Boden, wenn das Flugzeug den Start fortsetzten soll – gilt nicht für leichte Zweimots

Number: 14680 Question: Which of these statements about VMCG determination are correct or incorrect? I. VMCG may be determined using both lateral and directional control. II. During VMCG determination, the lateral deviation from the runway centreline may be not more than 30 ft.

A B C D

I is correct, II is correct. I is incorrect, II is incorrect. I is correct, II is incorrect I is incorrect, II is correct.

14681 - D

3854 - B

VMCG Minimum Control Speed – Ground minimale Kontrollgeschwindigkeit beim Startlauf am Boden, wenn das Flugzeug den Start fortsetzten soll – gilt nicht für leichte Zweimots

5314 - A

14680 - D

Number: 12972 Question: VMCA is certified with a bank angle of not more than 5° towards the operating engine (live engine low) because:

A B C D

at 5° bank (live engine low), sideslip is zero. the slip indicator at 5° bank (live engine low) is centred. although more bank reduces VMCA, too much bank may lead to fin stall. more than 5° bank (live engine low) would not reduce VMCA. The speed is established with failure of the critical engine, not more than 5 degrees of bank towards the live engine and the aircraft in the most critical configuration for approach. VMCA Minimum Control Speed – Airborne, Minimum Air Control Speed (siehe: Rotieren) – minimale Geschwindigkeit, bei der das Flugzeug unter näher spezifizierten Bedingungen geflogen und gesteuert werden kann; britisches Englisch: VMCA; US-Englisch: VMC

Number: 15099 Question: VMCA is the minimum speed at which directional control can be maintained when, amongst others: 1. maximum take-off thrust was set and is maintained on the remaining engines. 2. a sudden engine failure occurs on the most critical engine. 3. flaps are in any position. 4. the gear is either up or down. 5. the aeroplane is either in or out of ground effect. The combination that regroups all of the correct statements is:

A B C D

3, 4 and 5 are correct. 1, 2 and 4 are correct. 2, 3 and 5 are correct. 1 and 2 are correct.

The speed is established with failure of the critical engine, not more than 5 degrees of bank towards the live engine and the aircraft in the most critical configuration for approach.

VMCA Minimum Control Speed – Airborne, Minimum Air Control Speed (siehe: Rotieren) – minimale Geschwindigkeit, bei der das Flugzeug unter näher spezifizierten Bedingungen geflogen und gesteuert werden kann; britisches Englisch: VMCA; US-Englisch: VMC

12972 - C

15099 - D

If the critical engine fails during the approach and landing phase of flight, maintain the airspeed above VMCL, which is the minimum control speed during the approach and landing. The definition of VMCL is the minimum control speed at which it is possible to maintain directional control in the landing configuration with an angle of bank of not more than 5°. VMCL must be established with the: - Aircraft in its most critical landing configuration and all engines operating - Centre of gravity as far aft as possible and the aircraft at its maximum landing weight - Propeller on the failed engine (propeller aircraft only) in the position it achieves without pilot action, whilst maintaining a 3' glide slope - Go-around power/thrust setting on the operating engine VMCL is a fixed value for a given aircraft, but VMCA and VMCG both decrease with increasing altitude.

Number: 14876 Question: VMCL is the: A minimum control speed - with landing gear down, flaps up and all engines operating. B maximum speed in the landing configuration. C minimum control speed approach and landing. D minimum speed during landing with all engines operating. VMCL Minimum Control Speed for Landing Approach, Minimum Control Speed in Landing Configuration (alle Motoren funktionieren) – siehe: VMCG

Number: 13071 Question: Which of the following statements is correct? I. VMCL is the minimum control speed in the landing configuration. II. The speed VMCL is always limited by maximum rudder deflection.

A B C D

I is incorrect, II is incorrect. I is incorrect, II is correct. I is correct, II is correct. I is correct, II is incorrect.

Number: 13072 Question: Which of the following statements is correct? I When the critical engine fails during take-off the speed VMCL can be limiting. II The speed VMCL is always limited by maximum rudder deflection.

A B C D

I is correct, II is correct. I is incorrect, II is incorrect. I is incorrect, II is correct. I is correct, II is incorrect.

Number: 4548 Question: Which of the following statements is correct? I When the critical engine fails during take-off the speed VMCL can be limiting. II The speed VMCL can be limited by the available maximum roll rate.

A B C D

I is correct, II is correct. I is incorrect, II is correct. I is correct, II is incorrect. I is incorrect, II is incorrect.

14876 - C

13071 - D

13072 - B

4548 - B

If the critical engine fails during the approach and landing phase of flight, maintain the airspeed above VMCL, which is the minimum control speed during the approach and landing. The definition of VMCL is the minimum control speed at which it is possible to maintain directional control in the landing configuration with an angle of bank of not more than 5°. VMCL must be established with the: - Aircraft in its most critical landing configuration and all engines operating - Centre of gravity as far aft as possible and the aircraft at its maximum landing weight - Propeller on the failed engine (propeller aircraft only) in the position it achieves without pilot action, whilst maintaining a 3' glide slope - Go-around power/thrust setting on the operating engine VMCL is a fixed value for a given aircraft, but VMCA and VMCG both decrease with increasing altitude.

13073. Which of the following statements is correct? I When the critical engine fails during take-off the speed VMCL can be limiting. II The speed VMCL can be limited by the available maximum roll rate.

A B C D

I is correct, II is correct. I is incorrect, II is incorrect. I is correct, II is incorrect. I is incorrect, II is correct.

1310. Which of the following statements is correct? I VMCL is the minimum control speed in the landing configuration. II The speed VMCL can be limited by the available maximum roll rate.

A B C D

I is incorrect, II is incorrect. I is correct, II is correct. I is incorrect, II is correct. I is correct, II is incorrect.

13073 - D

13010 - B

Number: 14616 Question: The stall speed line in the manoeuvring load diagram runs through a point where the:

A B C D

speed = VB, load factor = gust load factor. speed = 0, load factor = +1. speed = VA, load factor = limit load factor. speed = VS, load factor = 0.

VS - Unaccelerated stall speed in the clean configuration

VA - Maximum design manoeuvring speed and is the highest speed where the aircraft will stall before it exceeds the maximum load factor. VB - Maximum design speed for maximum gust intensity.

Vc - Maximum design cruising speed. It is a given JAR25 requirement. VC is a design speed NOT an operating speed. VD is the design dive speed. If you go there the wings will come off. VMO/MMO may be equal to, but may not be greater than VC (the design cruise speed). The reasoning behind that is, that if you are flying at VMO/MMO, and you get a disturbance nose down, the speed WILL increase, hopefully you will be able to recover before VD.

Number: 14916 Question: Which factor should be taken into account when determining VA?

A B C D

The limit load factor. The safety factor. The calculation factor. The ultimate load factor.

14616 - C

14916 - A

Number: 15033 Question: Assuming no compressibility effects, the correct relationship between stall speed, limit load factor (n) and VA is:

A B C D

VA = VS * SQRT (n).

Number: 16811 Question: The significance of VA for jet transport aeroplanes is reduced at high cruising altitudes because:

A B C D

the engine has insufficient thrust to reach the limit load factor. at high altitudes the bank angle is normally limited to 15° to prevent exceeding the limit load factor. buffet onset limitations normally become limiting. the elevator deflection is limited to prevent exceeding the limit load factor. to buffet

stoßen schlagen ankämpfen

onset

Beginn {m} Anfang {m} Ausbruch {m} [einer Krankheit] Einsetzen {n} [einer Krankheit, des Alters, des Winters] Angriff {m} Einsatz {m} [Start / Einleitung eines Vorgangs], Attacke, Eintritt, Sturm {m}

Number: 4375 Question: Which of the following statements is true?

A

Flap extension in severe turbulence at constant speed moves the centre of pressure aft, which increases the structural limitation margins. B Flight in severe turbulence may lead to a stall and/or structural limitations being exceeded. C Flap extension in severe turbulence at constant speed increases both the stall speed and the structural limitation margins. D By increasing the flap setting in severe turbulence at constant speed the stall speed will be reduced and the risk for exceeding the structural limits will be decreased.

15033 - D

16811 - C

4375 - B

Number: 13211 Question: How does VA (EAS) alter when the aeroplane's mass decreases by 19%?

A B C D

4.36% reduction. 19% reduction. No change 10% reduction.

Once you set Cn at 2.5 for working out Va the answer for Va is Vs1g times sq root 2.5. If you reduce aircraft mass you get a new lower Vs1g and therefore a new lower Va. Because the lift equation has Vsq in it the reduction in Vs1g is proportional to sq root of the change in mass It is a quirk of the mathematics that, for small changes, inside 20%, the sq root of the proportional change is approximately half the original figure, so a 20% reduction in mass gives a 10% reduction in Va.

Manövergeschwindigkeit (VA) Die Manövergeschwindigkeit (engl. Design Maneuvering Speed o. a. Maximum Manoeuverspeed; in der Luftfahrt als VA abgekürzt) ist die Fluggeschwindigkeit (engl. Indicated Air Speed, IAS), bei der durch plötzliche volle Ruderausschläge die Struktur des Flugzeuges bis zur Grenze des Zulässigen belastet wird. Bedeutung: Volle Ruderausschläge sind oberhalb der Manövergeschwindigkeit nicht zulässig, da sie die Flugzeugstruktur überlasten würden. Aber auch unterhalb der Manövergeschwindigkeit gilt das nur für die einmalige Betätigung eines der drei Ruder (Höhen-, Quer- oder Seitenruder), wiederholte Ausschläge oder kombinierte Vollausschläge mehrerer Ruder gleichzeitig können die Struktur durchaus auch unterhalb der Manövergeschwindigkeit überlasten. Die Manövergeschwindigkeit VA wird oft mit der maximalen Geschwindigkeit in unruhiger Luft verwechselt. Diese wird VB oder VRA (RA wie rough air) bezeichnet und ist auf dem Fahrtmesser mit einem grünen (unterhalb von VRA) bzw. gelben (oberhalb von VRA) Bogen gekennzeichnet.

Abhängigkeit vom Fluggewicht Die VA wird auch bei kleinen Flugzeugen in Abhängigkeit vom Fluggewicht angegeben. - Ein hohes Fluggewicht bedingt dabei eine hohe Manövergeschwindigkeit, - ein niedriges Fluggewicht eine eher niedrige Manövergeschwindigkeit. Je schwerer ein Flugzeug, umso höher ist VA , denn mit mehr Treibstoff in den Tanks und folglich höherem Gewicht widersteht das Flugzeug wegen der Massenträgheit den aerodynamischen Kräften besser. Dabei spielt auch die Verteilung der Masse zwischen tragenden und nicht tragenden Teilen (also den Flügeln mit den Tanks einerseits und dem Cockpit andererseits) eine Rolle. Diese Feinheiten werden in den Handbüchern zu Kleinflugzeugen aber meist nicht berücksichtigt.

13211 - D

Number: 211 Question: The shape of the gust load diagram is also determinated by the following three vertical speed in ft/s (clean configuration) :

A B C D

25, 55, 75 35, 55, 66 15, 56, 65 25, 50, 66

Number: 11017 Question: The lift coefficient (CL) of an aeroplane in steady horizontal flight is 0.42. An increase in angle of attack of 1 degree increases CL by 0.1. A vertical up gust instantly changes the angle of attack by 3 degrees. The load factor will be:

A B C D

2.49 0.74 1.71 1.49

LF is lift / weight, or n=L/W. All other factors being equal lift and weight can be expressed as changes of Cl for the purpose of the calculation. Weight does not change, and so the Cl required is the same after the event as before (when considering the calculation). However lift increases as per the conditions in the question. In this case it increases by 3 x 0.1 =0.3.

So for the calculation the Cl after the event is 0.42+0.3=0.72. Divide this by the original Cl to get the answer. Mathematically it looks like: (0.42+(3x0.1)) / 0.42. Beware some questions ask for the change of load factor, in which case you have to subtract the original LF from the new one.

Number: 14109 Question: The lift coefficient (CL) of an aeroplane in steady horizontal flight is 0.35. An increase in angle of attack of 1 degree would increase CL by 0.079. If a vertical gust instantly changes the angle of attack by 2 degrees, the load factor will be:

A B C D

0.9 1.9 1.45 0.45

The lift coefficient (CL) of an aeroplane in steady horizontal flight is 0.35. An increase in angle of attack of 1 degree would increase CL by 0.079. If a vertical gust instantly changes the angle of attack by 2 degrees, the load factor will be: 1 deg increase = 0.079 x 2 = 0.155 NEW CL 0.35 + 0.155 = 0.508 / OLD CL 0.35 = 1.451

211 - D

11017 - C

14109 - C

Number: 16247 Question: An aeroplane maintains straight and level flight at a speed of 2 * VS. If a vertical gust causes a load factor of 2, the load factor n caused by the same gust at a speed of 1.3 VS would be:

A B C D

n = 1.3. n = 1.65. n = 4. n = 1.69.

It is 1 x 1.3/2 = 0.65 in normal steady flight you have 1g already, so just add 1 + 0.65 and you will get 1.65

Number: 14305 Question: VRA is:

A B C D

a speed just below Mach buffet. a speed just above low speed buffet. the stall speed in turbulent conditions. the recommended turbulence penetration airspeed.

16247 - B

14305 - D

Die Manövergeschwindigkeit VA wird oft mit der maximalen Geschwindigkeit in unruhiger Luft verwechselt. Diese wird VB oder VRA (RA wie rough air) bezeichnet und ist auf dem Fahrtmesser mit einem grünen (unterhalb von VRA) bzw. gelben (oberhalb von VRA) Bogen gekennzeichnet.

Number: 14673 Question: Which of these statements about the gust load factor on an aeroplane are correct or incorrect? I. When the mass decreases, the gust load factor increases. II. When the altitude decreases, the gust load factor increases.

A B C D

I is correct, II is correct. I is correct, II is incorrec I is incorrect, II is correct. I is incorrect, II is incorrect. Abhängigkeit vom Fluggewicht Die VA wird auch bei kleinen Flugzeugen in Abhängigkeit vom Fluggewicht angegeben. - Ein hohes Fluggewicht bedingt dabei eine hohe Manövergeschwindigkeit, - ein niedriges Fluggewicht eine eher niedrige Manövergeschwindigkeit. Je schwerer ein Flugzeug, umso höher ist VA , denn mit mehr Treibstoff in den Tanks und folglich höherem Gewicht widersteht das Flugzeug wegen der Massenträgheit den aerodynamischen Kräften besser. Dabei spielt auch die Verteilung der Masse zwischen tragenden und nicht tragenden Teilen (also den Flügeln mit den Tanks einerseits und dem Cockpit andererseits) eine Rolle. Diese Feinheiten werden in den Handbüchern zu Kleinflugzeugen aber meist nicht berücksichtigt.

Number: 14674 Question: Which of these statements about the gust load factor on an aeroplane are correct or incorrect? I. When the mass increases, the gust load factor increases. II. When the altitude increases, the gust load factor increases.

A B C D

I is correct, II is correct. I is correct, II is incorrect. I is incorrect, II is incorrect. I is incorrect, II is correct.

Number: 14675 Question: Which of these statements about the gust load factor on an aeroplane are correct or incorrect? I. When the wing area increases, the gust load factor increases. II. When the EAS increases, the gust load factor decreases.

A B C D

I is correct, II is incorrect. I is incorrect, II is correct. I is correct, II is correct. I is incorrect, II is incorrect.

14673 - A

14674 - C

14675 - A

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

A B C D

point 4. point 1. point 3. point 2.

2. From a polar curve of the entire aeroplane one can read:

A B C D

the minimum CL/CD ratio and the minimum drag. the maximum CL/CD ratio and maximum lift coefficient. the minimum drag coefficient and the maximum lift. the minimum drag and the maximum lift.

3. Which of the following parameters can be read from the parabolic polar diagram of an aeroplane?

A B C D

The minimum glide angle and the parasite drag coefficient. The aspect ratio of the wing and the induced drag coefficient. The minimum rate of descent and the induced drag. The induced drag and the parasite drag. The polar diagram's axes are Cl and Cd. Where the line crosses the Cl=0 axis there is no lift, and therefore no induced drag, so Cd at that point is from parasite drag only. Angle of glide is Cl/Cd, the lift/drag ratio. If the lift/drag ratio is 20:1 then the glide angle is 1 in 20. The tangent to the curve from the Cl/Cd origin shows the best Cl/Cd ratio, and thus the best glide angle you can get. Rate of descent would depend on what speed you were doing down the hill and that is not shown on the polar diagram. The polar diagram shows coefficients only. To get real world data you have to add in all the other bits of the lift and drag equations like S and rho and V.

Number: 14489 Question: During a straight steady climb: 1 - lift is less than weight. 2 - lift is greater than weight. 3 - load factor is less than 1. 4 - load factor is greater than 1. 5 - lift is equal to weight. 6 - load factor is equal to 1. Which of the following lists the correct statements ?

A B C D

1 and 6 1 and 3 2 and 4 5 and 6

To answer this question it is helpful to draw a diagram of the forces acting on the aircraft whilst it is in the climb, and resolve them accordingly. If you do this, you can see that the lift vector is acting at 90 degrees to the flight path. (The flight path being at an angle to the horizon) However, the weight vector will always act straight down, or at 90 degrees to the horizon. So only a part of this weight vector is acting in the direction opposite to the lift. (The other part of the weight vector acts in the same direction as the drag) Therefore, in order to keep the forces balanced, so the aircraft isn't accelerating: Lift = The part of the weight vector acting perpendicular to the flight path. To be completely mathematical L=W cos(gamma); where gamma is the flight path angle to the horizontal. So you can see that lift is now a fraction of the weight, so lift is less than weight. And mathematically this is shown because cos(gamma) can only have a value between 0 and 1. With it equalling 1 when gamma is zero (ie horizontal flight) and 0 when gamma is 90 degrees (ie full on vertical climb) The load factor bit comes from the definition of load factor, which is Lift/Weight. Now we've already determined that Lift is less than Weight in steady climb, so this fraction has a numerical value less than one. Again being mathematical, you can say from Lift = Weight x Cos(gamma) Lift/Weight = Cos(gamma) hence Load factor = Lift/Weight = Cos (gamma) And once again you can see that load factor can range from zero to one, with it equalling one, only in straight and level flight.

14489 - B

1. The lift of an aeroplane of weight W in a constant linear climb with a climb angle (gamma) is approximately:

A B C D

W / cos.gamma. W * (1-tan.gamma). W * (1-sin.gamma). W * cos.gamma.

Lift by definition act 90 deg to the relative airflow. So to calculate the lift you must consider the forces at right angles to the airflow. That is what the lift = Wcos Gamma

Number: 13137 Question: An aeroplane climbs to cruising level with a constant pitch attitude and maximum climb thrust, (assume no supercharger). How do the following variables change during the climb? (gamma = flight path angle)

A B C D

gamma remains constant, angle of attack remains constant, IAS decreases. gamma decreases, angle of attack increases, IAS remains constant. gamma decreases, angle of attack remains constant, IAS decreases. gamma decreases, angle of attack increases, IAS decreases.

If you are holding a constant pitch angle it means the angle between the longitudinal axis and the horizontal is constant. (constant nose position) If you are at max thrust at MSL, the thrust will decrease as you climb, so will have to increase the angle of attack to hold the nose position constant. That means the IAS will decrease. The gamma angle is the flight path inclination and that will decrease as well. At somepoint the thrust will be so low that you can only maintain S&L at the stalling angle of attack, so gamma will be zero and angle of attack and pitch angle will be the same.

3. A 50 ton twin engine aeroplane performs a straight, steady, wings level climb. If the lift/drag ratio is 12 and the thrust is 60 000N per engine, the climb gradient is: (assume g = 10m/s2)

A B C D

24% 3.7% 12% 15.7%

50 000 x10 = 500 000n drag =500 000/12=41666n Climb gradient = (Thrust -drag)/ weight X100 Thrust = 120 000 (dont forget the plane has 2 engines so (120 000 - 41666)/500000= 0,156 climb gradient = 0,1566x100 = 15,6 %

3. Given: Aeroplane mass: 50 000kg Lift/Drag ratio: 12 Thrust per engine: 28 000N Assumed g: 10m/s² For a straight, steady, wings level climb of a four-engine aeroplane, the one engine inoperative climb gradient is:

A B C D

8.5%. 8.0%. 2.9%. 9.7%.

5. Given: Aeroplane mass: 50 000kg Lift/Drag ratio: 12 Thrust per engine: 30 000N Assumed g: 10m/s² For a straight, steady, wings level climb of a three-engine aeroplane, the all engines climb gradient is:

A B C D

8.5%. 2.9%. 9.7%. 8.0%.

Put the mass into Newtons. 50000 x 10 = 500000N The drag is the lift required divided by 12 500000/12 = 41666N The angle of climb SIN THETA = T-D/W x 100 to express as a percentage. (90000-41666/500000) x 100 = 9.7%

6. Given: Aeroplane mass: 50 000kg Lift/Drag ratio: 12 Thrust per engine: 21 000N Assumed g: 10m/s² For a straight, steady, wings level climb of a four-engine aeroplane, the one engine inoperative climb gradient is:

A B C D

7.7 %. 8.5 %. 4.3 %. 6.0 %.

7. Given: Aeroplane mass: 50 000kg Lift/Drag ratio: 12 Thrust per engine: 21 000N Assumed g: 10m/s² For a straight, steady, wings level climb of a four-engine aeroplane, the all engines climb gradient is:

A B C D

4.3%. 8.5%. 6.0%. 7.7%.

8. When an aeroplane performs a straight steady climb with a 20% climb gradient, the load factor is equal to:

A B C D

1. 1.02. 0.83. 0.98.

Refer to diagram 'Forces in a climb' - PoF p5.3 Lift = Wcos(angle of climb). Angle of climb = 12° (20%). Assume W = 1 in level flight = lift Load Factor is ratio of Lift in Manoeuvre: Lift in Level Flight.

9. An aeroplane with a mass of 4000 kg is performing a co-ordinated level turn at a constant TAS of 160 kt and a bank angle of 45°. The lift is approximately:

A B C D

28000 N. 98000 N. 14000 N. 56000 N.

1. What is the approximate value of the lift of an aeroplane at a gross weight of 50000 N, in a horizontal co-ordinated 45 degrees banked turn?

A B C D

70000 N. 50000 N. 80000 N. 60000 N.

You are taking the square root of the load factor which is incorrect. "n" (load factor) = 1/Cos bank angle. You only take the square root if you are trying to find the new stalling speed. This question asks about the lift being generated.

Number: 175 Question: An aeroplane is in a level turn, at a constant TAS of 300 kt, and a bank angle of 45°. Its turning radius is: (given: g= 10 m/s²)

A B C D

2381 metres. 4743 metres. 3354 metres. 9000 metres.

1-A

Speed needs to be in m/s. 300kts is 153 m/s (about 1/2 is close enough for kts to m/s). 153^2/(10*tan45)=23716/10=2341 so 2381 is the closest answer.

175 - A

Number: 13465 Question: What is the approximate radius of a steady horizontal co-ordinated turn at a bank angle of 45° and a TAS of 200 kt?

A B C D

1.5 km. 2 km. 10 km. 1 km.

First off, the formula for motion in a circle, R = Vsqu/A refers to acceleration toward the centre, A. In a level turn, there are two accelerations acting on the aircraft, g and N. N is acceleration in the aircraft vertical, aka Load Factor. Because the aircraft is banked N is factored into one g vertically to hold the aircraft level, and A, the acceleration toward the centre. From the normal diagram you can see that geometrically, Tan AOB is A/g, so A = gTan AOB, and R = Vsqu/gTan AOB. Pure geometry! The diagram in our notes labels the arrows as forces (weight, Lift and CPforce), but the aircraft mass is determined, so acceleration and force are equivalent. It is now a second stage to find LF. Again, purely geometrically, N, or Load Factor is 1(g)/cos AOB

13465 - D

4. Approximately how long does it take to fly a complete circle during a horizontal steady co-ordinated turn with a bank angle of 45° and a TAS of 200 kt?

A B C D

95 s. 65 s. 650 s. 125 s.

The formula to calculate the circumference of a circle is: C = 2 x Pi x r = 2 x 3,1415 x 1060 m = 6660 m speed = distance / time --> time = distance / speed TAS = 200 kts = 103 m/s 6660 / 103 = 63 sec.

5. An aeroplane is in a steady horizontal turn at a TAS of 194.4 kt. The turn radius is 1000 m. The bank angle is: (assume g = 10 m/s2)

A B C D

30°. 60°. 45°. 50°.

6. An aeroplane performs a steady horizontal turn with a TAS of 200 kt. The turn radius is 2000 m. The load factor (n) is approximately:

A B C D

1.4. 2.0. 1.8. 1.1.

You require the bank angle to be able to work out the load factor because in a turn the load factor = 1/ Cos bank angle The formula for bank angle is Tan bank angle = TAS squared / g x radius You have to convert the TAS into SI units so 200 x 0.515 = 103 m/s 103 squared / 10 x 2000 = 0.53045 That is the tan value so the bank angle is 27.94371º We know “n” = 1 ÷ Cos bank angle, so “n” = 1.1319792

7. An aeroplane performs a steady co-ordinated horizontal turn with 20 degrees of bank and at 150 kt TAS. The same aeroplane with the same bank angle and speed, but at a lower mass will turn with:

A B C D

a larger turn radius. a higher turn rate. the same turn radius. a smaller turn radius.

If you look at the formula for radius and rate of turn you will see that weight does not appear? The ability for an aircraft to turn is independent of weight, you will often see questions that require you to have that knowledge.

8. What is the approximate diameter of a steady, level, co-ordinated turn with a bank angle of 30 degrees and a speed (TAS) of 500 kt?

A B C D

17 km. 23 km. 7 km. 13 km.

09. Two identical aeroplanes A and B, with the same mass, are flying steady level co-ordinated 20 degree bank turns. If the TAS of A is 130 kt and the TAS of B is 200 kt:

A B C D

the lift coefficient of A is less than that of B. the load factor of A and B are the same. the turn radius of A is greater than that of B. the rate of turn of A is less than that of B.

10. Two identical aeroplanes A and B, with the same mass, are flying steady level co-ordinated 20 degree bank turns. If the TAS of A is 130 kt and the TAS of B is 200 kt:

A B C D

the rate of turn of A is less than that of B. the turn radius of A is greater than that of B. the load factor of A is greater than that of B. the lift coefficient of A is greater than that of B.

Lift is the same. If V is high CL must be low and vice versa.

02-00 - TRANSONIC AERODYNAMICS

PRINCIPLES OF FLIGHT

02-01 The Mach number definition (23 Questions)

Number: 16327 Question: How does the Mach number change during a climb at constant IAS from sea level to 40,000 ft? A. B. C. D.

Decreases with increasing altitude. Initially remains constant until approximately 25,000 ft and then increases with increasing altitude. Increases with increasing altitude. Initially remains constant until approximately 25,000 ft and then decreases with increasing altitude.

Number: 16331 Question: Which statement with respect to the speed of sound is correct? A. Doubles if the temperature increases from 9° to 36° Centigrade. B. Increases always if the density of the air decreases. C. Is independent of altitude. D. Varies with the square root of the absolute temperature.

16327 - C;

16331 - D;

Number: 16275 Question: Assuming ISA conditions and a descent below the tropopause at constant Mach number and aeroplane mass, the: A. TAS remains constant. B. lift coefficient decreases. C. lift coefficient increases. D. IAS decreases.

Number: 16757 Question: Which of these statements about the transonic speed range is correct ? A. The airflow everywhere around the aeroplane is supersonic. B. The transonic speed range starts at Mcrit and extends to Mach numbers above M = 1. C. The transonic speed range starts at M = 0.5 and ends at Mcrit. D. The airflow everywhere around the aeroplane is subsonic.

16275 - B;

16757 - B;

Number: 1043 Question: Compressibility effects depend on: A. TAS. B. IAS. Right back at the beginning we said we would treat air as an incompressible fluid. C. Mach Number. Unfortunately it is not, and at higher Mach numbers we begin to see the effects of D. EAS. our mistake, like the difference between CAS and EAS followed by the formation of pressure waves and shockwave. These all depend on Mach number, although we often cite 300kt as the point at which we first notice things going wrong it is in fact M0.4

Number: 1044 Question: The formula for the Mach Number is: (a= speed of sound) A. M= a / TAS B. M= IAS / a C. M= TAS * a D. M= TAS / a

Number: 1045 Question: If the altitude is increased and the TAS remains constant in the troposphere under standard atmospheric conditions, the Mach number will: A. decrease. B. increase or decrease, depending on the type of aeroplane. C. increase. D. not change.

1043 - C;

1044 - D;

1045 - C;

Number: 1046 Question: Assuming ISA conditions, climbing at a constant Mach Number up the tropopause the TAS will: A. first increase, then decrease. B. increase. C. remain constant. D. decrease.

Number: 1047 Question: The speed of sound is determined only by: A. humidity. B. pressure. C. temperature. D. density.

Number: 2344 Question: An aeroplane is descending at a constant Mach number from FL 350. What is the effect on true airspeed? A. It remains constant. B. It increases as temperature increases. C. It decreases as altitude decreases. D. It decreases as pressure increases.

1046 - D;

1047 - C;

2344 - B;

Number: 3113 Question: The flight Mach number is 0.8 and the TAS is 400 kts. The speed of sound is: A. 600 kts B. 500 kts C. 320 kts D. 480 kts Number: 13100 Question: During a climb at a constant IAS, the Mach number will: A. increase initially and remain constant subsequently. B. decrease initially and increase subsequently. C. increase. D. remain constant.

Number: 13213 Question: if the Mach number is 0.8 and the TAS is 400 kt, what is the speed of sound? A. 500 kt. B. 320 kt. C. 600 kt. D. 480 kt.

Number: 13249 Question: In the transonic range the aeroplane characteristics are strongly determined by the: A. TAS. B. CAS. C. IAS. D. Mach number.

Number: 12995 Question: The Mach number is the ratio between the: A. TAS of the aeroplane and the speed of sound at sea level. B. IAS of the aeroplane and the speed of sound at sea level. C. IAS of the aeroplane and the speed of sound of the undisturbed flow. D. TAS of the aeroplane and speed of sound of the undisturbed flow.

3113 - B;

13100 - C;

13213 - A;

13249 - D;

12995 - D

Number: 13007 Question: A transonic Mach number is a Mach number: A. in the range between Mcrit and MMO. B. at which only supersonic local speeds occur. C. between the buffet onset Mach number and the Mach number corresponding to total supersonic flow. D. at which both subsonic and supersonic local speeds occur.

Number: 15426 Question: if the Mach number is 0.8 and the TAS is 480 kt, what is the speed of sound? A. 600 kt. B. 560 kt. C. 750 kt. D. 384 kt. Number: 1410 Question: The Mach number is the ratio between the: A. IAS of the aeroplane and the speed of sound of the undisturbed flow. B. IAS of the aeroplane and the speed of sound at sea level. C. TAS of the aeroplane and the speed of sound of the undisturbed flow. D. TAS of the aeroplane and the speed of sound at sea level. Number: 3875 Question: Which statement with respect to the speed of sound is correct? A. Doubles if the temperature increases from 9° to 36° Centigrade. B. Varies with the square root of the absolute temperature. C. Increases always if the density of the air decreases. D. Is independent of altitude.

13007 - D;

15426- B;

1410 - C;

3875 - B;

Number: 14578 Question: Regarding the transonic speed range: A. it starts at a Mach number just above 1. B. it starts at Mcrit and extends to Mach number equal to 1 C. it starts at a Mach number equal to 1. D. both subsonic and supersonic speeds exist in the flow around the aeroplane.

Number: 14579 Question: The subsonic speed range: A. implies both subsonic and supersonic speeds exist in the flow around the aeroplane. B. ends at M = 1. C. ends at Mcrit. D. ends at a Mach number just above M = 1.

14578 - D;

14579 - C;

Number: 14580 Question: Which of these statements about the supersonic speed range is correct? A. The supersonic speed range starts at a Mach number below M = 1 and extends to Mach numbers above M = 1. B. The airflow everywhere around the aeroplane is supersonic. C. The supersonic speed range starts at M = 1 and ends at Mcrit. D. The airflow around the aeroplane is transonic.

Number: 15003 Question: The speed range from approximately M=1.3 to approximately M=5 is called the: A. transonic range. B. supersonic range. C. subsonic range. D. hypersonic range.

14580 - B;

15003 - B;

02-00 - TRANSONIC AERODYNAMICS

PRINCIPLES OF FLIGHT

02-02 Normal shock waves - (67 Fragen)

Number: 16275 Question: Assuming ISA conditions and a descent below the tropopause at constant Mach number and aeroplane mass, the: A. TAS remains constant. B. lift coefficient decreases. C. lift coefficient increases. D. IAS decreases.

16275 - B

Number: 16450 Question: Behind a normal shock wave on an aerofoil section the local Mach number is: A. higher than before. B. equal to 1. C. less than 1. D. lower than before but still greater than 1.

Number: 204 Question: Air passes a normal shock wave. Which of the following statements is correct? A. The static pressure decreases. B. The static temperature decreases. C. The velocity increases. D. The static temperature increases.

Number: 206 Question: Two methods to increase the critical Mach number are: A. thick aerofoils and dihedral of the wing. B. thin aerofoils and dihedral of the wing. C. positive cambering of the aerofoil and sweepback of the wing. D. thin aerofoils and sweepback of the wing.

16450 - C;

204 - D;

206 - D;

Number: 952 Question: Which of these statements about "tuck under" are correct or incorrect? I. "Tuck under" is caused by an aft movement of the centre of pressure of the wing. II. "Tuck under" is caused by a reduction in the downwash angle at the location of the horizontal stabiliser. A. I is correct, II is correct. B. I is incorrect, II is correct. C. I is incorrect, II is incorrect. D. I is correct, II is incorrect.

952 - A;

Number: 962 Question: An A 310 aeroplane weighing 100 tons is turning at FL 350 at constant altitude with a bank of 50 degrees. Its flight Mach range between low-speed buffeting and high-speed buffeting goes from: A. M= 0.65 to M higher than 0.84 B. M= 0.74 to M= 0.84 C. M= 0.69 to M higher than 0.84 D. M= 0.72 to M higher than 0.84

ANSWER (c)

962 - C;

ANSWER Question Number: 962

Number: 1017 Question: How does stall speed (IAS) vary with altitude? A. It remains constant at lower altitudes but decreases at higher altitudes due to compressibility effects. B. It remains constant. C. It increases with increasing altitude, because the density decreases. D. It remains constant at lower altitudes but increases at higher altitudes due to compressibility effects. Basically, 1g stall speed in EAS stays constant. Higher up, because of compressibility, IAS is slightly higher than EAS, so IAS stall speed is a little higher. Additionally, at very high altitude Mach effects come in at EAS as low as the 1g stall speed, reducing available CLmax and further raising the 1g stall speed. See the buffet boundary charts in the notes. Whatever EAS and IAS/CAS are doing, the higher you go the greater TAS you get for any EAS or IAS. Number: 1018 Question: What data may be obtained from the Buffet Onset Boundary chart? A. The values of the Mach number at which low speed and shock stall occur at different weights and altitudes. B. The values of Mcrit at different weights and altitudes. C. The values of MMO at different weights and altitudes. D. The values of the Mach Number at which low speed and Mach Buffet occur at different weights and altitudes.

1017 - D

1018 - D

Number: 1019 Question: Mcrit is the free stream Mach Number at which: A. somewhere about the airframe Mach 1 is reached locally. B. Mach buffet occurs. C. the critical angle of attack is reached. D. shock stall occurs.

Number: 1020 Question: Which of the following (1) aerofoils and (2) angles of attack will produce the lowest Mcrit values? A. (1) thin and (2) large. B. (1) thick and (2) large. C. (1) thin and (2) small. D. (1) thick and (2) small. Number: 1021 Question: During which type of stall does the angle of attack have the smallest value? A. Shock stall. B. Accelerated stall. C. Deep stall. D. Low speed stall. Shock Stall is when the airflow over the wing becomes supersonic; the resulting shockwave causes flow separation... loss of lift. Number: 1022 Question: When the Mach number is slowly increased in straight and level flight the first shock waves will occur: A. somewhere on the horizontal tail. B. on the upper surface at the wing root. C. somewhere on the fin. D. on the lower surface of the wing.

Number: 1023 Question: The consequences of exceeding Mcrit in a swept-wing aeroplane may be: (assume no corrective devices, straight and level flight) buffeting - stoßend A. an increase in speed and a tendency to pitch up. B. engine unbalance and buffeting. C. buffeting of the aeroplane and a tendency to pitch up. D. buffeting of the aeroplane and a tendency to pitch down.

1019 - A

1020 - B

1021 - A

1022 - B

1023 - D

Number: 1026 Question: The maximum acceptable cruising altitude is limited by a minimum acceptable loadfactor because exceeding that altitude: A. a sudden necessary bankangle may exceed the limit load factor. buffet - stoß B. turbulence may exceed the limit load factor. C. Mach buffet will occur immediately. A common characteristic of shock wave induced separation is the D. turbulence may induce Mach buffet.

increasing severity of buffet intensity with increasing Mach number.

In fact, it is possible that the maximum angle of attack may not be achievable due to this severe buffet intensity. The aircraft's manoeuvring capability (load factor) is also reduced.

Number: 1032 Question: Shock stall: A. results from flow separation at the underside of the aerofoil. B. occurs as soon as Mcrit is exceeded. C. occurs when the lift coefficient, as a function of Mach number, reaches its maximum value. D. results from flow separation behind the bow wave.

This is called the shock stall. It differs from a conventional low-speed stall because it normally occurs at low angles of attack, although with increasing angles of attack the stall will occur at a lower Mach number.

Number: 1033 Question: In the transonic range the aeroplane characteristics are strongly determined by the: A. Mach number. B. CAS. C. TAS. D. IAS.

1026 - D

1032 - C

1033 - A

Number: 1034 Question: Which of the following flight phenomena can only occur at Mach numbers above the critical Mach number? A common characteristic of shock wave induced separation is the increasing severity of A. Elevator stall. buffet intensity with increasing Mach number. B. Mach buffet. C. Speed instability. In fact, it is possible that the maximum angle of attack may not be achievable due to this D. Dutch roll. severe buffet intensity. The aircraft's manoeuvring capability (load factor) is also reduced.

buffet - stoß Number: 1035 Question: Which of the following flight phenomena can occur at Mach numbers below the critical Mach number? A. Tuck under. B. Dutch roll. C. Shock stall. D. Mach buffet. The Spiral Dive... ...is probably responsible for more deaths than the dreaded spin... It is the one manoeuvre that will break any aeroplane ever built. There are two ways in which spiral dives are often entered: 1. From a steep turn. 2. Entry into poor weather without instrument capability. The steep turn is controlled by reference to the Earth's horizon. If we let the nose drop, the speed will increase and height will be lost rapidly. If you fly into cloud without the training and the instruments to fly blind, you will enter a spiral dive within 178 seconds! In cloud you'll see the height loss and the speed increase. The natural tendency is to pull back on the control column. But this deepens the spiral, increases the "g" force, and increases the airspeed. Spiral dive recognition and recovery You will recognise the spiral by the following: 1. 2. 3.

The speed is increasing rapidly. Height is lost at an increasing rate. The "g" force is increasing.

In a spin, 1. 2. 3.

The airspeed is low and oscillates up and down. The rate of descent (height loss) is steady. The "g" forces remain low.

To recover from the spiral dive you must do the following: 1. 2. 3.

1034 - B

Reduce the power; throttle back. Roll the wings level with coordinated rudder control. Ease out of the dive.

1035 - B

Number: 1036 Question: The Mach trim system will: A. adjust the elevator trim tab, depending on the Mach number. B. pump the fuel from tank to tank, depending on the Mach number. C. adjust the stabiliser, depending on the Mach number. D. keep the Mach number automatically constant. MACH TRIM To guard against nose tuck under, frequent pitch trim changes are required. This is carried out by a variable incidence tailplane, which is automatically positioned by a Mach trim system. This system is designed to aid aircraft longitudinal stability and ensures that the forward stick forces increase proportionally with increasing Mach number. It is operational at high Mach numbers in the transonic speed range.

Number: 1037 Question: The Mach trim system will prevent: A. buffeting. B. shock stall. C. Dutch roll. D. tuck under. Number: 1039 Question: The critical Mach number of an aeroplane is the free stream Mach number that produces the first sign of: By definition Mcrit is "the freestream Mn at which the first point on the A. a shock wave. aerofoil reaches M1.0". B. buffet. At M1.0 you are not supersonic, you are just "sonic" (moving at the same as C. supersonic flow. speed as the LSS) If you were accelerating, a split second later the flow D. local sonic flow. would be supersonic, but at Mcrit it is not yet.

Number: 1040 Question: The critical Mach Number of an aeroplane can be increased by: A. control surface deflection. Wing section design features used to increase MCRIT include: B. dihedral of the wings. » Low thickness/chord ratio C. vortex generators. » Maximum thickness well aft D. sweepback of the wings.

» Small leading edge radius of curvature Wing Planform has the most significant effect on MCRIT. Careful design not only delays the shock stall, but also significantly reduces the severity when it occurs. If a wing has sweep back, the effective chord (parallel to the aircraft's longitudinal axis) is lengthened, but the wing's thickness remains unchanged.

1036 - C

1037 - D

1039 - D

1040 - D

Number: 1042 Question: In transonic flight the ailerons will be less effective than in subsonic flight because: A. aileron deflection only partly affects the pressure distribution around the wing. B. aileron down deflection moves the shock wave forward. C. aileron deflection only affects the air in front of the shock wave. D. behind the shock wave pressure is lower. Because of the effect of the aileron on pressure distribution it will not be felt ahead of the shockwave.

Number: 2345 Question: A jet aeroplane is cruising at high altitude with a Mach number, that provides a buffet margin of 0.3g incremental. In order to increase the buffet margin to 0.4g incremental the pilot must: A. fly at a higher Mach number. margin - Rand, Grenze [Rand] B. extend the flaps to the first selection. buffet - stoß C. fly at a larger angle of attack. D. fly at a lower altitude and the same Mach number. incremental - schrittweise, zunehmend A common characteristic of shock wave induced separation is the increasing severity of buffet intensity with increasing Mach number. In fact, it is possible that the maximum angle of attack may not be achievable due to this severe buffet intensity. The aircraft's manoeuvring capability (load factor) is also reduced.

Number: 2351 Question: The critical Mach number of an aeroplane is the Mach number: A. at which the aeroplane has zero buffet margin. B. at which there is subsonic flow over all parts of the aeroplane. C. at which there is supersonic flow over a part of the aeroplane. D. above which, locally, supersonic flow exists somewhere over the aeroplane.

1042 - A

2345 - D

2351 - D

Number: 2353 Question: The Mach-trim function is installed on most commercial jets in order to minimize the adverse effects of: A. uncontrolled changes in stabiliser setting. MACH TRIM B. increased drag due to shock wave formation. To guard against nose tuck under, frequent pitch trim changes are required. This is carried out by a variable C. compressibility effects on the stabiliser. D. changes in the position of centre of pressure. incidence tailplane, which is automatically positioned by a Mach trim system. This system is designed to aid aircraft longitudinal stability and ensures that the forward stick forces increase proportionally with increasing Mach number. It is operational at high Mach numbers in the transonic speed range.

Number: 2354 Question: When comparing a rectangular wing and a swept back wing of the same wing area and wing loading (assume all other factors of importance remain constant), the swept back wing has the advantage of: Wing section design features A. greater strength. used to increase MCRIT include: B. higher critical Mach number. » Low thickness/chord ratio C. lower stall speed. D. increased longitudinal stability. » Maximum thickness well aft

» Small leading edge radius of curvature Wing Planform has the most significant effect on MCRIT. Careful design not only delays the shock stall, but also significantly reduces the severity when it occurs. If a wing has sweep back, the effective chord (parallel to the aircraft's longitudinal axis) is lengthened, but the wing's thickness remains unchanged.

Number: 2356 Question: Which statement is correct about a normal shock wave? A. The airflow expands when passing the aerofoil. B. The airflow changes direction. C. The airflow changes from supersonic to subsonic. D. The airflow changes from subsonic to supersonic.

Number: 3114 Question: When the air has passed through a normal shock wave the Mach number is: A. lower than before but still greater than 1. B. less than 1. C. higher than before. D. equal to 1.

2353 - D

2354 - B

2356 - C

3114 - B

Number: 3115 Question: When the air is passing through a shock wave the static temperature will: A. decrease and beyond a certain Mach number start increasing again. B. decrease. C. increase. D. stay constant.

Number: 3116 Question: When the air is passing through a shock wave the density will: A. increase. B. decrease and beyond a certain Mach number start increasing again. C. stay constant. D. decrease.

3115 - C

3116 - A

Number: 3117 Question: When air has passed through a shock wave the local speed of sound is: A. increased. B. decreased. C. decreased and beyond a certain Mach number start increasing again. D. not affected.

Number: 3119 Question: The loss of total pressure in a shock wave is due to the fact that: A. the speed reduction is too high. B. the friction in the boundary layer is higher. C. kinetic energy in the flow is converted into heat energy. D. the static pressure decrease is comparatively high.

Number: 3123 Question: Compared with an oblique shock wave at the same Mach number a normal shock wave has a: OBLIQUE SHOCK WAVE A. smaller expansion. An oblique shock wave is a compression wave and is similar to a normal shock B. smaller compression. wave, except that the airflow changes direction into a corner and its velocity C. higher compression. decreases to a lower supersonic value. D. higher expansion. As the air passes through an oblique shock wave its pressure, temperature and density all increase.

3117 - A

3119 - C

3123 - C

Number: 3124 Question: Compared with an oblique shock wave at the same Mach number a normal shock wave has a: A. higher total pressure. B. lower static temperature. C. higher total temperature. D. higher loss in total pressure. You have a loss of velocity (speed) which is dynamic pressure. Total pressure is Dynamic pressure + static pressure The static pressure will recover back to the original, but although the dynamic pressure also recovers a bit, it can never recover back to the original because some of it changed to temperature. Therefore the total pressure must have decreased. We believe there are two correct answers, so that one is worth an appeal. We think the total temperature must also be higher.

Number: 13058 Question: A shock wave on a lift generating wing will: A. reach its highest strength when flying at the critical Mach number. B. be situated at the greatest wing thickness when the aeroplane reaches the speed of sound. C. move slightly aft in front of a downward deflecting aileron. D. move forward as Mach number is increased. 1. The top shock will be at its weakest at Mcrit, it is only just forming 2. The shockwave moves aft as M increases 3. At M1 the shockwave has moved to the trailing edge When the aileron is deflected down the airflow over the top increases and the shockwave moves aft in the increased flow.

3124 - D

13058 - C

Number: 13112 Question: As the Mach number increases in straight and level flight, a shock wave on the upper surface of the wing will: A. move towards the trailing edge. B. not move. C. disappear. D. move towards the leading edge.

Number: 13134 Question: A normal shock wave is a discontinuity plane: A. across which the pressure drops suddenly. B. that is always normal to the surface. C. that is always normal to the local flow. D. across which the temperature drops suddenly.

Number: 13230 Question: In supersonic flight, all disturbances produced by an aeroplane are: A. very weak and negligible. B. outside a conical zone, dependent on the Mach number. C. within a conical zone, dependent on the Mach number. D. in front of the aeroplane. MACH CONE Only the region behind the oblique shock wave is affected by disturbances and is sometimes referred to as the zone of action. The region ahead of the oblique shock wave is not affected by the disturbances and is called the zone of silence. In three dimensions, the disturbances emanating from the moving body expand outward as spheres and not circles. When the speed is above Mach 1, these spheres are enclosed within a cone, called the Mach cone and it is within the Mach cone that disturbances are felt. If the source of the disturbance is a wing, the Mach lines generate two oblique plane waves forming a wedge.

13112 - A

13134 - C

13230 - C

Number: 12973 Question: When the Mach number is slowly increased in straight and level flight the first shock waves will occur: A. on the lower surface of the wing. B. somewhere on the fin. C. somewhere on the horizontal tail. D. on the upper surface at the wing root.

Number: 15601 Question: The sonic boom of an aeroplane flying at supersonic speed is created by: A. the centre of pressure, which is moving aft on the aerofoil. B. aerodynamic heating. C. the expansion flow behind the aeroplane. D. shock waves around the aeroplane.

Number: 15814 Question: What is the effect of aeroplane mass on shock wave intensity at constant Mach number? A. A change in mass does not influence shock wave intensity. B. Decreasing mass increases shock wave intensity at below standard temperatures. C. Decreasing mass increases shock wave intensity. D. Increasing mass increases shock wave intensity.

Number: 15923 Question: Which of these statements about Mcrit is correct? A. Shock waves cannot occur at speeds below Mcrit. B. Mcrit is always greater than 1. C. As speed increases above Mcrit, parasite drag decreases rapidly. D. Flight at any speed above Mcrit causes severe vibration of the aeroplane.

12973 - D

115601 - D

15814 - D

15923 - A

Number: 2069 Question: As altitude increases, the stall speed (IAS): A. initially remains constant but at higher altitudes increases. B. initially remains constant but at higher altitudes decreases. C. remains constant until the tropopause but at higher altitudes increases. D. remains constant regardless of altitude. Vs1g is a constant EAS but the well known compressibility correction means that displayed as an IAS it becomes slowly higher and higher as you climb. At Vs you are at max alpha and the airflow over the wing is peaking at much higher speeds than Mfs. Eventually, at very high altitude shockwave effects begin to downgrade achievable max alpha and Vs1g creeps even higher.

Number: 1673 Question: The critical Mach number of an aerofoil is the free stream Mach number at which: A. a "supersonic bell" appears on the upper surface. B. sonic speed (M=1) is first reached on the upper surface. C. the maximum operating temperature is reached. D. a shock wave appears on the upper surface.

Number: 1352 Question: The regime of flight from the critical Mach number up to approximately M = 1.3 is called the: A. hypersonic range. B. transonic range. C. subsonic range. D. supersonic range.

Number: 1353 Question: Just above the critical Mach number the first evidence of a shock wave will appear at the: A. upper side of the wing. B. lower side of the wing. C. trailing edge of the wing. D. leading edge of the wing.

2069 - A

1673 - B

1352 - B

1353 - A

Number: 1355 Question: Shock induced separation results in: A. decreasing drag. B. decreasing lift. C. increasing lift. D. constant lift.

This is called the shock stall. It differs from a conventional lowspeed stall because it normally occurs at low angles of attack, although with increasing angles of attack the stall will occur at a lower Mach number.

Number: 1356 Question: In the transonic range lift will decrease at the shock stall due to the: A. attachment of the shock wave on the trailing edge of the wing. B. appearance of the bow wave. C. separation of the boundary layer at the shock waves. D. first appearance of a shock wave at the upper side of the wing.

Number: 1357 Question: Is a transport aeroplane allowed to fly at a higher Mach number than the 'buffet-onset' Mach number in 1g flight? A. Yes, this causes no problems. B. Yes, if you want to fly fast at very high altitudes. C. No, this is not acceptable. D. Yes, but only during approach.

1355 - B

1356 - C

1357 - C

Number: 1411 Question: A normal shock wave: A. is a discontinuity plane in an airflow, in which the pressure drops suddenly. B. is a discontinuity plane in an airflow, which is always normal to the surface. C. can occur at different points on the aeroplane in transonic flight. D. is a discontinuity plane in an airflow, in which the temperature drops suddenly.

Number: 3859 Question: Whilst flying at a constant IAS and at n = 1, as the aeroplane mass decreases the value of Mcrit: A. is independent of the angle of attack. B. decreases. C. remains constant. D. increases. If you fly an aircraft at the same speed but at a lower weight, to maintain the value of lift you must reduce angle of attack. At a reduced angle of attack the air is accelerated less over the upper surface of the wing. Therefor the aircraft Mfs can be increased before the airflow over the wing would reach M 1.0 . By definition this is an incresae in Mcrit. Free Stream Mach Number (MFs) is the Mach number of the airflow sufficiently remote from the aircraft so as not to be affected by it, so that:

Number: 3474 Question: What is the influence of decreasing aeroplane weight on Mcrit at constant IAS? A. Mcrit decreases. B. Mcrit decreases as a result of flying at a greater angle of attack. C. Mcrit increases as a result of compressibility effects. D. Mcrit increases as a result of flying at a smaller angle of attack.

1411 - C

3859 - D

3474 - D

Number: 3500 Question: The application of the area rule on aeroplane design will decrease the: A. form drag. B. skin friction drag. C. induced drag. D. wave drag. TRANSONIC AREA RULE The Whitcomb area rule, also called the transonic area rule, is a design technique used to reduce an aircraft's drag at transonic and supersonic speeds, particularly between Mach 0.75 and 1.2. This is one of the most important operating speed ranges for commercial and military fixed-wing aircraft today. Cross-sectional area distribution along the body determines wave drag, largely independent of the actual shape.

Number: 3501 Question: Tuck under will happen: A. only above the critical Mach number. B. above or below the critical Mach number depending on the angle of attack. C. only at the critical Mach number. D. only below the critical Mach number.

3500 - D

3501 - A

Number: 3502 Question: High speed buffet is induced by A. a shift of the centre of gravity. B. expansion waves on the wing upper side. C. boundary layer control. D. boundary layer separation due to shock waves. Number: 3425 Question: To increase the critical Mach number a conventional aerofoil should: A. have a large camber. B. have a low thickness to chord ratio. Wing section design features used to increase MCRIT C. have a large leading edge radius. D. be used with a high angle of attack. include:

» Low thickness/chord ratio

» Maximum thickness well aft » Small leading edge radius of curvature Wing Planform has the most significant effect on MCRIT. Careful design not only delays the shock stall, but also significantly reduces the severity when it occurs. If a wing has sweep back, the effective chord (parallel to the aircraft's longitudinal axis) is lengthened, but the wing's thickness remains unchanged.

Number: 3426 Question: The critical Mach number can be increased by: A. positive dihedral of the wings. B. an increase in wing aspect ratio. C. a T-tail. D. sweepback of the wings.

3502 - D

3425 - B

3426 - D

Number: 3427S Question: Some aeroplanes have a 'waist' or 'coke bottle' contoured fuselage. This is done to: A. fit the engine intakes better to the fuselage. B. apply area rule. C. improve the low speed characteristics. D. increase the strength of the wing root junction.

A "coke bottle fuselage" is needed when a plane is designed to achieve trans-sonic or super-sonic speeds. The area rule for transonic drag minimization was postulated and demonstrated by NACA engineer Richard Whitcomb. The rule was that the total cross-sectional area of an airplane should smoothly increase and then decrease.

At the blackboard in the 1950s, Whitcomb sketches out the principle behind his Transonic Area Rule, which cut drag by reducing variations in a plane's total cross-sectional area

The result was that the fuselage cross section needed to be reduced beginning at the point of attachment of the leading edge of the wing and indented as necessary to compensate for the cross sectional area of the wing at each successive fuselage station going aft. The theory was provided to U.S. aerospace contractors in a confidential Research Memorandum in September 1952.

Number: 14157 Question: What is the relation between the mach angle (mu) and the corresponding mach number ? A. Tan Mu= 1/M B. sin mu = 1/M C. Sin2Mu = 1/M D. sin mu = M

3427 S - B

14157 - B

Number: 14667 Question: Which of these statements about an oblique shock wave are correct or incorrect? I. The static temperature in front of an oblique shock wave is lower than behind it. II. The static pressure in front of an oblique shock wave is lower than behind it. A. B. C. D.

I is correct, II is correct. I is incorrect, II is incorrect. I is correct, II is incorrect. I is incorrect, II is correct

Number: 14668 Question: Which of these statements about an oblique shock wave are correct or incorrect? I. The density in front of an oblique shock wave is lower than behind it. II. The total pressure in front of an oblique shock wave is higher than behind it. A. B. C. D.

I is incorrect, II is incorrect. I is correct, II is incorrect. I is correct, II is correct. I is incorrect, II is correct.

The question refers to total pressure, Static plus dynamic. Static pressure rises in the oblique shock wave and will still be higher behind the shock wave than in front of it. However the dynamic reduces as velocity decreases through the shock wave and pressure energy due to the movement of the air is converted to heat energy in the shock wave. The result is TOTAL PRESSURE goes down behind a shock wave.

14667 - A

14668 - C

Number: 14669 Question: Which of these statements about an oblique shock wave are correct or incorrect? I. The local speed of sound in front of an oblique shock wave is higher than behind it. II. The Mach number in front of an oblique shock wave is lower than behind it. A. B. C. D.

I is incorrect, II is incorrect. I is incorrect, II is correct. I is correct, II is incorrect. I is correct, II is correct.

Number: 14670 Question: Which of these statements about an oblique shock wave are correct or incorrect? I. The density behind an oblique shock wave is higher than in front of it. II. The local speed of sound behind an oblique shock wave is higher than in front of it. A. B. C. D.

I is incorrect, II is correct. I is incorrect, II is incorrect. I is correct, II is incorrect. I is correct, II is correct.

Number: 14671 Question: Which of these statements about an oblique shock wave are correct or incorrect? I. The Mach number behind an oblique shock wave is lower than in front of it. II. The total pressure behind an oblique shock wave is lower than in front of it. A. B. C. D.

I is incorrect, II is correct. I is incorrect, II is incorrect. I is correct, II is incorrect. I is correct, II is correct.

The question refers to total pressure, Static plus dynamic. Static pressure rises in the oblique shock wave and will still be higher behind the shock wave than in front of it. However the dynamic reduces as velocity decreases through the shock wave and pressure energy due to the movement of the air is converted to heat energy in the shock wave. The result is TOTAL PRESSURE goes down behind a shock wave.

14669 - A

14670 - D

14671 - D

Number: 14672 Question: Which of these statements about an oblique shock wave are correct or incorrect? I. The static temperature behind an oblique shock wave is higher than in front of it. II. The static pressure behind an oblique shock wave is higher than in front of it. A. B. C. D.

I is correct, II is correct. I is correct, II is incorrect. I is incorrect, II is correct. I is incorrect, II is incorrect.

Number: 15344 Question: What is the value of the Mach number if the Mach angle equals 45°? A. 1.4 B. 2.0 C. 1.2 D. 0.7

Number: 15179 Question: When supersonic airflow passes through an oblique shock wave, how do (1) static pressure, (2) density, and (3) local speed of sound change? A. B. C. D.

(1) remains constant, (2) decreases, (3) increases. (1) increases, (2) increases, (3) decreases. (1) decreases, (2) decreases, (3) decreases. (1) increases, (2) increases, (3) increases.

14672 - A

15344 - A

15179 - D

02-00 - TRANSONIC AERODYNAMICS

PRINCIPLES OF FLIGHT

02-03 - Means to avoid the effects of exceeding Mcrit (23 Fragen)

Number: 16495 Question: What is the effect of exceeding Mcrit on the stick force stability of an aeroplane with sweptback wings without any form of stability augmentation? A. A decrease, due to loss of lift in the wing root area. B. No effect, because stick force stability is independent of Mach number. C. No effect, because Mcrit is not relevant when considering stick force stability. D. An increase, due to shock wave formation in the wing root area.

Number: 16527 Question: When shock stall occurs, lift will decrease because: A. of the existence of a shock wave being located at the trailing edge of the wing. B. the bow wave appears. C. Mcrit is reached. D. flow separation occurs behind the shock wave.

Number: 16663 Question: Which of these statements on shock stall is correct? A. CLmax does not change as the Mach number increases. B. Shock stall is a stall due to flow separation caused by a shock wave. C. Shock stall is a stall due to flow separation at high angles of attack. D. Shock stall is caused by sudden loss of lift due to a rise in load factor.

16495 - A

16527 - D

16663 - B

Number: 16618 Question: If IAS remains constant, the effect of decreasing aeroplane mass is that Mcrit: A. remains unchanged. B. decreases as altitude increases and increases as altitude decreases. C. decreases, assuming the temperature remains constant. D. increases. Number: 1028 Question: Vortex generators on the upper side of the wing surface will: A. increase the critical Mach number. B. decrease the span wise flow at high Mach numbers. C. increase the magnitude of the shock wave. D. decrease shock wave induced flow separation. VORTEX GENERATORS These are small wing-like surfaces, which are fitted in front of the control surface and project vertically upward into the airstream. They operate by forcing high-energy air into the boundary layer, enabling it to overcome the adverse pressure gradient caused by the shock wave, and delaying its separation.

Number: 1029 Question: Vortex generators on the upper side of the wing: A. decrease critical Mach number. B. decrease wave drag. C. increase wave drag. D. increase critical Mach number. Number: 13113 Question: As altitude increased, the stall speed (IAS): A. initially remains constant and at higher altitudes decreases. B. remains constant regardless of altitude. C. remains constant until the tropopause and at higher altitudes increases. D. initially remains constant and at higher altitudes increases. Number: 13132 Question: During which type of stall does the angle of attack have the smallest value? A. Deep stall. B. Low speed stall. C. Shock stall. D. Accelerated stall. Shock Stall is when the airflow over the wing becomes supersonic; the resulting shockwave causes flow separation... loss of lift.

Number: 13156 Question: An aeroplane should be equipped with a Mach trimmer, if: A. at transonic Mach numbers the aeroplane demonstrates unconventional elevator stick force characteristics. B. stick force stability is independent of the airspeed and -altitude. C. at high airspeed and low altitude the aeroplane demonstrates unconventional elevator stick force characteristics. D. stick force per g strongly decreases at low Mach numbers.

16618 - D

1028 - D

1029 - B

13113 - D

13132 - C

13156 - A

Number: 13164 Question: An aeroplane is flying through the transonic range whilst maintaining straight and level flight. As the Mach number increases the centre of pressure of the wing will move aft. This movement requires: A. a higher IAS to compensate the nose down effect. B. a stability augmentation system to improve dynamic stability. C. a pitch up input of the stabiliser. D. much more thrust from the engines.

Number: 13232 Question: Mach buffet occurs: A. when the stall angle of attack is exceeded. B. when the Mach number has increased to Mcrit. C. at the Mach number at which shock wave induced boundary layer separation occurs. D. directly after exceeding Mcrit.

Number: 13239 Question: In the event of failure of the Mach trimmer: A. the Mach number must be limited. B. the aeroplane mass must be limited. C. the speed must be kept constant. D. the centre of gravity must be moved aft.

Number: 12970 Question: When an aerofoil section has accelerated from subsonic to supersonic speeds, its aerodynamic centre will have: A. shifted aft by about 10%. B. shifted slightly forward. C. shifted from approximately 25% to about 50% of the chord. D. not moved.

Number: 15652 Question: The movement of the aerodynamic centre of the wing when an aeroplane accelerates through the transonic range causes: As you move from subsonic to transonic the movement of the A. a decrease in static longitudinal stability. AC to 50% chord increases the wing contribution and B. a decrease in static directional stability. improves static longitudinal stability as the nosedown pitching C. an increase in static directional stability. effect about the CG is increased. D. an increase in static longitudinal stability. This may however lead to high trim drag or greater control effectiveness if a reduction in manoeuvrability is to be avoided. The downwash acting on the tailplane also results in an increased effective angle of attack.

Number: 15637 Question: What will happen if a large transport aeroplane slowly decelerates in level flight from its cruise speed in still air at high altitude? A. High speed buffet. B. CLMAX will increase due to shock stall. C. An accelerated stall. D. Stick shaker activation or low speed buffeting.

13164 - C

13232 - C

13239 - A

12970 - C

15652 - D

15637 - D

Number: 15694 Question: Shock induced separation can occur: A. behind a strong normal shock wave, independent of angle of attack. B. in supersonic flow only. C. at high Mach numbers, only at low angles of attack. D. at high Mach numbers, only at high angles of attack.

Shock Stall is when the airflow over the wing becomes supersonic; the resulting shockwave causes flow separation... loss of lift.

Number: 15801 Question: When altitude increases, the stall speed (IAS) will: A. increase due to decreasing Mcrit. B. decrease due to decreasing temperature and decreasing Mach number. C. increase due to increasing compressibility effects as a result of increasing Mach number. D. decrease due to increasing Mcrit. Number: 15947 Question: The position of the centre of pressure on an aerofoil of an aeroplane cruising at supersonic speed when compared with that at subsonic speed is: A. further aft. B. undetermined since there is no defined centre of pressure location at supersonic speeds. C. identical. D. further forward.

15694 - A

15801 - C

15947 - A

Number: 15882 Question: The increase in stall speed (IAS) with increasing altitude is due to: A. an increase in TAS. B. the larger angle of attack necessary in lower-density air to obtain the same lift as at sea level. C. exceedance of Mcrit. D. compressibility effects.

Number: 3499 Question: Vortex generators mounted on the upper wing surface will: A. decrease the stall speed by increasing spanwise flow on the wing. B. increase the effectiveness of the spoiler due to increase in parasite drag. C. decrease the interference drag of the trailing edge flaps. D. decrease the shock wave induced separation. VORTEX GENERATORS These are small wing-like surfaces, which are fitted in front of the control surface and project vertically upward into the airstream. They operate by forcing high-energy air into the boundary layer, enabling it to overcome the adverse pressure gradient caused by the shock wave, and delaying its separation.

Number: 14304 Question: Which type of buffet will occur if a jet aeroplane slowly accelerates in level flight from its cruise speed in still air at high altitude? A. Mach buffet. B. Low speed buffet. C. Flutter speed buffet. D. Accelerated stall buffet. Number: 15012 Question: What is the highest speed possible without supersonic flow over the wing? A. Initial Mach buffet speed. B. M = 1. C. Critical Mach number. D. Tuck under speed.

Number: 15109 Question: When the speed over an aerofoil section increases from subsonic to supersonic, its aerodynamic centre: A. moves from approximately 25% to about 50% of the chord. B. does not move. C. moves slightly forward. D. moves aft by approximately 10% of the chord.

15882 - D

3499 - D

14304 - A

15012 - C

15109 - A

02-00 - TRANSONIC AERODYNAMICS

PRINCIPLES OF FLIGHT

02-02 New Questions - (12 Fragen)

Number: 16847 Question: From the buffet onset graph of a given jet transport aeroplane it is determined that at FL 310 at a given mass buffet free flight is possible between M = 0.74 and M = 0.88. In what way would these numbers change if the aeroplane is suddenly pulled up e.g. in a traffic avoidance manoeuvre? A. Both Mach numbers increase. B. The lower Mach number increases and the higher Mach number decreases. C. Both Mach numbers decrease. D. The lower Mach number decreases and the higher Mach number increases.

16847 - B

Number: 10131 Question: The speed range between high- and low speed buffet: A. increases during a descent at a constant IAS. B. decreases during a descent at a constant Mach number. C. is always positive at Mach numbers below MMO. D. increases during climb.

Number: 12950 Question: What is the significance of the maximum allowed cruising altitude, based on the 1.3 g margin? At this altitude: A. exceeding a load factor of 1.3 will cause permanent deformation of this aeroplane. B. a manoeuvre with a load factor of 1.3 will cause buffet onset. C. a manoeuvre with a load factor of 1.3 will cause a Mach number at which accelerated low speed stall occurs. D. a manoeuvre with a load factor of 1.3 will cause Mcrit to be exceeded.

Number: 4566 Question: In order to provide an adequate "buffet boundary" at the commencement of the cruise a speed of 1.3Vs is used. At a mass of 120000 kg this is a CAS of 180 knots. If the mass of the aeroplane is increased to 135000 kg the value of 1.3Vs will be A. unaffected as Vs always occurs at the same angle of attack. B. increased to 191 knots, drag will decrease and air distance per kg of fuel will increase. C. increased to 191 knots, drag will increase and air distance per kg of fuel will decrease. D. increased to 202 knots but, since the same angle of attack is used, drag and range will remain the same. The stall speed increases as the square root of the weight change, not just in proportion. Thus a doubling of weight increases the stall speed not by a factor of 2 but by 1.414. In this case the difference is 135 ÷ 120 = 1.125 and the stall speed increases by the square root of that, 1.061.

Question NOT USED in UK exams

1031 - A

12950 - B

4566 - C

Number: 16096 Question: Which of these statements about wing sweepback are correct or incorrect? I. Increasing wing sweepback increases Mcrit. II. Increasing wing sweepback increases the drag divergence Mach number. A. I is incorrect, II is correct. B. I is correct, II is incorrect. C. I is incorrect, II is incorrect. D. I is correct, II is correct.

If a wing has sweep back, the effective chord (parallel to the aircraft's longitudinal axis) is lengthened, but the wing's thickness remains unchanged. This reduces the thickness/chord ratio of the wing, which results in a higher value of Mcrit, and delays the transonic drag rise.

Disadvantages 1) A reduction in the coefficient of lift, which increases stall speed. 2) Wing tip stalling leading to pitch-up 3) Low aspect ratio leading to increased induced drag at high angles of attack, which is particularly dangerous during takeoff and landing.

Number: 14666 Question: Which of these statements about wing sweepback are correct or incorrect? I. Increasing wing sweepback decreases Mcrit. II. Increasing wing sweepback increases the drag divergence Mach number. A. I is incorrect, II is correct. B. I is correct, II is incorrect. C. I is correct, II is correct. D. I is incorrect, II is incorrect.

16096 - D

114666 - A

Number: 13057 Question: Which statement concerning sweepback is correct? A. Sweepback provides a positive contribution to static lateral stability. B. A disadvantage of sweepback is that it decreases Mcrit. C. Sweepback increases speed stability at Mach numbers above Mcrit. D. Sweepback is mainly intended to increase static directional stability.

Number: 15629 Question: A supercritical wing: A. will be free of Mach buffet in the transonic range. B. will develop no transonic flow just above Mcrit. C. will develop no noticeable shock waves when flying just above Mcrit. D. always cruises at Mcrit. SUPERCRITICAL WINGS To reduce the severity of the shock stall and allow aircraft to travel faster, some modern jet transport category aircraft have supercritical wings. The point of maximum thickness is positioned close to the trailing edge and the upper surface has a very slight curvature. This ensures that the localised Mach number remains just above the critical Mach number and results in a fiattish pressure distribution over the majority of the upper surface.

Number: 15570 Question: In comparison to a conventional aerofoil section, typical shape characteristics of a supercritical aerofoil section are: A. a larger nose radius, flatter lower surface and negative as well as positive camber. B. a larger nose radius, flatter upper surface and negative as well as positive camber. C. a sharper pointed nose, flatter lower surface and positive camber at the rear of the aerofoil section. D. a sharper pointed nose, negative camber and a flatter upper surface. Number: 15549 Question: One advantage of a supercritical wing aerofoil over a conventional one is: A. it allows a wing of increased relative thickness to be used for approximately the same cruise Mach number. B. that there is no need for spoilers. C. a lower value of Mcrit at the same relative thickness. D. improved Dutch roll damping at cruise altitude. Number: 15527 Question: The effect of increasing angle of sweep is: A. a decrease in the critical Mach number. B. a decrease in stall speed. C. an increase in longitudinal stability. D. an increase in the critical Mach number.

13057 - A

15629 - C

15570 - B

15549 - A

15527 - D

Number: 16723 Question: Mcrit is increased by: A. dihedral, thin aerofoils and supercritical aerofoil sections. B. sweepback, thin aerofoils and area ruling. C. sweepback, area ruling and high aspect ratio. D. sweepback, dihedral and thin aerofoils.

A "coke bottle fuselage" is needed when a plane is designed to achieve trans-sonic or super-sonic speeds. The area rule for transonic drag minimization was postulated and demonstrated by NACA engineer Richard Whitcomb. The rule was that the total cross-sectional area of an airplane should smoothly increase and then decrease. The result was that the fuselage cross section needed to be reduced beginning at the point of attachment of the leading edge of the wing and indented as necessary to compensate for the cross sectional area of the wing at each successive fuselage station going aft. The theory was provided to U.S. aerospace contractors in a confidential Research Memorandum in September 1952.

6723 - B

At the blackboard in the 1950s, Whitcomb sketches out the principle behind his Transonic Area Rule, which cut drag by reducing variations in a plane's total cross-sectional area

03-00 - SUPERSONIC AERODYNAMICS

PRINCIPLES OF FLIGHT

03-00 Oblique Shock waves - (20 Fragen)

Number: 16327 Question: How does the Mach number change during a climb at constant IAS from sea level to 40,000 ft? A. B. C. D.

Decreases with increasing altitude. Initially remains constant until approximately 25,000 ft and then increases with increasing altitude. Increases with increasing altitude. Initially remains constant until approximately 25,000 ft and then decreases with increasing altitude.

Number: 16331 Question: Which statement with respect to the speed of sound is correct? A. Doubles if the temperature increases from 9° to 36° Centigrade. B. Increases always if the density of the air decreases. C. Is independent of altitude. D. Varies with the square root of the absolute temperature.

16327 - C;

16331 - D;

Number: 16275 Question: Assuming ISA conditions and a descent below the tropopause at constant Mach number and aeroplane mass, the: A. TAS remains constant. B. lift coefficient decreases. C. lift coefficient increases. D. IAS decreases.

Number: 16757 Question: Which of these statements about the transonic speed range is correct ? A. The airflow everywhere around the aeroplane is supersonic. B. The transonic speed range starts at Mcrit and extends to Mach numbers above M = 1. C. The transonic speed range starts at M = 0.5 and ends at Mcrit. D. The airflow everywhere around the aeroplane is subsonic.

16275 - B;

16757 - B;

Number: 1043 Question: Compressibility effects depend on: A. TAS. B. IAS. Right back at the beginning we said we would treat air as an incompressible fluid. C. Mach Number. Unfortunately it is not, and at higher Mach numbers we begin to see the effects of D. EAS. our mistake, like the difference between CAS and EAS followed by the formation of pressure waves and shockwave. These all depend on Mach number, although we often cite 300kt as the point at which we first notice things going wrong it is in fact M0.4

Number: 1044 Question: The formula for the Mach Number is: (a= speed of sound) A. M= a / TAS B. M= IAS / a C. M= TAS * a D. M= TAS / a

Number: 1045 Question: If the altitude is increased and the TAS remains constant in the troposphere under standard atmospheric conditions, the Mach number will: A. decrease. B. increase or decrease, depending on the type of aeroplane. C. increase. D. not change.

1043 - C;

1044 - D;

1045 - C;

Number: 1046 Question: Assuming ISA conditions, climbing at a constant Mach Number up the tropopause the TAS will: A. first increase, then decrease. B. increase. C. remain constant. D. decrease.

Number: 1047 Question: The speed of sound is determined only by: A. humidity. B. pressure. C. temperature. D. density.

Number: 2344 Question: An aeroplane is descending at a constant Mach number from FL 350. What is the effect on true airspeed? A. It remains constant. B. It increases as temperature increases. C. It decreases as altitude decreases. D. It decreases as pressure increases.

1046 - D;

1047 - C;

2344 - B;

Number: 3113 Question: The flight Mach number is 0.8 and the TAS is 400 kts. The speed of sound is: A. 600 kts B. 500 kts C. 320 kts D. 480 kts Number: 13100 Question: During a climb at a constant IAS, the Mach number will: A. increase initially and remain constant subsequently. B. decrease initially and increase subsequently. C. increase. D. remain constant.

Number: 13213 Question: if the Mach number is 0.8 and the TAS is 400 kt, what is the speed of sound? A. 500 kt. B. 320 kt. C. 600 kt. D. 480 kt.

Number: 13249 Question: In the transonic range the aeroplane characteristics are strongly determined by the: A. TAS. B. CAS. C. IAS. D. Mach number.

Number: 12995 Question: The Mach number is the ratio between the: A. TAS of the aeroplane and the speed of sound at sea level. B. IAS of the aeroplane and the speed of sound at sea level. C. IAS of the aeroplane and the speed of sound of the undisturbed flow. D. TAS of the aeroplane and speed of sound of the undisturbed flow.

3113 - B;

13100 - C;

13213 - A;

13249 - D;

12995 - D

Number: 13007 Question: A transonic Mach number is a Mach number: A. in the range between Mcrit and MMO. B. at which only supersonic local speeds occur. C. between the buffet onset Mach number and the Mach number corresponding to total supersonic flow. D. at which both subsonic and supersonic local speeds occur.

Number: 15426 Question: if the Mach number is 0.8 and the TAS is 480 kt, what is the speed of sound? A. 600 kt. B. 560 kt. C. 750 kt. D. 384 kt. Number: 1410 Question: The Mach number is the ratio between the: A. IAS of the aeroplane and the speed of sound of the undisturbed flow. B. IAS of the aeroplane and the speed of sound at sea level. C. TAS of the aeroplane and the speed of sound of the undisturbed flow. D. TAS of the aeroplane and the speed of sound at sea level. Number: 3875 Question: Which statement with respect to the speed of sound is correct? A. Doubles if the temperature increases from 9° to 36° Centigrade. B. Varies with the square root of the absolute temperature. C. Increases always if the density of the air decreases. D. Is independent of altitude.

13007 - D;

15426- B;

1410 - C;

3875 - B;

Number: 14578 Question: Regarding the transonic speed range: A. it starts at a Mach number just above 1. B. it starts at Mcrit and extends to Mach number equal to 1 C. it starts at a Mach number equal to 1. D. both subsonic and supersonic speeds exist in the flow around the aeroplane.

Number: 14579 Question: The subsonic speed range: A. implies both subsonic and supersonic speeds exist in the flow around the aeroplane. B. ends at M = 1. C. ends at Mcrit. D. ends at a Mach number just above M = 1.

14578 - D;

14579 - C;

Number: 14580 Question: Which of these statements about the supersonic speed range is correct? A. The supersonic speed range starts at a Mach number below M = 1 and extends to Mach numbers above M = 1. B. The airflow everywhere around the aeroplane is supersonic. C. The supersonic speed range starts at M = 1 and ends at Mcrit. D. The airflow around the aeroplane is transonic.

Number: 15003 Question: The speed range from approximately M=1.3 to approximately M=5 is called the: A. transonic range. B. supersonic range. C. subsonic range. D. hypersonic range.

14580 - B;

15003 - B;

05-01 - CONTROL (8 Fragen)

Number: 16809 Question: Which statement about elevators is correct? A. The elevator is the primary control surface for control about the longitudinal axis and is operated by a forward or backward movement of the control wheel or stick. B. The elevator is the primary control surface for control about the lateral axis and is operated by a forward or backward movement of the control wheel or stick. C. The elevator is used only to trim an aeroplane and is normally operated by a dedicated control wheel, which is usually situated close to the throttle. D. The elevator is the primary control surface for control about the longitudinal axis and is operated by a left or right rotation of the control wheel. Number: 933 Question: Rotation about the lateral axis is called: A. slipping. B. rolling. C. yawing. D. pitching. Number: 934 Question: Rolling is the rotation of the aeroplane about the: A. wing axis. B. longitudinal axis. C. lateral axis. D. vertical axis. Number: 12990 Question: The pitch angle is defined as the angle between the: A. chord line and the horizontal plane. B. speed vector axis and the longitudinal axis. C. longitudinal axis and the chord line. D. longitudinal axis and the horizontal plane. The angle of incidence is fixed, but the angle of attack changes in flight. Likewise, do not confuse the 'Pitch Angle' or 'Pitch Attitude' of the aircraft with the angle of attack. For any given angle of attack, the pitch angle can vary. Similarly for any given pitch angle, the angle of attack can also vary.

16809 - B

933 - D

934 - B

12990 - D

Number: 14574 Question: An aeroplane's pitch angle is defined as the angle between its: A. longitudinal axis and the horizontal plane. B. lateral axis and the horizontal plane. C. speed vector and its longitudinal axis. D. speed vector and the horizontal plane. The angle of incidence is fixed, but the angle of attack changes in flight. Likewise, do not confuse the 'Pitch Angle' or 'Pitch Attitude' of the aircraft with the angle of attack. For any given angle of attack, the pitch angle can vary. Similarly for any given pitch angle, the angle of attack can also vary.

Number: 14586 Question: An aeroplane's bank angle is defined as the angle between its: A. lateral axis and the horizontal plane. B. speed vector and the horizontal plane. C. longitudinal axis and the horizontal plane. D. speed vector and longitudinal axis.

Number: 14634 Question: Rotation around the longitudinal axis is called: A. pitching. B. yawing. C. slipping. D. rolling.

Number: 14635 Question: Rotation around the normal axis is called: A. rolling. B. slipping. C. yawing. D. pitching.

14574 - A

14586 - A

14634 - D

14635 - C

05-08 - CONTROL (23 Fragen)

Number: 2361 Question: One advantage of a movable-stabiliser system compared with an elevator trim system is that: A. the system's complexity is reduced. B. it is a more effective means of trimming. C. it leads to greater stability in flight. D. the complete system (structure and control mechanism) weighs less.

Number: 13075 Question: Which statement in respect of a trimmable horizontal stabiliser is correct? A. An aeroplane with a forward cg requires the stabiliser leading edge to be higher than for one with an aft cg in the same trimmed condition. B. An aeroplane with a forward cg requires the stabiliser leading edge to be lower than for one with an aft cg in the same trimmed condition. C. At the forward C.G. limit , stabiliser trim is adjusted fully nose down to obtain maximum elevator authority at rotation during take-off. D. Because take-off speeds do not vary with centre of gravity location, the need for stabiliser adjustment is dependent on flap position only. If the C.G is moved aft it results in a larger nose down pitching moment which has to be compensated for by placing a download on the tail plane. This can be achieved by deflecting the elevator up, although this results in high trim drag at higher airspeeds. On large aircraft small movements of the variable incidence stabilizer allow the trimming to be carried out, but with reduced trim drag. An upward deflection of the elevator is required to trim the aircraft, so think of the stabilizer as a large elevator with its trailing edge moving in the same direction as the elevator. To achieve this you must therefore lower the leading edge of the stabilizer to produce the necessary download. Note: Once in a trimmed condition on fully powered flying controls the elevator will then align itself with the stabilizer to reduce the tail load and also the loads acting on the hinges and servo actuator. On power assisted controls there is however no guarantee that this occurs due to downwash and configuration effects that can alter the flow over the surface. It is also fair to say that trim drag also increases fuel consumption.

Number: 13214 Question: If the elevator trim tab is deflected up, the cockpit trim indicator shows: A. nose down. B. nose left. C. nose up. D. neutral.

2361 - B

13075 - B

13214 - A

Number: 13241 Question: In straight and level flight, as speed is reduced: A. the elevator is deflected further downwards and the trim tab further upwards. B. both elevator and trim tab are deflected further upwards. C. the elevator and trim tab do not move. D. the elevator is deflected further upwards and the trim tab further downwards.

Number: 12954 Question: What is the effect of elevator trim tab adjustment on the static longitudinal stability of an aeroplane? A. Aeroplane nose down trim increases the static longitudinal stability. B. Aeroplane nose up trim increases the static longitudinal stability. C. Depends on the value of stick force/g. D. No effect.

Number: 15573 Question: When comparing a stabiliser trim system with an elevator trim system, which of these statements is correct? A. an elevator trim is able to compensate larger changes in pitching moments. B. an elevator trim is more suitable for aeroplanes with a large CG range. C. a stabiliser trim is able to compensate larger changes in pitching moments. D. a stabiliser trim is more sensitive to flutter. The Variable Incidence Horizontal Stabiliser provides pitch trim on most transport category aircraft. Varying the angle of incidence has the same effect as moving the elevator, but is aerodynamically more efficient, particularly at high airspeeds, and can provide a considerable trim range.

Number: 15734 Question: In straight flight, as speed is reduced, whilst trimming to keep the stick force zero: A. the elevator and trim tab do not move. B. the elevator is deflected further upwards and the trim tab further downwards. C. the elevator is deflected further downwards and the trim tab further upwards. D. both elevator and trim tab are deflected further upwards.

Number: 1415 Question: If the elevator trim tab is deflected up, the cockpit trim indicator presents: A. nose-up. B. neutral. C. nose-left. D. nose-down.

13241 - D

12954 - D

15573 - C

15734 - B

1415 - D

Number: 3882 Question: What is the position of the elevator in relation to the trimmable horizontal stabiliser of a power assisted aeroplane that is in trim? A. The position depends on speed, the position of slats and flaps and the position of the centre of gravity. B. The elevator is always deflected slightly downwards in order to have sufficient remaining flare capability. C. The elevator deflection (compared with the stabiliser position) is always zero. D. At a forward CG the elevator is deflected upward and at an aft CG the elevator is deflected downward. In the power assisted hydraulic controls you move the control surface then the hydraulics assist you, so you do have feedback from the controls. To give you that feel there will always be some angle between the control surface and the tail plane, except if is at absolute zero. When trimmed you will have no forces on the stick, but the moment you move it out of the trimmed position you must experience the air load on the control surface.

The Variable Incidence Horizontal Stabiliser provides pitch trim on most transport category aircraft. Varying the angle of incidence has the same effect as moving the elevator, but is aerodynamically more efficient, particularly at high airspeeds, and can provide a considerable trim range.

Number: 3887 Question: When a jet transport aeroplane takes off with the CG at the forward limit and the trimmable horizontal stabiliser (THS) is positioned at the maximum allowable nose down position for take-off: A. rotation will require a higher than normal stick force. B. rotation will be normal using the normal rotation technique. C. early nose wheel raising will take place. D. there will be a tendency to over-rotate.

Number: 3476 Question: What should be usually done to perform a landing with the stabiliser jammed in the cruise flight position? A. if possible, relocate as many passengers as possible to the front of the cabin. B. use the Mach trimmer until after landing. C. choose a lower landing speed than normal. D. choose a higher landing speed than normal and/or use a lower flapsetting for landing.

3882 - A

3887 - A

3476 - D

Number: 3477 Question: In general transport aeroplanes with power assisted flight controls are fitted with an adjustable stabiliser instead of trim tabs on the elevator. This is because: A. trim tab deflection increases VMO. B. effectiveness of trim tabs is insufficient for those aeroplanes. C. mechanical adjustment of trim tabs creates too many problems. D. the pilot does not feel the stick forces at all.

Number: 14485 Question: In straight flight, as speed is increased, whilst trimming to keep the stick force zero: A. the elevator is deflected further upwards and the trim tab further downwards. B. both elevator and trim tab are deflected further downwards. C. the elevator and trim tab do not move. D. the elevator is deflected further downwards and the trim tab further upwards.

Number: 14306 Question: The most important factor determining the required position of the Trimmable Horizontal Stabiliser (THS) for take off is the: A. total mass of the aeroplane. B. position of the aeroplane''s centre of gravity. C. centre of gravity position of the fuel. D. stall speed.

Number: 14310 Question: What is the position of the elevator in relation to the trimmable horizontal stabiliser of an aeroplane with fully hydraulically operated flight controls that is in trim? A. At a forward CG, the elevator is deflected upward and at an aft CG, it is deflected downward. B. The position depends on speed, the position of flaps and slats and the position of the centre of gravity. C. Elevator deflection is zero. D. The elevator is always deflected slightly downward in order to have sufficient remaining flare capability.

3477 - B

14485 - D

14306 - B

14310 - C

Number: 14311 Question: What is the effect on landing speed when a trimmable horizontal stabiliser jams at high IAS? A. In most cases, no effect. B. No effect with a forward CG. C. No effect when landing on a high elevation runway. D. In most cases, a higher than normal landing speed is required.

Number: 14597 Question: When a jet transport aeroplane takes off with the CG at the forward limit and the trimmable horizontal stabiliser (THS) is positioned at the maximum allowable nose up position for take-off: A. rotation will require a higher than normal stick force. B. early nose wheel raising will take place. C. rotation will be normal using the normal rotation technique. D. there will be a tendency to over-rotate.

14311 - D

14597 - C

Number: 14598 Question: When a jet transport aeroplane takes off with the CG at the aft limit and the trimmable horizontal stabiliser (THS) is positioned at the maximum allowable nose up position for take-off: A. early nose wheel raising will take place. B. rotation will require higher than normal stick force. C. there will be a tendency to under-rotate. D. rotation will be normal using the normal rotation technique.

Number: 14659 Question: Comparing the differences between a horizontal trimmable stabiliser and an elevator trim tab, which of these statements are correct or incorrect? I. The effects of a trim tab runaway are more serious. II. A jammed trim tab causes less control difficulty. A. B. C. D.

I is incorrect, II is incorrect. I is correct, II is incorrect. I is incorrect, II is correct. I is correct, II is correct.

14598 - A

14659 - C

Number: 14682 Question: Which of these statements about a trimmable horizontal stabiliser is correct? A. At the aft CG limit, stabiliser trim is adjusted fully nose up to obtain maximum elevator authority at rotation during take-off. B. A trimmed aeroplane with an aft CG requires the stabiliser leading edge to be lower than in the case of a forward CG in the same condition. C. Because take-off speeds do not vary with CG position, the need for stabiliser adjustment is dependent on flap position only. D. A trimmed aeroplane with an aft CG requires the stabiliser leading edge to be higher than in the case of a forward CG in the same condition.

Number: 14690 Question: When comparing an elevator trim system with a stabiliser trim system, which of these statements is correct? A. an elevator trim produces lower trim drag B. an elevator trim is able to compensate larger changes in pitching moments. C. an elevator trim is more sensitive to flutter. D. an elevator trim is more suitable for aeroplanes with a large CG range. Number: 14691 Question: When comparing a stabiliser trim system with an elevator trim system, which of these statements is correct? A. an elevator trim is more suitable for aeroplanes with a large CG range. B. a stabiliser trim is not as capable to compensate large changes in pitching moments. C. an elevator trim is able to compensate larger changes in pitching moments. D. a stabiliser trim is less sensitive to flutter.

Number: 14692 Question: When comparing a stabiliser trim system with an elevator trim system, which of these statements is correct? A. a elevator trim is able to compensate larger changes in pitching moments. B. an elevator trim is less suitable for aeroplanes with a large CG range. C. an elevator trim is more suitable for aeroplanes with a large CG range. D. a stabiliser trim is more sensitive to flutter.

14682 - D

14690 - C

14691 - D

14692 - B

05-07 - CONTROL (1 Fragen)

Number: 3888 Question: When flutter damping of control surfaces is obtained by mass balancing, these weights will be located with respect to the hinge of the control surface: A. behind the hinge. B. above the hinge. C. in front of the hinge. D. below the hinge.

3888 - C

05-06 - CONTROL (17 Fragen)

Number: 16469 Question: What is the primary input for an artificial feel system? A. TAS. To prevent over-controlling and overstressing the aircraft, some form of artificial feel is B. Mach number. incorporated in the control system, so C. Static pressure. that the control forces experienced represent those of a manually contro lled aircraft. D. IAS. On transport category aircraft feel forces are provided by - spring units, - Pitot-static Q-feel units, - or a combination of both. Artificial feel systems normally also incorporate a self-centring mechanism, so that the flight deck controls automatically return to their neutral positions when released, centralising the control surfaces.

Number: 1002 Question: Stick forces, provided by an elevator feel system, depend on: A. stabiliser position, static pressure. B. stabiliser position, total pressure. C. elevator deflection, static pressure. D. elevator deflection, dynamic pressure.

Number: 2355 Question: Which kind of ''tab'' is commonly used in case of manual reversion of fully powered flight controls? A. Spring tab. B. Servo tab. C. Anti-balance tab. D. Balance tab. Servo Tab is directly controlled by the pilot through a pivot point and movement of the tab supplies the hinge moment necessary to move the main control surface. Movement of the tab causes the control surface to move to a new position of equilibrium in a direction of travel opposite to that of the tab (i.e. tab down, control surface up). In practice, the servo tab lacks effectiveness at low airspeeds when large control defiections are required. to lack - fehlen This is because the amount of airfiow passing over the tab is too low to produce the necessary hinge moment and hence the required defiection.

Number: 2363 Question: Which statement is correct about a spring tab? A. At high IAS it behaves like a servo tab. B. Its main purpose is to increase stick force per g. C. At low IAS it behaves like a servo tab. D. At high IAS it behaves like a fixed extension of the elevator.

16469 - D

1002 - D

2355 - B

2363 - A

Number: 13087 Question: Which three aerodynamic means decrease manoeuvring stick forces? A. Spring tab - horn balance - bobweight. Horn Balance B. Servo tab - trim tab - balance tab. C. Servo tab - horn balance - spring tab. D. Spring tab - trim tab - mass balancing weight.

Number: 12975 Question: When power assisted controls are used for pitch control: A. they only function in combination with an elevator trim tab. B. aerodynamic balancing of the control surfaces is meaningless. C. trimming is superfluous. D. a part of the aerodynamic forces is still felt on the column.

Number: 15514 Question: The tab in the figure represents: A. a servo tab. B. a balance tab . C. an anti-balance tab. D. a trim tab. Servo Tab is directly controlled by the pilot through a pivot point and movement of the tab supplies the hinge moment necessary to move the main control surface. Movement of the tab causes the control surface to move to a new position of equilibrium in a direction of travel opposite to that of the tab (i.e. tab down, control surface up). In practice, the servo tab lacks effectiveness at low airspeeds when large control defiections are required. to lack - fehlen This is because the amount of airfiow passing over the tab is too low to produce the necessary hinge moment and hence the required defiection.

13087 - C

12975 - D

15514 - A

Number: 15524 Question: The tab in the figure represents: A. a trim tab. B. a control tab. C. an antibalance tab. D. a balance tab that also functions as a trim tab.

Balance Tabs They are connected to the tailplane by a mechanical linkage that causes them to move in the opposite direction to the control surface

Number: 3849 Question: When power assisted controls are used for pitch control: A. they only function in combination with an elevator trim tab. B. trimming is superfluous. C. aerodynamic balancing of the control surfaces is meaningless. D. a part of the aerodynamic forces is still felt on the column.

Number: 3861 Question: Examples of aerodynamic balancing of control surfaces are: A. balance tab, horn balance, and mass balance. B. servo tab, spring tab, seal between the wing trailing edge and the leading edge of control surface. C. spring tab, servo tab, and power assisted control. D. mass in the nose of the control surface, horn balance and mass balance.

15524 - D

3849 - D

3861 - B

Number: 3870 Question: An aeroplane has a servo tab controlled elevator. What will happen if the elevator jams during flight? jams - klemmt A. The pitch control forces double. B. Pitch control is lost. C. Pitch control sense is reversed. D. The servo-tab now works as a negative trim-tab. Only the tab is left functioning. It is not connected to the elevator but works in the unnatural sense to drive the elevator in the natural sense, so with the elevator locked you only have very limited control from the tab only and in the unnatural sense - and, of course, you have no trim or balance function left.

Number: 3880 Question: A horn balance in a control system has the following purpose: A. to obtain mass balancing. B. to decrease the effective longitudinal dihedral of the aeroplane. C. to prevent flutter. D. to decrease stick forces.

Horn Balance

Number: 3657 Question: Examples of aerodynamic balancing of control surfaces are: A. weight in the nose of the control surface, horn balance. B. Fowler flaps, upper and lower rudder. C. upper and lower rudder, seal between wing's trailing edge and leading edge of a control surface. D. seal between wing's trailing edge and leading edge of a control surface, horn balance. The answer would make sense if it said "a seal between the wing trailing edge and the leading edge of the control surface", but even that would be a bit untidy because in a seal balance the seal is in the gap between the wing housing for the control and the leading edge of the control. The use of wing "trailing edge" is confusing

Number: 3654 Question: Which statement about a primary control surface controlled by a servo tab, is correct? A. The servo tab can also be used as a balance tab. B. Due to the effectiveness of the servo tab the control surface area can be smaller. C. The control effectiveness of the primary surface is increased by servo tab deflection. D. The position is undetermined during taxiing, in particular with tailwind. The function of a servo tab is very different from a balance tab. With a servo tab control system movement of the pilot’s flight controls moves the servo tab. The servo tab at the trailing edge of the main flying control surface produces a aerodynamic force to move the control surface. The servo tab is displaced in the opposite direction in which the flight control surface moves. ie if you wish to pitch the aircraft nose up, servo tab is deflected down and moves the elevator up. The system requires airflow from leading edge to trailing edge, when taxiing in a tailwind the effectiveness of this type of control is reduced.

3870 - C

3880 - D

3657 - D

3654 - D

Number: 14857 Question: Which aerodynamic design features can be used to reduce control forces? A. Balance tab, control surfaces with increased area behind the hinge, artificial feel system. B. Horn balance, balance tab, servo tab. C. Servo tab, bobweight, control surfaces with increased area. D. Mass balance, horn balance, artificial feel system.

Number: 14858 Question: In general, control forces are reduced by: A. a servo tab, spring tab and mass balancing. B. a horn balance, servo tab and spring tab. C. mass balancing, a trim tab and spring tab. D. a balance tab, forward shift of the CG and a servo tab.

Number: 15068 Question: Artificial feel is required: A. when the flight control surfaces are fitted with control tabs or trim tabs. B. with power assisted flight controls. C. with fully powered flight controls. D. when there is a trimmable stabiliser.

14857 - B

14858 - B

15068 - C

05-05 - CONTROL (5 Fragen)

Number: 16637 Question: Yaw is followed by roll because the: A. rolling motion generated by rudder deflection causes a speed increase of the outer wing which increases the lift on that wing so the aeroplane starts to roll in the direction of the turn. B. yawing motion generated by rudder deflection causes a speed increase of the inner wing, which increase the lift on that wing so that the aeroplane starts to roll in the same direction as the yaw. C. rudder is located above the longitudinal axis and when it is deflected, it causes a rolling moment in the same direction as the yaw. D. yawing motion generated by rudder deflection causes a speed increase of the outer wing, which increases the lift on that wing so that the aeroplane starts to roll in the same direction as the yaw.

Number: 16726 Question: Rotation about the longitudinal axis of an aeroplane can be achieved by: A. aileron deflection and/or rudder deflection. B. speed brake extension or wing flap deflection. C. elevator deflection and/or slat extension. D. symmetrical spoiler deflection and/or elevator deflection.

Number: 178 Question: If the nose of an aeroplane yaws left, this causes: A. a roll to the right. B. an increase in lift on the left wing. C. a roll to the left. D. a decrease in relative airspeed on the right wing.

Number: 961 Question: Which moments or motions interact in a dutch roll? A. Pitching and adverse yaw. B. Pitching and yawing. C. Pitching and rolling. D. Rolling and yawing.

Number: 13212 Question: If the nose of an aeroplane yaws left, this causes: A. a decrease in relative airspeed on the right wing. B. a roll to the right. C. a roll to the left. D. an increase in lift on the left wing.

16637 - D

16726 - A

178 - C

961 - D

13212 - C

05-04 - CONTROL (19 Fragen) Number: 16415 Question: The function of ailerons is to rotate the aeroplane about the: A. yaw axis. B. lateral axis. C. normal axis. D. longitudinal axis. Number: 16303 Question: Outboard ailerons (if present) are normally used: A. in low speed flight only. B. when the landing gear is up. C. in high speed flight only. D. at transonic and supersonic speeds only. On aircraft fitted with two sets of ailerons, (inboard and outboard), as the airspeed increases the aerodynamic loads on the ailerons tend to twist the wing at the tips, where it is more fiexible. To overcome this tendency, some aircraft use the technique of locking the outboard ailerons in the faired or neutral position, and use an inboard aileron/spoiler combination above fiap retraction speeds to provide the necessary roll control.

Number: 1001 Question: During initiation of a turn with speedbrakes extended, the roll spoiler function induces a spoiler deflection: A. downward on the upgoing wing and upward on the downgoing wing. B. on the downgoing wing only. C. upward on the upgoing wing and downward on the downgoing wing. D. on the upgoing wing only.

16415 - D

16303 - A

1001 - A

Number: 2364 Question: How is adverse yaw compensated for during entry into and roll out from a turn? A. Differential aileron deflection. B. Anti-balanced rudder control. C. Servo tabs. D. Horn-balanced controls.

Number: 3109 Question: One method to compensate adverse yaw is: A. an anti-balance tab. B. a differential aileron. C. a balance panel. D. a balance tab.

Number: 3110 Question: Flaperons are controls which combine the function of: A. flaps and elevator. Some commercial aircraft have ailerons that droop when flaps are extended to B. ailerons and elevator. increase the lift co-efficient, these are flaperons. C. ailerons and flaps. D. flaps and speed brakes. There are some other aircraft (mainly military) that have split ailerons that can act as speed brakes as well, but these are not called flaperons.

Number: 13125 Question: An example of differential aileron deflection during initiation of left turn is: A. Left aileron: 5° up. Right aileron: 2° down. B. Left aileron: 2° down. Right aileron: 5° up. C. Left aileron: 2° up. Right aileron: 5° down. D. Left aileron: 5° down. Right aileron: 2° up.

Number: 13162 Question: An aeroplane is provided with spoilers and both inboard and outboard ailerons. Roll control during cruise is provided by: On aircraft fitted with two sets of ailerons, (inboard and outboard), as the A. inboard ailerons and roll spoilers. airspeed increases the aerodynamic loads on the ailerons tend to twist the B. inboard and outboard ailerons. wing at the tips, where it is more fiexible. C. outboard ailerons only. To overcome this tendency, some aircraft use the technique of locking the D. outboard ailerons and roll spoilers. outboard ailerons in the faired or neutral position, and use an inboard aileron/spoiler combination above fiap retraction speeds to provide the necessary roll control.

2364 - A

3109 - B

3110 - C

13125 - A

13162 - A

Number: 13236 Question: In what phase of flight are the outboard ailerons (if fitted) not active? A. Landing with a strong and gusty crosswind, to avoid over-controlling the aeroplane. B. Cruise. C. Take-off, until lift-off. D. Approach. On aircraft fitted with two sets of ailerons, (inboard and outboard), as the airspeed increases the aerodynamic loads on the ailerons tend to twist the wing at the tips, where it is more fiexible. To overcome this tendency, some aircraft use the technique of locking the outboard ailerons in the faired or neutral position, and use an inboard aileron/spoiler combination above fiap retraction speeds to provide the necessary roll control.

Number: 12974 Question: When roll spoilers are extended, the part of the wing on which they are mounted: A. experiences a reduction in lift, which generates the desired rolling moment. In addition there is a local increase in drag, which suppresses adverse yaw. B. is forced downwards as a reaction to the increased drag. C. stalls. This causes a difference in lift between both wings, which generates the desired rolling moment. D. experiences extra drag, which generates a yawing moment. The speed difference between both wings generates the desired rolling moment.

Number: 15608 Question: Aileron deflection causes a rotation around the longitudinal axis by: A. changing the wing drag and the two wings therefore produce different lift values resulting in a moment about the longitudinal axis. B. aileron secondary effect. C. causing sideslip, which generates a rolling moment. D. changing the wing camber and the two wings therefore produce different lift values resulting in a moment about the longitudinal axis.

13236 - B

12974 - A

15608 - D

Number: 15554 Question: When a turn is initiated, adverse yaw is: A. a momentary yawing motion opposite to the turn due to an incorrect differential aileron movement. B. the tendency of an aeroplane to yaw in the opposite direction of turn mainly due to the difference in induced drag on each wing. C. the tendency of an aeroplane to yaw in the same direction of turn due to the different wing speeds. D. the tendency of an aeroplane to yaw in the opposite direction of turn mainly due to the difference in aileron form drag.

Number: 1416 Question: Differential aileron deflection: A. increases the CLmax. B. is required to achieve the required roll rate. C. equals the drag of the right and left aileron. D. is required to keep the total lift constant when ailerons are deflected.

Number: 1417 Question: An example of differential aileron deflection during initiation of left turn is: A. Left aileron: 5° up. Right aileron: 2° down. B. Left aileron: 2° down. Right aileron: 5° up. C. Left aileron: 2° up. Right aileron: 5° down. D. Left aileron: 5° down. Right aileron: 2° up.

Number: 3869 Question: Which phenomenon is counteracted with differential aileron deflection? A. Adverse yaw. B. Turn co-ordination. counteracted - entgegengearbeitet , entgegengewirkt C. Sensitivity for spiral dive. D. Aileron reversal.

Number: 3835 Question: When are outboard ailerons (if present) de-activated? A. Flaps (and/or slats) extended or speed below a certain value. B. Landing gear extended. C. Landing gear retracted. D. Flaps (and slats) retracted or speed above a certain value. On aircraft fitted with two sets of ailerons, (inboard and outboard), as the airspeed increases the aerodynamic loads on the ailerons tend to twist the wing at the tips, where it is more fiexible. To overcome this tendency, some aircraft use the technique of locking the outboard ailerons in the faired or neutral position, and use an inboard aileron/spoiler combination above fiap retraction speeds to provide the necessary roll control.

15554 - B

1416 - C

1417 - A

3869 - A

3835 - D

Number: 3830 Question: Which of the following statements concerning control is correct? A. Hydraulically powered control surfaces do not need mass balancing. B. On some aeroplanes, the servo tab also serves as a trim tab. C. In general the maximum downward elevator deflection is larger than upward. D. In a differential aileron control system the control surfaces have a larger upward than downward maximum deflection.

Number: 3504 Question: A jet aeroplane equipped with inboard and outboard ailerons is cruising at its normal cruise Mach number. In this case: A. only the spoilers will be active, not the ailerons. B. only the inboard ailerons are active. C. the inboard and outboard ailerons are active. D. only the outboard aileron are active. On aircraft fitted with two sets of ailerons, (inboard and outboard), as the airspeed increases the aerodynamic loads on the ailerons tend to twist the wing at the tips, where it is more fiexible. To overcome this tendency, some aircraft use the technique of locking the outboard ailerons in the faired or neutral position, and use an inboard aileron/spoiler combination above fiap retraction speeds to provide the necessary roll control.

Number: 14914 Question: What are the primary roll controls on a conventional aeroplane? A. Asymmetrically extended leading edge flaps. B. The rudder. C. The ailerons. D. Symmetrically deflected spoilers.

3830 - D

3504 - B

14914 - C

05-03 - CONTROL (2 Fragen)

Number: 3868 Question: An advantage of locating the engines at the rear of the fuselage, in comparison to a location beneath the wing, is : A. easier maintenance of the engines. beneath - darunter - unterhalb B. a wing which is less sensitive to flutter. C. less influence of thrust changes on longitudinal control. D. less influence on lateral/directional stability characteristics such as dutch roll.

Number: 3490 Question: Given two identical aeroplanes with wing mounted engines, one fitted with jet engines and the other with counter rotating propellers, what happens following an engine failure? A. More roll tendency for the propeller aeroplane. B. The same yaw tendency for both aeroplanes regardless of left or right engine failure. C. The same roll tendency for both aeroplanes. D. Less roll tendency for the propeller aeroplane.

3868 - C

3490 - A

05-02 - CONTROL (14 Fragen)

Number: 16412 Question: The elevator deflection required for a given manoeuvre will be: A. the same at all speeds. B. larger at high IAS when compared to low IAS. C. larger for an aft CG position when compared to a forward position. D. larger for a forward CG position when compared to an aft position.

Number: 16821 Question: Which statement is about CG limits is correct? A. The forward CG limit is mainly determined by the amount of pitch control available from the elevator. B. The aft CG limit is determined by the maximum elevator deflection available. C. If the aft CG limit is correctly chosen, the forward CG limit is automatically determined as well. D. The forward CG limit is determined by stability considerations only.

Number: 929 Question: The centre of gravity moving aft will: A. decrease the elevator up effectiveness B. increase the elevator up effectiveness C. increase or decrease the elevator up effectiveness, depending on wing location D. not affect the elevator up or down effectiveness

16412 - D

16821 - A

929 - B

Number: 957 Question: When the cg position is moved forward, the elevator deflection for a manoeuvre with a given load factor greater than 1 will be: A. larger. B. smaller. C. dependent on trim position. D. unchanged.

Number: 10133 Question: What is the effect of an aft shift of the centre of gravity on (1) static longitudinal stability and (2) the required control deflection for a given pitch change? A. B. C. D.

(1) reduces (2) reduces. (1) increases (2) increases. (1) reduces (2) increases. (1) increases (2) reduces.

Number: 13146 Question: A jet transport aeroplane exhibits pitch up when thrust is suddenly increased from an equilibrium condition, because the thrust line is below the: exhibits - aufwisen A. CG. B. centre of pressure. C. drag line of action. D. neutral point.

Number: 15921 Question: For a given elevator deflection, aeroplane longitudinal manoeuvrability increases when: A. flaps are retracted at constant IAS. B. IAS decreases. C. the CG moves forward. D. the CG moves aft.

957 - A

10133 - A

13146 - A

15921 - D

Number: 15874 Question: What kind of horizontal control surface is shown in the figure? A. Elevator. B. Canard. C. Frise type control. D. All-flying tail.

Number: 3506 Question: How does positive camber of an aerofoil affect static longitudinal stability ? It has A. positive effect, because the lift vector rotates backward at increasing angle of attack. B. no effect, because camber of the aerofoil produces a constant pitch down moment coefficient, independent of angle of attack. C. positive effect, because the centre of pressure shifts rearward at increasing angle of attack D. negative effect, because the lift vector rotates forward at increasing angle of

Number: 14591 Question: Low speed pitch-up can be caused by a significant thrust: A. increase with engines located on the rear fuselage. B. decrease with engines located on the rear fuselage. C. increase with podded engines located beneath a low-mounted wing. D. decrease with podded engines located beneath a low-mounted wing.

15874 D

3506 - B

14591 - C

Number: 14651 Question: The elevator deflection required for a given manoeuvre will be: A. the same at all speeds. B. larger for an aft CG position when compared to a forward position. C. smaller at high IAS when compared to low IAS. D. the same for all CG positions.

Number: 14652 Question: For a given elevator deflection, aeroplane longitudinal manoeuvrability decreases when: A. the CG moves aft. B. the CG moves forward C. flaps are retracted at constant IAS. D. IAS increases.

Number: 14653 Question: Aeroplane manoeuvrability increases for a given control surface deflection when: A. the CG moves forward. B. IAS decreases. C. flaps are retracted at constant IAS. D. IAS increases.

Number: 14980 Question: One advantage of mounting the horizontal tailplane on top of the vertical fin is: A. to decrease the susceptibility to deep stall. B. to improve the aerodynamic efficiency of the vertical fin. C. to decrease fuel consumption by creating a tail-heavy situation. D. that it does not require a de-icing system. Just as winglets improve the efficiency of a wing by reducing tip vortices, a horizontal stabiliser at the top of the fin will do the same to the fin.

14651 - C

14652 - B

14653 - D

14980 - B

05-01 - CONTROL (8 Fragen)

Number: 16809 Question: Which statement about elevators is correct? A. The elevator is the primary control surface for control about the longitudinal axis and is operated by a forward or backward movement of the control wheel or stick. B. The elevator is the primary control surface for control about the lateral axis and is operated by a forward or backward movement of the control wheel or stick. C. The elevator is used only to trim an aeroplane and is normally operated by a dedicated control wheel, which is usually situated close to the throttle. D. The elevator is the primary control surface for control about the longitudinal axis and is operated by a left or right rotation of the control wheel. Number: 933 Question: Rotation about the lateral axis is called: A. slipping. B. rolling. C. yawing. D. pitching. Number: 934 Question: Rolling is the rotation of the aeroplane about the: A. wing axis. B. longitudinal axis. C. lateral axis. D. vertical axis. Number: 12990 Question: The pitch angle is defined as the angle between the: A. chord line and the horizontal plane. B. speed vector axis and the longitudinal axis. C. longitudinal axis and the chord line. D. longitudinal axis and the horizontal plane. The angle of incidence is fixed, but the angle of attack changes in flight. Likewise, do not confuse the 'Pitch Angle' or 'Pitch Attitude' of the aircraft with the angle of attack. For any given angle of attack, the pitch angle can vary. Similarly for any given pitch angle, the angle of attack can also vary.

16809 - B

933 - D

934 - B

12990 - D

Number: 14574 Question: An aeroplane's pitch angle is defined as the angle between its: A. longitudinal axis and the horizontal plane. B. lateral axis and the horizontal plane. C. speed vector and its longitudinal axis. D. speed vector and the horizontal plane. The angle of incidence is fixed, but the angle of attack changes in flight. Likewise, do not confuse the 'Pitch Angle' or 'Pitch Attitude' of the aircraft with the angle of attack. For any given angle of attack, the pitch angle can vary. Similarly for any given pitch angle, the angle of attack can also vary.

Number: 14586 Question: An aeroplane's bank angle is defined as the angle between its: A. lateral axis and the horizontal plane. B. speed vector and the horizontal plane. C. longitudinal axis and the horizontal plane. D. speed vector and longitudinal axis.

Number: 14634 Question: Rotation around the longitudinal axis is called: A. pitching. B. yawing. C. slipping. D. rolling.

Number: 14635 Question: Rotation around the normal axis is called: A. rolling. B. slipping. C. yawing. D. pitching.

14574 - A

14586 - A

14634 - D

14635 - C

05-02 - CONTROL (14 Fragen)

Number: 16412 Question: The elevator deflection required for a given manoeuvre will be: A. the same at all speeds. B. larger at high IAS when compared to low IAS. C. larger for an aft CG position when compared to a forward position. D. larger for a forward CG position when compared to an aft position.

Number: 16821 Question: Which statement is about CG limits is correct? A. The forward CG limit is mainly determined by the amount of pitch control available from the elevator. B. The aft CG limit is determined by the maximum elevator deflection available. C. If the aft CG limit is correctly chosen, the forward CG limit is automatically determined as well. D. The forward CG limit is determined by stability considerations only.

Number: 929 Question: The centre of gravity moving aft will: A. decrease the elevator up effectiveness B. increase the elevator up effectiveness C. increase or decrease the elevator up effectiveness, depending on wing location D. not affect the elevator up or down effectiveness

16412 - D

16821 - A

929 - B

Number: 957 Question: When the cg position is moved forward, the elevator deflection for a manoeuvre with a given load factor greater than 1 will be: A. larger. B. smaller. C. dependent on trim position. D. unchanged.

Number: 10133 Question: What is the effect of an aft shift of the centre of gravity on (1) static longitudinal stability and (2) the required control deflection for a given pitch change? A. B. C. D.

(1) reduces (2) reduces. (1) increases (2) increases. (1) reduces (2) increases. (1) increases (2) reduces.

Number: 13146 Question: A jet transport aeroplane exhibits pitch up when thrust is suddenly increased from an equilibrium condition, because the thrust line is below the: exhibits - aufwisen A. CG. B. centre of pressure. C. drag line of action. D. neutral point.

Number: 15921 Question: For a given elevator deflection, aeroplane longitudinal manoeuvrability increases when: A. flaps are retracted at constant IAS. B. IAS decreases. C. the CG moves forward. D. the CG moves aft.

957 - A

10133 - A

13146 - A

15921 - D

Number: 15874 Question: What kind of horizontal control surface is shown in the figure? A. Elevator. B. Canard. C. Frise type control. D. All-flying tail.

Number: 3506 Question: How does positive camber of an aerofoil affect static longitudinal stability ? It has A. positive effect, because the lift vector rotates backward at increasing angle of attack. B. no effect, because camber of the aerofoil produces a constant pitch down moment coefficient, independent of angle of attack. C. positive effect, because the centre of pressure shifts rearward at increasing angle of attack D. negative effect, because the lift vector rotates forward at increasing angle of

Number: 14591 Question: Low speed pitch-up can be caused by a significant thrust: A. increase with engines located on the rear fuselage. B. decrease with engines located on the rear fuselage. C. increase with podded engines located beneath a low-mounted wing. D. decrease with podded engines located beneath a low-mounted wing.

15874 D

3506 - B

14591 - C

Number: 14651 Question: The elevator deflection required for a given manoeuvre will be: A. the same at all speeds. B. larger for an aft CG position when compared to a forward position. C. smaller at high IAS when compared to low IAS. D. the same for all CG positions.

Number: 14652 Question: For a given elevator deflection, aeroplane longitudinal manoeuvrability decreases when: A. the CG moves aft. B. the CG moves forward C. flaps are retracted at constant IAS. D. IAS increases.

Number: 14653 Question: Aeroplane manoeuvrability increases for a given control surface deflection when: A. the CG moves forward. B. IAS decreases. C. flaps are retracted at constant IAS. D. IAS increases.

Number: 14980 Question: One advantage of mounting the horizontal tailplane on top of the vertical fin is: A. to decrease the susceptibility to deep stall. B. to improve the aerodynamic efficiency of the vertical fin. C. to decrease fuel consumption by creating a tail-heavy situation. D. that it does not require a de-icing system. Just as winglets improve the efficiency of a wing by reducing tip vortices, a horizontal stabiliser at the top of the fin will do the same to the fin.

14651 - C

14652 - B

14653 - D

14980 - B

05-03 - CONTROL (2 Fragen)

Number: 3868 Question: An advantage of locating the engines at the rear of the fuselage, in comparison to a location beneath the wing, is : A. easier maintenance of the engines. beneath - darunter - unterhalb B. a wing which is less sensitive to flutter. C. less influence of thrust changes on longitudinal control. D. less influence on lateral/directional stability characteristics such as dutch roll.

Number: 3490 Question: Given two identical aeroplanes with wing mounted engines, one fitted with jet engines and the other with counter rotating propellers, what happens following an engine failure? A. More roll tendency for the propeller aeroplane. B. The same yaw tendency for both aeroplanes regardless of left or right engine failure. C. The same roll tendency for both aeroplanes. D. Less roll tendency for the propeller aeroplane.

3868 - C

3490 - A

05-04 - CONTROL (19 Fragen) Number: 16415 Question: The function of ailerons is to rotate the aeroplane about the: A. yaw axis. B. lateral axis. C. normal axis. D. longitudinal axis. Number: 16303 Question: Outboard ailerons (if present) are normally used: A. in low speed flight only. B. when the landing gear is up. C. in high speed flight only. D. at transonic and supersonic speeds only. On aircraft fitted with two sets of ailerons, (inboard and outboard), as the airspeed increases the aerodynamic loads on the ailerons tend to twist the wing at the tips, where it is more fiexible. To overcome this tendency, some aircraft use the technique of locking the outboard ailerons in the faired or neutral position, and use an inboard aileron/spoiler combination above fiap retraction speeds to provide the necessary roll control.

Number: 1001 Question: During initiation of a turn with speedbrakes extended, the roll spoiler function induces a spoiler deflection: A. downward on the upgoing wing and upward on the downgoing wing. B. on the downgoing wing only. C. upward on the upgoing wing and downward on the downgoing wing. D. on the upgoing wing only.

16415 - D

16303 - A

1001 - A

Number: 2364 Question: How is adverse yaw compensated for during entry into and roll out from a turn? A. Differential aileron deflection. B. Anti-balanced rudder control. C. Servo tabs. D. Horn-balanced controls.

Number: 3109 Question: One method to compensate adverse yaw is: A. an anti-balance tab. B. a differential aileron. C. a balance panel. D. a balance tab.

Number: 3110 Question: Flaperons are controls which combine the function of: A. flaps and elevator. Some commercial aircraft have ailerons that droop when flaps are extended to B. ailerons and elevator. increase the lift co-efficient, these are flaperons. C. ailerons and flaps. D. flaps and speed brakes. There are some other aircraft (mainly military) that have split ailerons that can act as speed brakes as well, but these are not called flaperons.

Number: 13125 Question: An example of differential aileron deflection during initiation of left turn is: A. Left aileron: 5° up. Right aileron: 2° down. B. Left aileron: 2° down. Right aileron: 5° up. C. Left aileron: 2° up. Right aileron: 5° down. D. Left aileron: 5° down. Right aileron: 2° up.

Number: 13162 Question: An aeroplane is provided with spoilers and both inboard and outboard ailerons. Roll control during cruise is provided by: On aircraft fitted with two sets of ailerons, (inboard and outboard), as the A. inboard ailerons and roll spoilers. airspeed increases the aerodynamic loads on the ailerons tend to twist the B. inboard and outboard ailerons. wing at the tips, where it is more fiexible. C. outboard ailerons only. To overcome this tendency, some aircraft use the technique of locking the D. outboard ailerons and roll spoilers. outboard ailerons in the faired or neutral position, and use an inboard aileron/spoiler combination above fiap retraction speeds to provide the necessary roll control.

2364 - A

3109 - B

3110 - C

13125 - A

13162 - A

Number: 13236 Question: In what phase of flight are the outboard ailerons (if fitted) not active? A. Landing with a strong and gusty crosswind, to avoid over-controlling the aeroplane. B. Cruise. C. Take-off, until lift-off. D. Approach. On aircraft fitted with two sets of ailerons, (inboard and outboard), as the airspeed increases the aerodynamic loads on the ailerons tend to twist the wing at the tips, where it is more fiexible. To overcome this tendency, some aircraft use the technique of locking the outboard ailerons in the faired or neutral position, and use an inboard aileron/spoiler combination above fiap retraction speeds to provide the necessary roll control.

Number: 12974 Question: When roll spoilers are extended, the part of the wing on which they are mounted: A. experiences a reduction in lift, which generates the desired rolling moment. In addition there is a local increase in drag, which suppresses adverse yaw. B. is forced downwards as a reaction to the increased drag. C. stalls. This causes a difference in lift between both wings, which generates the desired rolling moment. D. experiences extra drag, which generates a yawing moment. The speed difference between both wings generates the desired rolling moment.

Number: 15608 Question: Aileron deflection causes a rotation around the longitudinal axis by: A. changing the wing drag and the two wings therefore produce different lift values resulting in a moment about the longitudinal axis. B. aileron secondary effect. C. causing sideslip, which generates a rolling moment. D. changing the wing camber and the two wings therefore produce different lift values resulting in a moment about the longitudinal axis.

13236 - B

12974 - A

15608 - D

Number: 15554 Question: When a turn is initiated, adverse yaw is: A. a momentary yawing motion opposite to the turn due to an incorrect differential aileron movement. B. the tendency of an aeroplane to yaw in the opposite direction of turn mainly due to the difference in induced drag on each wing. C. the tendency of an aeroplane to yaw in the same direction of turn due to the different wing speeds. D. the tendency of an aeroplane to yaw in the opposite direction of turn mainly due to the difference in aileron form drag.

Number: 1416 Question: Differential aileron deflection: A. increases the CLmax. B. is required to achieve the required roll rate. C. equals the drag of the right and left aileron. D. is required to keep the total lift constant when ailerons are deflected.

Number: 1417 Question: An example of differential aileron deflection during initiation of left turn is: A. Left aileron: 5° up. Right aileron: 2° down. B. Left aileron: 2° down. Right aileron: 5° up. C. Left aileron: 2° up. Right aileron: 5° down. D. Left aileron: 5° down. Right aileron: 2° up.

Number: 3869 Question: Which phenomenon is counteracted with differential aileron deflection? A. Adverse yaw. B. Turn co-ordination. counteracted - entgegengearbeitet , entgegengewirkt C. Sensitivity for spiral dive. D. Aileron reversal.

Number: 3835 Question: When are outboard ailerons (if present) de-activated? A. Flaps (and/or slats) extended or speed below a certain value. B. Landing gear extended. C. Landing gear retracted. D. Flaps (and slats) retracted or speed above a certain value. On aircraft fitted with two sets of ailerons, (inboard and outboard), as the airspeed increases the aerodynamic loads on the ailerons tend to twist the wing at the tips, where it is more fiexible. To overcome this tendency, some aircraft use the technique of locking the outboard ailerons in the faired or neutral position, and use an inboard aileron/spoiler combination above fiap retraction speeds to provide the necessary roll control.

15554 - B

1416 - C

1417 - A

3869 - A

3835 - D

Number: 3830 Question: Which of the following statements concerning control is correct? A. Hydraulically powered control surfaces do not need mass balancing. B. On some aeroplanes, the servo tab also serves as a trim tab. C. In general the maximum downward elevator deflection is larger than upward. D. In a differential aileron control system the control surfaces have a larger upward than downward maximum deflection.

Number: 3504 Question: A jet aeroplane equipped with inboard and outboard ailerons is cruising at its normal cruise Mach number. In this case: A. only the spoilers will be active, not the ailerons. B. only the inboard ailerons are active. C. the inboard and outboard ailerons are active. D. only the outboard aileron are active. On aircraft fitted with two sets of ailerons, (inboard and outboard), as the airspeed increases the aerodynamic loads on the ailerons tend to twist the wing at the tips, where it is more fiexible. To overcome this tendency, some aircraft use the technique of locking the outboard ailerons in the faired or neutral position, and use an inboard aileron/spoiler combination above fiap retraction speeds to provide the necessary roll control.

Number: 14914 Question: What are the primary roll controls on a conventional aeroplane? A. Asymmetrically extended leading edge flaps. B. The rudder. C. The ailerons. D. Symmetrically deflected spoilers.

3830 - D

3504 - B

14914 - C

05-05 - CONTROL (5 Fragen)

Number: 16637 Question: Yaw is followed by roll because the: A. rolling motion generated by rudder deflection causes a speed increase of the outer wing which increases the lift on that wing so the aeroplane starts to roll in the direction of the turn. B. yawing motion generated by rudder deflection causes a speed increase of the inner wing, which increase the lift on that wing so that the aeroplane starts to roll in the same direction as the yaw. C. rudder is located above the longitudinal axis and when it is deflected, it causes a rolling moment in the same direction as the yaw. D. yawing motion generated by rudder deflection causes a speed increase of the outer wing, which increases the lift on that wing so that the aeroplane starts to roll in the same direction as the yaw.

Number: 16726 Question: Rotation about the longitudinal axis of an aeroplane can be achieved by: A. aileron deflection and/or rudder deflection. B. speed brake extension or wing flap deflection. C. elevator deflection and/or slat extension. D. symmetrical spoiler deflection and/or elevator deflection.

Number: 178 Question: If the nose of an aeroplane yaws left, this causes: A. a roll to the right. B. an increase in lift on the left wing. C. a roll to the left. D. a decrease in relative airspeed on the right wing.

Number: 961 Question: Which moments or motions interact in a dutch roll? A. Pitching and adverse yaw. B. Pitching and yawing. C. Pitching and rolling. D. Rolling and yawing.

Number: 13212 Question: If the nose of an aeroplane yaws left, this causes: A. a decrease in relative airspeed on the right wing. B. a roll to the right. C. a roll to the left. D. an increase in lift on the left wing.

16637 - D

16726 - A

178 - C

961 - D

13212 - C

05-06 - CONTROL (17 Fragen)

Number: 16469 Question: What is the primary input for an artificial feel system? A. TAS. To prevent over-controlling and overstressing the aircraft, some form of artificial feel is B. Mach number. incorporated in the control system, so C. Static pressure. that the control forces experienced represent those of a manually contro lled aircraft. D. IAS. On transport category aircraft feel forces are provided by - spring units, - Pitot-static Q-feel units, - or a combination of both. Artificial feel systems normally also incorporate a self-centring mechanism, so that the flight deck controls automatically return to their neutral positions when released, centralising the control surfaces.

Number: 1002 Question: Stick forces, provided by an elevator feel system, depend on: A. stabiliser position, static pressure. B. stabiliser position, total pressure. C. elevator deflection, static pressure. D. elevator deflection, dynamic pressure.

Number: 2355 Question: Which kind of ''tab'' is commonly used in case of manual reversion of fully powered flight controls? A. Spring tab. B. Servo tab. C. Anti-balance tab. D. Balance tab. Servo Tab is directly controlled by the pilot through a pivot point and movement of the tab supplies the hinge moment necessary to move the main control surface. Movement of the tab causes the control surface to move to a new position of equilibrium in a direction of travel opposite to that of the tab (i.e. tab down, control surface up). In practice, the servo tab lacks effectiveness at low airspeeds when large control defiections are required. to lack - fehlen This is because the amount of airfiow passing over the tab is too low to produce the necessary hinge moment and hence the required defiection.

Number: 2363 Question: Which statement is correct about a spring tab? A. At high IAS it behaves like a servo tab. B. Its main purpose is to increase stick force per g. C. At low IAS it behaves like a servo tab. D. At high IAS it behaves like a fixed extension of the elevator.

16469 - D

1002 - D

2355 - B

2363 - A

Number: 13087 Question: Which three aerodynamic means decrease manoeuvring stick forces? A. Spring tab - horn balance - bobweight. Horn Balance B. Servo tab - trim tab - balance tab. C. Servo tab - horn balance - spring tab. D. Spring tab - trim tab - mass balancing weight.

Number: 12975 Question: When power assisted controls are used for pitch control: A. they only function in combination with an elevator trim tab. B. aerodynamic balancing of the control surfaces is meaningless. C. trimming is superfluous. D. a part of the aerodynamic forces is still felt on the column.

Number: 15514 Question: The tab in the figure represents: A. a servo tab. B. a balance tab . C. an anti-balance tab. D. a trim tab. Servo Tab is directly controlled by the pilot through a pivot point and movement of the tab supplies the hinge moment necessary to move the main control surface. Movement of the tab causes the control surface to move to a new position of equilibrium in a direction of travel opposite to that of the tab (i.e. tab down, control surface up). In practice, the servo tab lacks effectiveness at low airspeeds when large control defiections are required. to lack - fehlen This is because the amount of airfiow passing over the tab is too low to produce the necessary hinge moment and hence the required defiection.

13087 - C

12975 - D

15514 - A

Number: 15524 Question: The tab in the figure represents: A. a trim tab. B. a control tab. C. an antibalance tab. D. a balance tab that also functions as a trim tab.

Balance Tabs They are connected to the tailplane by a mechanical linkage that causes them to move in the opposite direction to the control surface

Number: 3849 Question: When power assisted controls are used for pitch control: A. they only function in combination with an elevator trim tab. B. trimming is superfluous. C. aerodynamic balancing of the control surfaces is meaningless. D. a part of the aerodynamic forces is still felt on the column.

Number: 3861 Question: Examples of aerodynamic balancing of control surfaces are: A. balance tab, horn balance, and mass balance. B. servo tab, spring tab, seal between the wing trailing edge and the leading edge of control surface. C. spring tab, servo tab, and power assisted control. D. mass in the nose of the control surface, horn balance and mass balance.

15524 - D

3849 - D

3861 - B

Number: 3870 Question: An aeroplane has a servo tab controlled elevator. What will happen if the elevator jams during flight? jams - klemmt A. The pitch control forces double. B. Pitch control is lost. C. Pitch control sense is reversed. D. The servo-tab now works as a negative trim-tab. Only the tab is left functioning. It is not connected to the elevator but works in the unnatural sense to drive the elevator in the natural sense, so with the elevator locked you only have very limited control from the tab only and in the unnatural sense - and, of course, you have no trim or balance function left.

Number: 3880 Question: A horn balance in a control system has the following purpose: A. to obtain mass balancing. B. to decrease the effective longitudinal dihedral of the aeroplane. C. to prevent flutter. D. to decrease stick forces.

Horn Balance

Number: 3657 Question: Examples of aerodynamic balancing of control surfaces are: A. weight in the nose of the control surface, horn balance. B. Fowler flaps, upper and lower rudder. C. upper and lower rudder, seal between wing's trailing edge and leading edge of a control surface. D. seal between wing's trailing edge and leading edge of a control surface, horn balance. The answer would make sense if it said "a seal between the wing trailing edge and the leading edge of the control surface", but even that would be a bit untidy because in a seal balance the seal is in the gap between the wing housing for the control and the leading edge of the control. The use of wing "trailing edge" is confusing

Number: 3654 Question: Which statement about a primary control surface controlled by a servo tab, is correct? A. The servo tab can also be used as a balance tab. B. Due to the effectiveness of the servo tab the control surface area can be smaller. C. The control effectiveness of the primary surface is increased by servo tab deflection. D. The position is undetermined during taxiing, in particular with tailwind. The function of a servo tab is very different from a balance tab. With a servo tab control system movement of the pilot’s flight controls moves the servo tab. The servo tab at the trailing edge of the main flying control surface produces a aerodynamic force to move the control surface. The servo tab is displaced in the opposite direction in which the flight control surface moves. ie if you wish to pitch the aircraft nose up, servo tab is deflected down and moves the elevator up. The system requires airflow from leading edge to trailing edge, when taxiing in a tailwind the effectiveness of this type of control is reduced.

3870 - C

3880 - D

3657 - D

3654 - D

Number: 14857 Question: Which aerodynamic design features can be used to reduce control forces? A. Balance tab, control surfaces with increased area behind the hinge, artificial feel system. B. Horn balance, balance tab, servo tab. C. Servo tab, bobweight, control surfaces with increased area. D. Mass balance, horn balance, artificial feel system.

Number: 14858 Question: In general, control forces are reduced by: A. a servo tab, spring tab and mass balancing. B. a horn balance, servo tab and spring tab. C. mass balancing, a trim tab and spring tab. D. a balance tab, forward shift of the CG and a servo tab.

Number: 15068 Question: Artificial feel is required: A. when the flight control surfaces are fitted with control tabs or trim tabs. B. with power assisted flight controls. C. with fully powered flight controls. D. when there is a trimmable stabiliser.

14857 - B

14858 - B

15068 - C

05-07 - CONTROL (1 Fragen)

Number: 3888 Question: When flutter damping of control surfaces is obtained by mass balancing, these weights will be located with respect to the hinge of the control surface: A. behind the hinge. B. above the hinge. C. in front of the hinge. D. below the hinge.

3888 - C

05-08 - CONTROL (23 Fragen)

Number: 2361 Question: One advantage of a movable-stabiliser system compared with an elevator trim system is that: A. the system's complexity is reduced. B. it is a more effective means of trimming. C. it leads to greater stability in flight. D. the complete system (structure and control mechanism) weighs less.

Number: 13075 Question: Which statement in respect of a trimmable horizontal stabiliser is correct? A. An aeroplane with a forward cg requires the stabiliser leading edge to be higher than for one with an aft cg in the same trimmed condition. B. An aeroplane with a forward cg requires the stabiliser leading edge to be lower than for one with an aft cg in the same trimmed condition. C. At the forward C.G. limit , stabiliser trim is adjusted fully nose down to obtain maximum elevator authority at rotation during take-off. D. Because take-off speeds do not vary with centre of gravity location, the need for stabiliser adjustment is dependent on flap position only. If the C.G is moved aft it results in a larger nose down pitching moment which has to be compensated for by placing a download on the tail plane. This can be achieved by deflecting the elevator up, although this results in high trim drag at higher airspeeds. On large aircraft small movements of the variable incidence stabilizer allow the trimming to be carried out, but with reduced trim drag. An upward deflection of the elevator is required to trim the aircraft, so think of the stabilizer as a large elevator with its trailing edge moving in the same direction as the elevator. To achieve this you must therefore lower the leading edge of the stabilizer to produce the necessary download. Note: Once in a trimmed condition on fully powered flying controls the elevator will then align itself with the stabilizer to reduce the tail load and also the loads acting on the hinges and servo actuator. On power assisted controls there is however no guarantee that this occurs due to downwash and configuration effects that can alter the flow over the surface. It is also fair to say that trim drag also increases fuel consumption.

Number: 13214 Question: If the elevator trim tab is deflected up, the cockpit trim indicator shows: A. nose down. B. nose left. C. nose up. D. neutral.

2361 - B

13075 - B

13214 - A

Number: 13241 Question: In straight and level flight, as speed is reduced: A. the elevator is deflected further downwards and the trim tab further upwards. B. both elevator and trim tab are deflected further upwards. C. the elevator and trim tab do not move. D. the elevator is deflected further upwards and the trim tab further downwards.

Number: 12954 Question: What is the effect of elevator trim tab adjustment on the static longitudinal stability of an aeroplane? A. Aeroplane nose down trim increases the static longitudinal stability. B. Aeroplane nose up trim increases the static longitudinal stability. C. Depends on the value of stick force/g. D. No effect.

Number: 15573 Question: When comparing a stabiliser trim system with an elevator trim system, which of these statements is correct? A. an elevator trim is able to compensate larger changes in pitching moments. B. an elevator trim is more suitable for aeroplanes with a large CG range. C. a stabiliser trim is able to compensate larger changes in pitching moments. D. a stabiliser trim is more sensitive to flutter. The Variable Incidence Horizontal Stabiliser provides pitch trim on most transport category aircraft. Varying the angle of incidence has the same effect as moving the elevator, but is aerodynamically more efficient, particularly at high airspeeds, and can provide a considerable trim range.

Number: 15734 Question: In straight flight, as speed is reduced, whilst trimming to keep the stick force zero: A. the elevator and trim tab do not move. B. the elevator is deflected further upwards and the trim tab further downwards. C. the elevator is deflected further downwards and the trim tab further upwards. D. both elevator and trim tab are deflected further upwards.

Number: 1415 Question: If the elevator trim tab is deflected up, the cockpit trim indicator presents: A. nose-up. B. neutral. C. nose-left. D. nose-down.

13241 - D

12954 - D

15573 - C

15734 - B

1415 - D

Number: 3882 Question: What is the position of the elevator in relation to the trimmable horizontal stabiliser of a power assisted aeroplane that is in trim? A. The position depends on speed, the position of slats and flaps and the position of the centre of gravity. B. The elevator is always deflected slightly downwards in order to have sufficient remaining flare capability. C. The elevator deflection (compared with the stabiliser position) is always zero. D. At a forward CG the elevator is deflected upward and at an aft CG the elevator is deflected downward. In the power assisted hydraulic controls you move the control surface then the hydraulics assist you, so you do have feedback from the controls. To give you that feel there will always be some angle between the control surface and the tail plane, except if is at absolute zero. When trimmed you will have no forces on the stick, but the moment you move it out of the trimmed position you must experience the air load on the control surface.

The Variable Incidence Horizontal Stabiliser provides pitch trim on most transport category aircraft. Varying the angle of incidence has the same effect as moving the elevator, but is aerodynamically more efficient, particularly at high airspeeds, and can provide a considerable trim range.

Number: 3887 Question: When a jet transport aeroplane takes off with the CG at the forward limit and the trimmable horizontal stabiliser (THS) is positioned at the maximum allowable nose down position for take-off: A. rotation will require a higher than normal stick force. B. rotation will be normal using the normal rotation technique. C. early nose wheel raising will take place. D. there will be a tendency to over-rotate.

Number: 3476 Question: What should be usually done to perform a landing with the stabiliser jammed in the cruise flight position? A. if possible, relocate as many passengers as possible to the front of the cabin. B. use the Mach trimmer until after landing. C. choose a lower landing speed than normal. D. choose a higher landing speed than normal and/or use a lower flapsetting for landing.

3882 - A

3887 - A

3476 - D

Number: 3477 Question: In general transport aeroplanes with power assisted flight controls are fitted with an adjustable stabiliser instead of trim tabs on the elevator. This is because: A. trim tab deflection increases VMO. B. effectiveness of trim tabs is insufficient for those aeroplanes. C. mechanical adjustment of trim tabs creates too many problems. D. the pilot does not feel the stick forces at all.

Number: 14485 Question: In straight flight, as speed is increased, whilst trimming to keep the stick force zero: A. the elevator is deflected further upwards and the trim tab further downwards. B. both elevator and trim tab are deflected further downwards. C. the elevator and trim tab do not move. D. the elevator is deflected further downwards and the trim tab further upwards.

Number: 14306 Question: The most important factor determining the required position of the Trimmable Horizontal Stabiliser (THS) for take off is the: A. total mass of the aeroplane. B. position of the aeroplane''s centre of gravity. C. centre of gravity position of the fuel. D. stall speed.

Number: 14310 Question: What is the position of the elevator in relation to the trimmable horizontal stabiliser of an aeroplane with fully hydraulically operated flight controls that is in trim? A. At a forward CG, the elevator is deflected upward and at an aft CG, it is deflected downward. B. The position depends on speed, the position of flaps and slats and the position of the centre of gravity. C. Elevator deflection is zero. D. The elevator is always deflected slightly downward in order to have sufficient remaining flare capability.

3477 - B

14485 - D

14306 - B

14310 - C

Number: 14311 Question: What is the effect on landing speed when a trimmable horizontal stabiliser jams at high IAS? A. In most cases, no effect. B. No effect with a forward CG. C. No effect when landing on a high elevation runway. D. In most cases, a higher than normal landing speed is required.

Number: 14597 Question: When a jet transport aeroplane takes off with the CG at the forward limit and the trimmable horizontal stabiliser (THS) is positioned at the maximum allowable nose up position for take-off: A. rotation will require a higher than normal stick force. B. early nose wheel raising will take place. C. rotation will be normal using the normal rotation technique. D. there will be a tendency to over-rotate.

14311 - D

14597 - C

Number: 14598 Question: When a jet transport aeroplane takes off with the CG at the aft limit and the trimmable horizontal stabiliser (THS) is positioned at the maximum allowable nose up position for take-off: A. early nose wheel raising will take place. B. rotation will require higher than normal stick force. C. there will be a tendency to under-rotate. D. rotation will be normal using the normal rotation technique.

Number: 14659 Question: Comparing the differences between a horizontal trimmable stabiliser and an elevator trim tab, which of these statements are correct or incorrect? I. The effects of a trim tab runaway are more serious. II. A jammed trim tab causes less control difficulty. A. B. C. D.

I is incorrect, II is incorrect. I is correct, II is incorrect. I is incorrect, II is correct. I is correct, II is correct.

14598 - A

14659 - C

Number: 14682 Question: Which of these statements about a trimmable horizontal stabiliser is correct? A. At the aft CG limit, stabiliser trim is adjusted fully nose up to obtain maximum elevator authority at rotation during take-off. B. A trimmed aeroplane with an aft CG requires the stabiliser leading edge to be lower than in the case of a forward CG in the same condition. C. Because take-off speeds do not vary with CG position, the need for stabiliser adjustment is dependent on flap position only. D. A trimmed aeroplane with an aft CG requires the stabiliser leading edge to be higher than in the case of a forward CG in the same condition.

Number: 14690 Question: When comparing an elevator trim system with a stabiliser trim system, which of these statements is correct? A. an elevator trim produces lower trim drag B. an elevator trim is able to compensate larger changes in pitching moments. C. an elevator trim is more sensitive to flutter. D. an elevator trim is more suitable for aeroplanes with a large CG range. Number: 14691 Question: When comparing a stabiliser trim system with an elevator trim system, which of these statements is correct? A. an elevator trim is more suitable for aeroplanes with a large CG range. B. a stabiliser trim is not as capable to compensate large changes in pitching moments. C. an elevator trim is able to compensate larger changes in pitching moments. D. a stabiliser trim is less sensitive to flutter.

Number: 14692 Question: When comparing a stabiliser trim system with an elevator trim system, which of these statements is correct? A. a elevator trim is able to compensate larger changes in pitching moments. B. an elevator trim is less suitable for aeroplanes with a large CG range. C. an elevator trim is more suitable for aeroplanes with a large CG range. D. a stabiliser trim is more sensitive to flutter.

14682 - D

14690 - C

14691 - D

14692 - B

07-00 NEW QUESTIONS

PRINCIPLES OF FLIGHT

07-01 New Questions (20 Questions)

Number: 16083 Question: Which of these statements about an oblique shock wave are correct or incorrect? I. The density behind an oblique shock wave is lower than in front of it. II. The local speed of sound behind an oblique shock wave is lower than in front of it. A. I is incorrect, II is correct. B. I is incorrect, II is incorrect. C. I is correct, II is incorrect. D. I is correct, II is correct.

OBLIQUE SHOCK WAVE An oblique shock wave is a compression wave and is similar to a normal shock wave, except that the airflow changes direction into a corner and its velocity decreases to a lower supersonic value. The wave angle depends on the Mach number of the approaching flow and the angle of the wedge. This type of shock wave is weaker than the normal shock wave, but the energy loss still has to be overcome by the aircraft engines. As the air passes through an oblique shock wave its pressure, temperature and density all increase.

16083 - B

Number: 16165 Question: Which of these statements about the gust load factor on an aeroplane are correct or incorrect? I. When the wing area decreases, the gust load factor decreases. II. When the EAS decreases, the gust load factor decreases. A. I is correct, II is incorrect. B. I is incorrect, II is correct. C. I is incorrect, II is incorrect. D. I is correct, II is correct.

GUST LOAD Speed VB (max gust intensity speed) +/- 66 ft/sec. Speed VC (design cruise speed) +/- 50 ft/sec. Speed VD (design dive speed) +/- 25 ft/sec.

Number: 16185 Question: Considering subsonic incompressible airflow through a Venturi, which statement is correct? I. The dynamic pressure in the undisturbed airflow is lower than in the throat. II. The total pressure in the undisturbed airflow is higher than in the throat. A. I is incorrect, II is correct. B. I is correct, II is correct. C. I is incorrect, II is incorrect. D. I is correct, II is incorrect.

16165 - D

16185 - D

Number: 16187 Question: Considering subsonic incompressible airflow through a Venturi, which statement is correct? I. The dynamic pressure in the throat is higher than in the undisturbed airflow. II. The total pressure in the throat is higher than in the undisturbed airflow. A. I is incorrect, II is incorrect. B. I is incorrect, II is correct. C. I is correct, II is correct. D. I is correct, II is incorrect.

Number: 16192 Question: Which statement, about an aeroplane leaving ground effect at constant angle of attack, is correct? I. The lift coefficient CL decreases. II. The induced drag coefficient CDi decreases. A. I is incorrect, II is correct. B. I is correct, II is correct. C. I is incorrect, II is incorrect. D. I is correct, II is incorrect.

GROUND EFFECT

Ground effect occurs because the surface alters the airflow pattern around the wings. Primarily, the surface restricts the formation of the wingtip vortices. This results in a reduction in the amount of induced downwash behind the wing and increases the wing's effective angle of attack Upwash and downwash are reduced, causing the effective angle of attack of the wing to increase, (ref: Fig. 5.12). Therefore, when an aircraft is "in ground effect" lift will generally be increased and induced drag (CD) will be decreased. In addition, the reduced downwash will effect both longitudinal stability because of CP movement, and the pitching moment because of changes to the effective angle of attack of the tailplane, (Ref: Fig. 5.21).

16187 - D

16192 - D

Number: 16194 Question: Which of these statements about induced drag are correct or incorrect? I. Induced drag decreases as angle of attack decreases. II. At constant load factor, induced drag increases with decreasing aeroplane mass. A. I is correct, II is correct. B. I is incorrect, II is correct. C. I is incorrect, II is incorrect. D. I is correct, II is incorrect. The speed at which total drag is a minimum (V md) occurs when the induced and parasite drag are equal

Limit load factors are based on the maximum weight of the aircraft.

16194 - D

Number: 16196 Question: Which of these statements about flutter are correct or incorrect? I. Aero-elastic coupling affects flutter characteristics. II. Occurrence of flutter is independent of IAS. A. I is incorrect, II is correct. B. I is correct, II is correct. C. I is correct, II is incorrect. D. I is incorrect, II is incorrect.

Flutter is caused by the combined effects of - changes in the pressure distribution around the control surface

FLUTTER - MASS BALANCE

- with changing angles of attack (aerodynamic forces), and the forces due to the elastic nature of the aircraft structure itself (aeroelastic forces). If these forces become coincident and act in phase with each other, the resultant oscillations quickly increase in amplitude, and if left unchecked, may ultimately lead to structural failure.

To help eliminate fiutter in flight, manually operated control surfaces are generally mass balanced. Attaching weights forward of the hinge line brings the centre of gravity of the control surface to the hinge-line, thus altering the period of vibration and the liability to fiutter. These additional weights are usually installed internally along the leading edge of the control surface, inside the horn balance, or on an arm attached to the surface

Number: 16199 Question: Comparing the differences between a horizontal trimmable stabiliser and an elevator trim tab, which statement is correct? I. A trim tab is less suitable for jet transport aeroplanes because of their large speed range. II. A trim tab is a more powerful means of trimming. A. I is correct, II is correct. B. I is correct, II is incorrect. C. I is incorrect, II is incorrect. D. I is incorrect, II is correct.

16196 - C

16199 - B

Number: 16203 Question: Which of these statements are correct or incorrect regarding a sideslip, with the relative airflow coming from the left, on an aeroplane that exhibits both directional and lateral stability? I. The initial tendency of the nose of the aeroplane is to move to the left. II. The initial tendency of the left wing is to move down. A. I is incorrect, II is incorrect. B. I is incorrect, I is correct. C. I is correct, II is correct. D. I is correct, II is incorrect.

Number: 16206 Question: Which of these statements about an oblique shock wave are correct or incorrect? I. The density behind an oblique shock wave is higher than in front of it. II. The local speed of sound behind an oblique shock wave is lower than in front of it. A. I is incorrect, II is incorrect. B. I is correct, II is incorrect. C. I is correct, II is correct. D. I is incorrect, II is correct.

OBLIQUE SHOCK WAVE An oblique shock wave is a compression wave and is similar to a normal shock wave, except that the airflow changes direction into a corner and its velocity decreases to a lower supersonic value. The wave angle depends on the Mach number of the approaching flow and the angle of the wedge. This type of shock wave is weaker than the normal shock wave, but the energy loss still has to be overcome by the aircraft engines. As the air passes through an oblique shock wave its pressure, temperature and density all increase.

16203 - D

16206 - B

Number: 16216 Question: 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 increases as the angle of attack increases. II. The strength of wing tip vortices decreases as the aspect ratio increases. A. I is correct, II is correct. B. I is incorrect, II is incorrect. C. I is incorrect, II is correct. D. I is correct, II is incorrect.

ASPECT RATIO

S = WING AREA, sq. m (b x c) b = SPAN, m c = AVERAGE CHORD, m AR = ASPECT RATIO AR = b/c AR = b²/S

16216 - A

Number: 71 Question: The lift coefficient of a symmetrical aerofoil section at zero angle of attack is: A. negative. B. positive, if the centre of pressure coincides with the centre of gravity. C. positive. D. zero.

Number: 3857 Question: Which of the following wing planforms gives the highest local profile lift coefficient at the wingroot ? A. Rectangular. B. Tapered. C. Positive angle of sweep. D. Elliptical. The tapered airfoil is desirable from the standpoint of weight and stiffness, but again is not as efficient aerodynamically as the elliptical wing. In order to preserve the aerodynamic efficiency of the elliptical wing, rectangular and tapered wings are sometimes "tailored" through use of wing twist and variation in airfoil sections until they provide as nearly as possible the elliptical wing's lift distribution. While it is true that the elliptical wing provides the best lift coefficients before reaching an incipient stall, it gives little advance warning of a complete stall, and lateral control may be difficult because of poor aileron effectiveness. In comparison, the rectangular wing has a tendency to stall first at the wing root and provides adequate stall warning, adequate aileron effectiveness, and is usually quite stable. It is, therefore, favored in the design of low cost, low speed airplanes. Stall progression patterns for various wing planforms are graphically depicted in Figure 17-13. Note that it is possible for the trailing edge of the inboard portion of the rectangular wing to be stalled while the rest of the wing is developing lift. This is a very desirable characteristic, and along with simplicity of construction is the reason why this type of wing is so popular in light airplanes, despite certain structural and aerodynamic inefficiencies.

71 - D

3857 - A

The elliptical wing is the ideal subsonic planform since it provides for a minimum of induced drag for a given aspect ratio, though as we shall see, its stall characteristics in some respects are inferior to the rectangular wing. It is also comparatively difficult to construct.

Number: 6641 Question: Left rudder input will cause: A. Left yaw about the vertical axis and right roll about the longitudinal axis. B. Left yaw about the vertical axis and left roll about the longitudinal axis. C. Right yaw about the vertical axis and right roll about the longitudinal axis. D. Right yaw about the vertical axis and left roll about the longitudinal axis.

Number: 6642 Question: If a turbulent gust causes an aeroplane to roll: A. The down going wing experiences an increase in angle of attack. B. The down going wing experiences a decrease in angle of attack. C. The angle of attack depends on whether the aeroplane changes speed. D. The down going wing has no angle of attack.

Number: 6645 Question: Which forces produce the necessary normal acceleration to make an aircraft turn? A. Centrifugal force. B. The vertical component of lift. C. The horizontal component of lift. D. The horizontal component of weight.

6641 - B

6642 - A

6645 - C

Number: 6647 Question: When an aircraft is rolled to port, adverse yaw will be reduced by: A. Frise ailerons producing increased profile leading edge drag on both surfaces. B. The down going aileron producing a greater angle of deflection than the up going aileron. C. A Frise aileron being effective on the port wing. D. The leading edge of the down going aileron protruding into the airflow. port {adj} - Backbord - links

ADVERSE AILERON YAW negatiwe Wendemoment Differential ailerons are designed so that the up-going aileron is deflected through a greater angle than the down-going aileron.

Frise ailerons are designed so that the leading edge of the aileron projects beneath the wing when the aileron is defiected upward.

Number: 6650 Question: The inputs to the Q feel unit are from: A. Pitot pressure and total head pressure. B. Static pressure and temperature. C. Pitot and static pressures. D. Altitude and pitot pressure.

On transport category aircraft feel forces are provided by spring units, Pitot-static Q-feel units, or a combination of both. Artificial feel systems normally also incorporate a self-centring mechanism, so that the flight deck controls automatically return to their neutral positions when released, centralising the control surfaces. Like spring feel units, Q-feel Units are fitted in the operating linkage between the fiight deck controls and the power control units

6647 -C

6650 - C

Number: 6652 Question: The reasons for having a trim system on a powered assisted flying controls is: A. Relieve stresses on the trim tab. B. Enables the pilot to maintain control in case of hydraulic failure. C. Relieve stresses on the hydraulic actuators. D. Enables the stick force to be reduced to zero.

Number: 6654 Question: Deflecting the elevator up, when the trim tab is in neutral, will cause the tab to: A. Remain in line with the elevator. B. Move up relative to the elevator chord line. C. Remain in line with the tailplane. D. Move down relative to the elevator chord line.

6652 - D

6654 - A

07-00 NEW QUESTIONS

PRINCIPLES OF FLIGHT

07-02 New Questions (20 Questions)

Number: 6655 Question: What is the advantage of a variable incidence tailplane over a fixed incidence tailplane with elevator and trim tab? A. Linkages and mechanism less complicated. B. Elevator movement is restricted at high speed. C. Less trim drag and maximum elevator authority retained. D. Increased flight stability and less weight.

6655 - C

Number: 6658 Question: An aircraft is equipped with an all flying tailplane which has a combined anti-balance and trimming tab. The top of the trim wheel is moved forward. Which of the following statements is most correct? A. The tab moves up, so that more effect is required when the pilot attempts to move the control column to the rear. B. The tab moves down, so that more effort is required when the pilot attempts to move the control column to the rear. C. The tab moves up, so that less effort is required when the pilot attempts to move the control column to the rear. D. The tab moves down, so that less effort is required when the pilot attempts to move the control column to the rear.

Number: 6659 Question: Which statement about a jet transport aeroplane is correct during take-off with the CG at the forward limit and the trimmable horizontal stabiliser (THS) positioned at the maximum allowable aeroplane down position? A. Rotation will be normal. B. If the THS position is just within the limits of the green band, the take off warning system will be activated. C. The rotation will require extra stick force. D. Early nose wheel raising will take place.

6658 - A

6659 - C

Number: 6661 Question: Comparing the differences between a horizontal trimmable stabiliser and an elevator trim tab, which of these statements are correct or incorrect? I. A stabiliser trim is more suitable to cope with the large trim changes generated by the high lift devices on most jet transport aeroplanes. II. A stabiliser trim runaway causes less control difficulty. A. I is correct, II is correct. B. I is incorrect, II is correct. C. I is incorrect, II is incorrect. D. I is correct, II is incorrect.

Number: 6667 Question: The manoeuvring speed VA, expressed as indicated airspeed of a transport aeroplane: A. Depends on aeroplane mass and is independent of pressure altitude. B. Is independent of aeroplane mass, but dependent on pressure altitude. C. Depends on aeroplane mass and pressure altitude. D. Is a constant value.

VA - Maximum design manoeuvring speed and is the highest speed where the aircraft will stall before it exceeds the maximum load factor.

6661 - D

6667 - C

Number: 6668 Question: When flying at speeds above VA: A. The aircraft cannot be stalled. B. Full elevator deflection may result in damage to the airframe or structural failure. C. The aircraft may self-destruct in a turn. D. An overspeed warning will be activated.

Manoeuvres at speeds above point A therefore have the potential to cause permanent deformation to the structure or structural failure if the ultimate load is exceeded. This does not mean that any manoeuvre at a speed greater than point A will always cause structural damage; manoeuvres may be performed safely provided that the limit load factor is not exceeded.

There is of course a safety factor on the airframe of 1·5 so complete failure of the structure will not occur at the load factor of 2,5 but at 2,5 x 1,5 = 3,75.

6668 - B

Number: 6671 Question: By what percentage does VA (EAS) alter when the aeroplane's weight decreases by 19%? A. No change. B. 4.36% lower. C. 19% lower. D. 10% lower.

Number: 6673 Question: An aircraft flying at a given EAS is subjected to a positive gust of 50kt EAS. Which of the following correctly describes the increase in positive G felt by the aircraft? A. More at high aircraft weight. B. Less at altitude. C. More with a high aspect ratio straight wing. D. More with a swept wing.

Number: 6677 Question: Which of these statements about the gust load factor on an aeroplane are correct or incorrect? I. When the slope of the lift curve versus angle of attack curve decreases, the gust load factor decreases. II. When the wing loading decreases, the gust load factor decreases. A. I is correct, II is incorrect. B. I is incorrect, II is correct. C. I is correct, II is correct. D. I is incorrect, II is incorrect.

Vertical gusts will affect the load factor (n) by changing the angle of attack of the wing.

6671 - D

6673 - C

6677 - A

Number: 6678 Question: The twisting of a propeller blade from root to tip has been made to: A. To ensure its optimum thrust is always achieved at take off. B. To prevent excessive stress at the blade tip at high RPM. C. To provide its greatest thrust toward the blade root. D. Provide a constant angle of attack from root to tip.

Angle of Attack is the angle between the relative airflow and the blade section chord line.

The propeller blades twist from the hub to the tip (i. e. the blade angle reduces toward the tip) so that all blade sections along the entire The blades twist more at the root than the tip to maintain a constant angle of attack along the complete blade. Blade Angle = The angle between the chord line of each section and plane of rotation.

WIG ROOT BLADE ANGLE AGLE OF ATTACK

6678 - D

GROß CONSTANT

WING TIP KLEIN CONSTANT

Number: 6683 Question: For a given RPM of a fixed pitch propeller, the blade angle of attack will: A. Remain constant when the TAS increases. B. Increase when the TAS decreases. C. Remain constant when the TAS decreases. D. Decrease when the TAS decreases.

Angle of Attack is the angle between the relative airflow and the blade section chord line.

6683 - B

Number: 6686 Question: Increasing the camber on propeller blades will, if all else is the same: A. Increase the power absorption capability. B. Increase the propeller solidity. C. Increase the propeller efficiency. D. Give the aircraft greater range.

Number: 6691 Question: A propeller rotating anti-clockwise when viewed from the front, during the take-off ground roll will: A. Produce an increased load on the left wheel due to gyroscopic effect. B. Produce an increased load on the right wheel due to torque reaction. C. Produce an increased load on the left wheel due to torque reaction. D. Produce an increased load on the right wheel due to gyroscopic effect.

6686 - A

6691 - C

Number: 6694 Question: Counter rotating propellers have the effect of: A. Increasing the torque but decreasing the gyroscopic effect. B. Decreasing the torque but increasing the gyroscopic effect. C. Increasing the torque and gyroscopic effects. D. Cancelling out the torque and gyroscopic effects.

counter - entgegensetzen

Number: 6695 Question: In a single engine a/c with clockwise rotating propeller, a left yaw is generated due to: A. The torque effect. B. Higher helix angle. C. Higher lift on the right wing. D. The slipstream, striking the fin on the left side.

Number: 6697 Question: For a tail wheel aircraft with a right handed propeller, at the start of the take off run, asymmetric blade effect causes: A. No effect. B. Nose down pitch (tail up) Asymmetric Blade Effect C. Yaw to right. D. Yaw to left.

6694 - D

6695 - D

6697 - D

Number: 6699 Question: An aircraft in flight is affected by loads. These may be classified as: A. Tensile, shear, twisting and stretching. B. Thrust, drag, lift and weight. C. Compressive, tensile, shear and torsional. D. Compressive, bending, shear and torsional.

6699 - C

Number: 6700 Question: In order to climb with the speed for maximum climb rate, the aircraft should be flown with the IAS at which: A. The best ratio of speed / drag is reached. B. The best ratio of thrust / drag is prevailing. C. The best lift / drag ratio is prevailing. D. The power excess is maximal.

Vy - BEST RATE OF CLIMB

Vy - TURBOJET AEROPLANE

Vy - PROPELLER AEROPLANE

Vx - BEST ANGLE OF CLIMB

Vx - TURBOJET AEROPLANE

VMD VMP Blue Area

6700 - D

- Velocity for Minimum Drag - Velocity für Minimum Power Required - Exces Thrust available

Vx - PROPELLER AEROPLANE

Number: 6703 Question: Ignoring thrust effects in a steady straight climb at a climb angle 'gamma', the lift of an aeroplane with weight W is: A. W * (1-tan gamma) B. W * cos gamma C. W / cos gamma D. W * (1-sin gamma)

6703 - B

Number: 6704 Question: The maximum glide range of an aircraft will depend on wind and: A. Minimum Lift / Drag ratio. B. CL MAX. C. The ratio to lift to drag which varies according to angle of attack. D. Speed for minimum power required.

6704 - C

07-00 NEW QUESTIONS

PRINCIPLES OF FLIGHT

07-03 New Questions (20 Questions)

Number: 6707 Question: What action must the pilot take to maintain altitude and airspeed when turning in a jet aircraft? A. Increase angle of attack. B. Increase thrust. C. Decrease the turn radius. D. Increase angle of attack and thrust.

To maintain a steady balanced turn at a constant altitude, the greater the angle of bank, the greater the centripetal force and the greater the total lift requirement.

6707- D

Number: 6708 Question: In a steady turn at constant height: A. The radius of turn depends upon the weight and load factor. B. The rate of turn depends upon the weight, TAS and angle of bank. C. The rate of turn depends upon the TAS and angle of bank. D. The radius of turn depends only upon load factor.

Number: 6711 Question: During a correctly balanced turn: A. The lift force balances the aircraft weight. B. The lift force provides a centripetal force and a force that opposes the weight of the aircraft. C. The centrifugal force directly balances the weight of the aircraft. D. The thrust is a component of the centrifugal force.

6708 - C

6711 - B

Number: 6712 Question: Two identical aeroplanes A and B, with the same mass, are flying steady level co-ordinated 20° bank turns. If the TAS of A is 130 kt and the TAS of B is 200 kt: A. The turn radius of A is greater than that of B. B. The load factor of A is greater than that of B. C. The lift coefficient of A is less that that of B. D. The rate of turn of A is greater than that of B.

6712 - D

Number: 16157 Question: Which statement about propeller noise is correct? I. Propeller noise remains the same when the blade tip speed increases. II. For a given engine and propeller blade shape, a decrease in the number of propeller blades allows for a reduction in propeller noise. A. I is incorrect, II is incorrect. B. I is correct, II is correct. C. I is correct, II is incorrect. D. I is incorrect, II is correct.

Number: 16158 Question: Which statement is correct regarding the gyroscopic effect of a clockwise rotating propeller on a single engine aeroplane? I. Pitch up produces left yaw. II. Right yaw produces pitch up. A. I is correct, II is incorrect. B. I is incorrect, II is incorrect. C. I is correct, II is correct. D. I is incorrect, II is correct.

16157 - A

16158 - B

Number: 16161 Question: Considering subsonic incompressible airflow through a Venturi, which statement is correct? I. The dynamic pressure in the undisturbed airflow is the same as in the throat. II. The total pressure in the undisturbed airflow is higher than in the throat. A. I is incorrect, II is correct. B. I is correct, II is incorrect. C. I is incorrect, II is incorrect. D. I is correct, II is correct.

Number: 16163 Question: Considering subsonic incompressible airflow through a Venturi, which statement is correct? I. The static pressure in the undisturbed airflow is lower than in the throat. II. The speed in the undisturbed airflow is the same as in the throat. A. I is incorrect, II is incorrect. B. I is incorrect, II is correct. C. I is correct, II is incorrect. D. I is correct, II is correct.

16161 - C

16163 - A

Number: 16166 Question: 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 decreases. II. The strength of wing tip vortices decreases as the aspect ratio increases. A. I is correct, II is incorrect. B. I is incorrect, II is correct. C. I is correct, II is correct. D. I is incorrect, II is incorrect.

ASPECT RATIO

S = WING AREA, sq. m (b x c) b = SPAN, m c = AVERAGE CHORD, m AR = ASPECT RATIO AR = b/c AR = b²/S

16166 - C

Number: 16191 Question: Which statement, about an aeroplane leaving ground effect at constant angle of attack, is correct? I. The lift coefficient CL decreases. II. The induced drag coefficient CDi remains constant. A. I is correct, II is correct. B. I is correct, II is incorrect. C. I is incorrect, II is correct. D. I is incorrect, II is incorrect.

Ground effect occurs because the surface alters the airflow pattern around the wings. Primarily, the surface restricts the formation of the wingtip vortices. This results in a reduction in the amount of induced downwash behind the wing and increases the wing's effective angle of attack Upwash and downwash are reduced, causing the effective angle of attack of the wing to increase, (ref: Fig. 5.12). Therefore, when an aircraft is "in ground effect" lift will generally be increased and induced drag (CD) will be decreased. In addition, the reduced downwash will effect both longitudinal stability because of CP movement, and the pitching moment because of changes to the effective angle of attack of the tailplane, (Ref: Fig. 5.21).

16191 - B

Number: 16201 Question: Comparing the differences between a horizontal trimmable stabiliser and an elevator trim tab, which of these statements are correct or incorrect? I. The effects of a stabiliser trim runaway are more serious. II. A jammed stabiliser trim causes less control difficulty. A. I is correct, II is incorrect. B. I is incorrect, II is incorrect. C. I is correct, II is correct. D. I is incorrect, II is correct.

16201 - A

Number: 16218 Question: Which of these statements are correct or incorrect regarding a sideslip, with the relative airflow coming from the right, on an aeroplane that exhibits both directional and lateral stability? I. The initial tendency of the nose of the aeroplane is to move to the left. II. The initial tendency of the left wing is to move down. A. I is incorrect, II is incorrect. B. I is correct, II is incorrect. C. I is correct, II is correct. D. I is incorrect, II is correct.

Number: 728 Question: The effect of a high wing with zero dihedral is as follows: A. Positive dihedral effect B. Negative dihedral effect C. Zero dihedral effect D. Its only purpose is to ease aeroplane loading

16218 - D

728 - A

Number: 15390 Question: Assuming all bodies have the same cross-sectional area and are in motion, place these bodies in order of decreasing pressure drag. The correct answer is: A. 3, 4, 1, 2. B. 1, 2, 4, 3. C. 2, 1, 3, 4. D. 2, 1, 4, 3.

Number: 15299 Question: Assuming all bodies have the same cross-sectional area and are in motion, place these bodies in order of increasing pressure drag. The correct answer is: A. 2, 1, 4, 3. B. 3, 1, 4, 2. C. 4, 3, 2, 1. D. 3, 4, 1, 2.

15390 - D

15299 - D

Number: 6585 Question: The type of flap which extends rearward from the trailing edge of the wing as it is lowered is called: A. a fowler flap. B. a split flap. C. a zap flap. D. a Krueger flap.

Because of the combined effects of increased area and camber, the Fowler flap gives the greatest increase in lift of the flaps considered, and also gives the least drag because of the slot and the reduction of thickness : chord ratio.

6585 - A

Number: 6598 Question: The least energy loss through a normal shockwave occurs when the local Mach number is: A. Just below 1.0 M B. Well above 1.0 M The combined effect of the energy C. Just above 1.0 M loss across the shock wave, and D. Exactly 1.0 M the turbulent wake behind the shock wave, is called wave drag

Number: 6643 Question: Roll is: A. Rotation about the normal axis. B. Rotation about the longitudinal axis. C. Rotation about the longitudinal axis due to speed brake selection. D. Due to aileron deflection and is motion about the lateral axis.

6598 - C

6643 - B

Number: 6644 Question: When inner and outer ailerons are mounted, outer ailerons are used: A. During take off only. B. When flaps are in landing configuration only. C. At high speeds. D. At low speed.

Number: 6646 Question: In an aircraft fitted with spoilers for lateral control, and not deployed as speed brakes, a roll to the right is initiated by: A. Right spoiler extended, left spoiler retracted. B. Left spoiler extended, right spoiler retracted. C. Both spoilers extended. D. Right spoiler extended, but left spoiler extended further.

6644 - D

6646 - A

07-00 NEW QUESTIONS

PRINCIPLES OF FLIGHT

07-04 New Questions (20 Questions)

Number: 6649 Question: Aerodynamic balance can be obtained by: A. A weight mounted forward of the control surface hinge line. B. An internal balance (the leading edge of the aileron is housed within a box inside the wing trailing edge, vented to atmosphere, with a sea from the wing to the leading edge of the aileron. C. The down going aileron moving through a smaller angle than the up going aileron. D. An external balance, provided by a seal from the wing to the trailing edge of the aileron.

Number: 6653 Question: Which of the following is an advantage of engines mounted on the rear fuselage over those mounted in wing pods? A. Wings can have a lighter form of construction. B. The wing is less likely to suffer from flutter. C. Easier maintenance access. D. Longitudinal trim is less affected by changes in thrust.

6649 - B

6653 - D

Number: 6656 Question: Some airplanes have spring tabs mounted into the control system. This is to provide: A. Reduced control surface loads at all speeds. B. Constant spring tension to a trim system. C. A reduction in the pilots effort to move the controls against high air loads. D. Feel feedback in a control system.

Number: 6657 Question: Which kind of tab is commonly used in case of manual reversion of fully powered flight controls? A. Balance tab. B. Anti-balance tab. C. Spring tab. D. Servo tab. Servo Tab is directly controlled by the pilot through a pivot point and movement of the tab supplies the hinge moment necessary to move the main control surface. Movement of the tab causes the control surface to move to a new position of equilibrium in a direction of travel opposite to that of the tab (i.e. tab down, control surface up). In practice, the servo tab lacks effectiveness at low airspeeds when large control defiections are required. to lack - fehlen This is because the amount of airfiow passing over the tab is too low to produce the necessary hinge moment and hence the required defiection.

Spring Servo Tab overcomes the low-speed problems associated with a servo tab by including a spring box in the system. The spring tension is such that the tab does not come into operation until the stick force exceeds a predetermined value. At low airspeeds, the spring tension prevents movement of the servo tab and any control input by the pilot moves the control surface and tab as one piece. At higher airspeeds, the springs compress and the tab moves by way of the pivot point in the opposite direction to the control surface, providing the necessary aerodynamic assistance.

6656 - C

6657 - D

Number: 6660 Question: When a jet transport aeroplane takes off with the CG at the aft limit and the trimmable horizontal stabiliser (THS) is positioned at the maximum allowable nose up position for take-off: A. Rotation will be normal using the normal rotation technique. B. Early nose wheel raising will take place. C. There will be a tendency to under-rotate. D. Rotation will require higher than normal stick force.

Number: 6663 Question: Control surface flutter: A. Occurs at high angles of attack. B. Provides additional lift for take-off and landing in the event of engine failure. C. Is a destructive vibration that must be damped out within the flight envelope. D. Is a means of predicting the critical safe life of the wing. Flutter is caused by the combined effects of - changes in the pressure distribution around the control surface - with changing angles of attack (aerodynamic forces), and the forces due to the elastic nature of the aircraft structure itself (aeroelastic forces). If these forces become coincident and act in phase with each other, the resultant oscillations quickly increase in amplitude, and if left unchecked, may ultimately lead to structural failure.

To help eliminate fiutter in flight, manually operated control surfaces are generally mass balanced. Attaching weights forward of the hinge line brings the centre of gravity of the control surface to the hinge-line, thus altering the period of vibration and the liability to fiutter. These additional weights are usually installed internally along the leading edge of the control surface, inside the horn balance, or on an arm attached to the surface

6660 - B

6663 - C

Number: 6664 Question: Aileron reversal can be caused by: A. Fries type ailerons at low angles of attack. B. Neither A nor C. C. Both A and C. D. Twisting of the wing above reversal speed. AEROELASTIC DISTORTION (AILERON REVERSAL) During high-speed fiight, any aileron defiection may cause the wing to twist about its torsional axis. This is because the wing is flexible and the ailerons are near the wingtips, where the wings are less rigid. The actual torsional rigidity of a wing depends on its structure, but is normally strong enough to prevent any distortion at low airspeeds. Aileron power, however, increases as the square of forward airspeed, whereas the torsional stiffness of a wing remains constant with speed.

Number: 6665 Question: What is the relationship of VMO and MMO, in a climb and descent? A. If climbing at VMO, Mach number is decreasing. B. If climbing at VMO, it is possible to exceed MMO. C. If climbing at MMO, Indicated Airspeed is increasing. D. If descending at MMO, VMO cannot be exceeded.

VMO I MMO - Maximum operating speed limit. VMO is normally expressed as lAS. MMO is the same as VMO, but is stated as a Mach number and is the high altitude limiting speed since it is achieved before limiting lAS in the less dense air. In the climb and maintaining a constant lAS, the Mach number will increase, therefore MMO may be exceeded. Similarly, when descending at MMO, care should be taken not to exceed VMO.

6664 - D

6665 - B

Number: 6669 Question: On FAR 23 airplane, the limit load factor in normal category is: A. +3,8 G B. 4,4 G C. +6,0 G D. +3,2 G

CATEGORY

LIMIT LOAD

Normal 3.8 to –1.52 Utility (mild acrobatics, including spins) 4.4 to –1.76 Acrobatic 6.0 to –3.0 NORMAL - For airplanes with gross weight of more than 4,000 pounds, the limit load factor is reduced. To the limit loads given above, a safety factor of 50 percent is added.

Number: 6672 Question: An aircraft is flown at 20% below its normal weight. Because of this, VA will be: A. 20% higher. B. 20% lower. C. 10% higher. D. 10% lower. VA - Maximum design manoeuvring speed and is the highest speed where the aircraft will stall before it exceeds the maximum load factor.

Number: 6674 Question: Which of these statements about the gust load factor on an aeroplane are correct or incorrect? I. When the mass decreases, the gust load factor increases. II. When the altitude increases, the gust load factor increases. A. I is incorrect, II is correct. B. I is correct, II is correct. C. I is incorrect, II Is incorrect. D. I is correct, II is incorrect.

6669 - A

6672 - D

6674 - D

Number: 6675 Question: Which of these statements about the gust load factor on an aeroplane are correct or incorrect? I. When the slope of the lift versus angle of attack curve increases, the gust load factor decreases. II. When the wing loading increases, the gust load factor increases. A. I is incorrect, II is incorrect. B. I is correct, II is incorrect. C. I is correct, II is correct. D. I is incorrect, II is correct.

Number: 6676 Question: Which of these statements about the gust load factor on an aeroplane are correct or incorrect? I. When the slope of the lift versus angle of attack curve increases, the gust load factor increases. II. When the wing loading increases, the gust load factor increases. A. I is incorrect, II is correct. B. I is correct, II is correct. C. I is correct, II is incorrect. D. I is incorrect, II is incorrect.

6675 - A

6676 - C

Number: 6679 Question: With a fixed pitch propeller increasing speed will … propeller alpha and increasing power and therefore propeller RPM will … propeller alpha. A. decrease, decrease. During take-off, the angle of attack During cruise conditions, the angle of attack is B. increase, increase. is large because the airspeed is small and so forward airspeeds are limited to C. increase, decrease. low and the rotational velocity is prevent engine overspeed. D. decrease, increase. high.

Number: 6680 Question: A typical fixed pitch propeller (C-172) is designed to achieve its optimum angle of attack at: A. Cruise speed. B. Low forward speeds, such as during take off. C. Maximum speed for high performance. D. Rest to case engine starting.

6679 - D

6680 - A

Number: 6681 Question: A variable pitch propeller during take-off will move towards: A. Fine pitch to ensure minimum aerodynamic drag is generated. B. Coarse pitch to ensure the best angle of attack is achieved. C. Fine pitch to ensure that the engine can develop its maximum power. D. Coarse pitch to achieve the highest possible thrust.

Feathered When the chord line of the blade is parallel to the airflow, therefore preventing wind milling. Coarse Pitch The maximum cruising pitch in normal operation. Flight Fine Pitch The minimum pitch obtainable in flight. Ground Fine Pitch The minimum torque position for ground operation and is sometimes referred to as superfine pitch. Reverse Pitch An aerodynamic brake position used for braking and sometimes ground manoeuvring. It is achieved by accelerating air forward by the blade going into a negative angle. Alpha Range The fiight operating range, from flight fine pitch to coarse pitch. Beta Range From fiight fine pitch to reverse pitch which is the ground operating range and is hydromechanically controlled by a flight deck power lever.

6681 - C

Number: 6684 Question: The greatest drag produced by the variable pitch propeller on a piston engine will occur when the propeller is: A. Used during a powered glide. B. Stopped in coarse pitch. C. Windmilling. D. Stopped in fine pitch. WINDMILLING BLADE SECTION Reducing the rotational velocity (rpm) while maintaining the airspeed eventually causes the blade angle of attack to become negative.

When this occurs, the total reaction acts in a rearward direction and its components alter their orientation.

Number: 6685 Question: With a propeller feathered: A. There will be minimum drag on the propeller. B. The best windmilling speed is achieved. C. There will be minimum left to drag ratio. D. The engine will turn over just fast enough to lubricate it. Feathered When the chord line of the blade is parallel to the airflow, therefore preventing wind milling. Coarse Pitch The maximum cruising pitch in normal operation. Flight Fine Pitch The minimum pitch obtainable in flight. Ground Fine Pitch The minimum torque position for ground operation and is sometimes referred to as superfine pitch.

6684 - C

6685 - A

Number: 6687 Question: The number of blades in a propeller would be increased: A. To enable a longer undercarriage to be fitted. B. To reduce noise. C. To increase power absorption capability. D. To increase the efficiency of the variable pitch mechanism.

A further consideration is the number and the shape of the blades used. Increasing the aspect ratio of the blades reduces drag but the amount of thrust produced depends on blade area, so using high-aspect blades can result in an excessive propeller diameter. A further balance is that using a smaller number of blades reduces interference effects between the blades, but to have sufficient blade area to transmit the available power within a set diameter means a compromise is needed. Increasing the number of blades also decreases the amount of work each blade is required to perform, limiting the local Mach number - a significant performance limit on propellers. When the airflow over the tip of the blade reaches its critical speed, drag and torque resistance increase rapidly and shock waves form creating a sharp increase in noise. Aircraft with conventional propellers, therefore, do not usually fly faster than Mach 0.6. There have been propeller aircraft which attained up to the Mach 0.8 range, but the low propeller efficiency at this speed makes such applications rare.

Number: 6688 Question: The more blades a propeller has, the more power it is able to absorb. The limitation on blade number from an aerodynamic standpoint is: A. Engine speed if the engine is not geared. B. The blade diameter as compared to the maximum width. C. The loss of efficiency as the propeller tip approaches sonic speed. D. The loss of efficiency of one blade if it follows to the path of the preceding blade to closely. A further consideration is the number and the shape of the blades used. Increasing the aspect ratio of the blades reduces drag but the amount of thrust produced depends on blade area, so using high-aspect blades can result in an excessive propeller diameter. A further balance is that using a smaller number of blades reduces interference effects between the blades, but to have sufficient blade area to transmit the available power within a set diameter means a compromise is needed. Increasing the number of blades also decreases the amount of work each blade is required to perform, limiting the local Mach number - a significant performance limit on propellers. When the airflow over the tip of the blade reaches its critical speed, drag and torque resistance increase rapidly and shock waves form creating a sharp increase in noise. Aircraft with conventional propellers, therefore, do not usually fly faster than Mach 0.6. There have been propeller aircraft which attained up to the Mach 0.8 range, but the low propeller efficiency at this speed makes such applications rare.

6687 - C

6688 - D

07-00 NEW QUESTIONS

PRINCIPLES OF FLIGHT

07-05 New Questions (20 Questions)

Number: 6689 Question: The propeller noise can be minimised by: A. Reduce the propeller area. B. Reduce the RPM of the engine. C. Decrease the angle of attack of the propeller. D. Increase number of blades.

Number: 6692 Question: Which of the following would change the magnitude of the gyroscopic precession effect of the propeller? A. Rate of roll. B. Propeller blade angle. C. TAS. D. Propeller RPM.

Number: 6693 Question: The first action in the event of propeller runaway (overspeed conditions), should be to: A. Feather the propeller. B. Reduce the RPM lever setting. C. Close the throttle. D. Push the RPM. Lever fully forward.

Number: 6698 Question: The four forces of lift, weight, thrust and drag in level flight act through: A. The C of P. B. The Aft limit. C. The Aerodynamic Centre. D. The C of G.

6689 - D

6692 - D

6693 - C

6698 - D

Number: 6702 Question: In a steady climb: A. Thrust equals the weight component along the flight path and lift equals the sum of the components of drag and weigh along the flight path. B. If the angle of climb is 20°, lift equals weight times sin (20°). C. Thrust equals drag plus the weight component perpendicular to the flight path and lift equals the weight component along the flight path. D. Thrust equals drag plus the weight component along the flight path and lift equals the weight component perpendicular to the flight path.

Number: 6709 Question: In a steady banked turn the lift will: A. Equal the weight. B. Equal the resultant of weight and centrifugal force. C. Equal the centrifugal force. D. Equal the centrifugal force minus the weight.

6702 - D

6709 - B

Number: 6713 Question: In a steady co-ordinated horizontal turn, lift is: A. Equal to the aeroplane weight. B. Greater than in straight and level flight because it must balance the centripetal force. C. Greater than in straight and level flight, because it must balance the weight and generate the centripetal force. D. greater than in straight and level flight, because it must generate the centrifugal force.

Number: 6714 Question: Which are the two most important parameters to determine the value of VMCG? A. Engine thrust and rudder deflection. B. Engine thrust and flap setting. VMCG - Minimum Control Speed – Ground C. Air density and rudder deflection. minimale Kontrollgeschwindigkeit beim Startlauf am Boden, wenn das Flugzeug den Start fortsetzten soll – gilt nicht für leichte D. Engine thrust and air density. Zweimots

67113 - C

6714 - A

Number: 16128 Question: Which statement about propeller noise is correct? I. Propeller noise decreases when the blade tip speed increases. II. For a given engine and propeller blade shape, an increase in the number of propeller blades allows for a reduction in propeller noise. A. I is correct, II is correct. B. I is incorrect, II is incorrect. C. I is correct, II is incorrect. D. I is incorrect, II is correct.

Number: 16159 Question: Which statement is correct for a propeller of given diameter and at constant RPM? I. Assuming blade shape does not change power absorption is independent of the number of blades. II. Power absorption decreases if the mean chord of the blades increases. A. I is incorrect, II is correct. B. I is correct, II is incorrect. C. I is correct, II is correct. D. I is incorrect, II is incorrect.

Number: 16209 Question: Which of these statements about the gust load factor on an aeroplane are correct or incorrect? I. When the mass increases, the gust load factor decreases. II. When the altitude decreases, the gust load factor decreases. A. I is incorrect, II is incorrect. B. I is correct, II is correct. C. I is correct, II is incorrect. D. I is incorrect, II is correct.

16128 - D

16159 - D

16209 - C

Number: 6648 Question: The following is true concerning a balance tab. It is: A. only operating at high speed. B. a form of aerodynamic balance. C. a form of mass balance. D. used to increase feel. Balance Tabs They are connected to the tailplane by a mechanical linkage that causes them to move in the opposite direction to the control surface

Inset Hinge

The inset hinge places the hinge-line inside the control surface. This reduces the length of the moment arm and therefore the size of the hinge moment, thus reducing the overall stick force. This is known as control surface overbalance,

6648 - B

Horn Balance

Number: 6651 Question: When flutter damping of control surfaces is obtained by mass balancing, these weights will be located with respect to the hinge of the control surface: A. Behind the hinge. B. Above the hinge. C. Below the hinge. D. In front of the hinge. Flutter is caused by the combined effects of - changes in the pressure distribution around the control surface - with changing angles of attack (aerodynamic forces), and the forces due to the elastic nature of the aircraft structure itself (aeroelastic forces). If these forces become coincident and act in phase with each other, the resultant oscillations quickly increase in amplitude, and if left unchecked, may ultimately lead to structural failure. To help eliminate fiutter in flight, manually operated control surfaces are generally mass balanced. Attaching weights forward of the hinge line brings the centre of gravity of the control surface to the hinge-line, thus altering the period of vibration and the liability to fiutter. These additional weights are usually installed internally along the leading edge of the control surface, inside the horn balance, or on an arm attached to the surface

Number: 6662 Question: Wing flutter may be caused by a: A. Combination of roll control reversal and low speed stall. B. Combination of bending and torsion of the wing structure. C. Combination of fuselage bending and wing torsion. D. Aerodynamic wing stall at high speed. Flutter is caused by the combined effects of - changes in the pressure distribution around the control surface - with changing angles of attack (aerodynamic forces), and the forces due to the elastic nature of the aircraft structure itself (aeroelastic forces). If these forces become coincident and act in phase with each other, the resultant oscillations quickly increase in amplitude, and if left unchecked, may ultimately lead to structural failure.

6651 - D

6662 - B

Number: 6670 Question: If an aircraft is flying at a speed above VA: A. The aircraft can not stall. B. The airplane can destroy itself when in a bank. C. An overspeed warning will be indicated. D. A full elevator deflection could cause a structural damage.

VA - Maximum design manoeuvring speed and is the highest speed where the aircraft will stall before it exceeds the maximum load factor.

Number: 6682 Question: A fixed pitch propeller is usually: A. At its optimum angle on take off. B. At too fine and angle of take off. C. At too coarse an angle in the cruise. D. At too course of an angle of take off. Feathered When the chord line of the blade is parallel to the airflow, therefore preventing wind milling. Coarse Pitch The maximum cruising pitch in normal operation. Flight Fine Pitch The minimum pitch obtainable in flight. Ground Fine Pitch The minimum torque position for ground operation and is sometimes referred to as superfine pitch.

Number: 6696 Question: A propeller rotating clockwise as seen from the rear tends to rotate the aircraft to the: A. Left around the vertical axis, and to the right around the longitudinal axis. B. Right around the vertical axis, and to the right around the longitudinal axis. C. Right around the vertical axis, and to the left around the longitudinal axis. D. Left around the vertical axis, and to the left around the longitudinal axis.

6670 - D

6682 - D

6696 - D

Number: 6701 Question: Flying at the maximum rate of climb speed (Vy) you will obtain maximum: A. Altitude in the shortest time. B. Altitude at maximum boost setting. C. Altitude in the shortest distance. D. Altitude in the shortest distance and time.

Number: 6705 Question: During the glide, the forces acting on an aircraft are: A. Lift, drag and weight. B. Drag, thrust and weight. C. Lift, weight and thrust. D. Thrust, lift and drag.

Number: 6710 Question: In a level banked turn, the stalling speed will: A. Decrease. B. Increase. C. very inversely with wing loading. D. Remain the same.

6701 - A

6705 - A

6710 - B

07-00 NEW QUESTIONS

PRINCIPLES OF FLIGHT

07-06 New Questions (96 Questions)

Number: 6576 Question: A wing stalling angle: A. Decreased in any turn. B. Increased in a high rate of turn. C. Unaffected by a turn. D. Decreased in a high rate of turn.

Most light aircraft tend to stall when the wing reaches an angle of attack of approximately 15 - 16° in any phase of night, regardless of the airspeed, provided that the aircraft configuration is not altered

Number: 6601 Question: In supersonic flight, any disturbance around a body affects the flow only: A. outside the Mach cone. MACH CONE B. in front of the normal shockwave. Only the region behind the oblique shock wave is affected by C. In front of the body. disturbances and is sometimes referred to as the zone of action. D. within the Mach cone.

The region ahead of the oblique shock wave is not affected by the disturbances and is called the zone of silence. In three dimensions, the disturbances emanating from the moving body expand outward as spheres and not circles. When the speed is above Mach 1, these spheres are enclosed within a cone, called the Mach cone and it is within the Mach cone that disturbances are felt.

6576 - C

6601 - D

Number: 6561 Question: The continuity equation states: If the area of a tube is increasing, the speed of the subsonic and incompressible flow inside is: A. Not changing. B. Increasing. C. Decreasing. D. Sonic.

Number: 6578 Question: What is the percentage increase in stall speed in a 45° bank turn? A. 41% B. 10% C. 45% D. 19%

6561 - C

6578 - D

Number: 12957 Question: What is the approximate value of the lift of an aeroplane at a gross weight of 50000 N, in a horizontal co-ordinated 45 degrees banked turn? A. 80000 N. By what percentage does the lift increase in a steady level turn at 45o of bank, compared to B. 60000 N. straight and level: C. 70000 N. D. 50000 N. 1 ÷ cos 45 = 1.41 = 41% 50,000N x 1.41 = 70500

12957 - C

Number: 6594 Question: If ice is present on the leading edge of the wings, it may increase the landing distance due to a higher Vth with: A. 30-40% B. 5-10% C. 10-20% D. 40-50%

Number: 6614 Question: Dynamic longitudinal stability requires: A. An effective elevator. B. A small CG range. C. A variable incidence (trimming) tailplane. D. Positive static longitudinal stability.

6594 - D

6614 - D

Number: 208 Question: The relationship between the stall speed VS and VA (EAS) for a large transport aeroplane can be expressed in the following formula: (SQRT= square root) A. Va= VA SQRT(3.75) B. VS= VA SQRT(3.75) VS - Unaccelerated stall speed in the clean configuration C. VS= VA SQRT(2.5) D. VA= VS SQRT(2.5) VA - Maximum design manoeuvring speed and is the highest speed where the aircraft will stall before it exceeds the maximum load factor.

Number: 6575 Question: When an aeroplane is in ground effect: A. Drag and lift are both reduced. B. Drag is decreased, lift is increased. C. Drag is increased, lift is decreased. D. Drag and lift are both increased.

Number: 6570 Question: Which statement is correct about the CI and angle of attack? A. For an asymmetric aerofoil with positive camber, if angle of attack is greater than 0, CI = 0 B. For a symmetric aerofoil, if angle of attack = 0, CI = 0 C. For an asymmetric aerofoil, if angle of attack = 0, CI = 0 D. For a symmetric aerofoil, if angle of attack = 0, CI is not equal to 0

208 - D

6575 - B

6570 - B

Number: 16657 Question: The correct sequence of cross-sections representing propeller blade twist is: A. sequence 1. B. sequence 4. C. sequence 3. D. sequence 2.

The propeller blades twist from the hub to the tip (i. e. the blade angle reduces toward the tip) so that all blade sections along the entire The blades twist more at the root than the tip to maintain a constant angle of attack along the complete blade. Blade Angle = The angle between the chord line of each section and plane of rotation.

WIG ROOT BLADE ANGLE AGLE OF ATTACK

16657 - A

GROß CONSTANT

WING TIP KLEIN CONSTANT

Number: 6572 Question: What is the effect of increasing wing aspect ratio on induced drag? A. It is increased because high aspect ratio has greater frontal area. B. It is unaffected because there is no relation between aspect ratio and induced drag. C. It is reduced because the effect of wing-tip vortices is reduced. D. It is increased because high aspect ratio produces greater downwash.

Number: 6620 Question: Moving the C of G of an aircraft aft inflight will: A. Increase longitudinal stability. B. Increase the angle of attack. C. Reduce longitudinal stability. D. Have no effect on longitudinal stability.

Number: 6630 Question: Which of the following features will reduce lateral stability? A. Anhedral of the wing. B. Tip tanks. C. Dihedral of the wing. D. Streamlining the wing root.

Number: 6639 Question: Rudder controls: A. Pitch B. Roll C. Yaw D. Turn

6572 - C

6620 - C

6630 - A

6639 - C

Number: 6610 Question: Reducing the thickness/chord ratio on a wing will: A. Delay the onset of shockwave formation. B. All of the above. C. Reduce the transonic variations in lift coefficient. D. Reduce the transonic variations in drag coefficient. Wing section design features used to increase MCRIT include: » Low thickness/chord ratio

» Maximum thickness well aft » Small leading edge radius of curvature With increasing Mach number, the flow over the wing continues to accelerate and eventually reaches a sonic value at a particular point on the wing, normally the point of maximum thickness

Number: 16310 Question: Which of these statements about weight or mass is correct? A. In the SI system the unit of measurement for weight is the kilogram. B. The weight of an object is independent of the acceleration due to gravity. C. The mass of an object depends on the acceleration due to gravity. D. In the SI system the unit of measurement for weight is the Newton.

6610 - B

16310 - D

Number: 15826 Question: A typical curve representing propeller efficiency of a fixed pitch propeller versus TAS at constant RPM is: A. diagram 3. B. diagram 2. C. diagram 1. D. diagram 4.

Maximum efficiency occurs at one airspeed and one particular blade angle of attack.

The highest efficiency obtained by a propeller is 85% to 88%. The blade angle is usually set so that the speed for maximum efficiency is close to the cruising speed. At any other airspeed, the efficiency is relatively low and only a small proportion of the power being delivered by the engine is used to propel the aircraft.

15826 - D

Number: 6611 Question: Compared to a normal transonic airfoil section a supercritical section has: A. A flatter top surface. B. A flatter bottom surface. C. A more cambered top surface. D. A very sharp leading edge.

Wing section design features used to increase MCRIT include: » Low thickness/chord ratio

» Maximum thickness well aft » Small leading edge radius of curvature With increasing Mach number, the flow over the wing continues to accelerate and eventually reaches a sonic value at a particular point on the wing, normally the point of maximum thickness

6611 - A

Number: 6562 Question: Bernoulli's equation is: (note: rho is density, Pstat is static pressure, Pdyn is dynamic pressure, Ptot is total pressure) A. Pdyn + 1/2 * rho * IAS^2 = constant B. Ptot + 1/2 * rho * TAS^2 = constant C. Pstat + 1/2 * rho *TAS^2 = constant D. Pstat + 1/2 * rho * IAS^2 = constant

BERNOULLI'S THEOREM

Number: 16655 Question: The correct sequence of cross-sections representing propeller blade twist is: A. sequence 1. B. sequence 3. C. sequence 4. D. sequence 2.

The propeller blades twist from the hub to the tip (i. e. the blade angle reduces toward the tip) so that all blade sections along the entire length operate at the same angle of attack. The blades twist more at the root than the tip to maintain a constant angle of attack along the complete blade. Blade Angle = The angle between the chord line of each section and plane of rotation.

6562 - C

16665 - D

Number: 16775 Question: Which of these statements about the pitching moment coefficient versus angle of attack lines in the annex is correct? A. The CG position is further aft at line 4 when compared with line 1. B. In its curved part line 2 illustrates a decreasing static longitudinal stability at high angles of attack. C. Static longitudinal stability is greater at line 3 when compared with line 4 at low and moderate angles of attack. D. The CG position is further forward at line 4 when compared with line 1.

16775 - D

Number: 6613 Question: If an aircraft has static longitudinal instability, it: A. May or may not be dynamically stable, depending on momentum and damping factors. B. Will be dynamically stable. C. Will be dynamically unstable. D. Will be dynamically stable only at a low speed.

6613 - C

Number: 6609 Question: Compared to straight wings of the same airfoil section swept wings … the onset of the transonic drag rise and have a … CD in supersonic flight: A. delay, lower B. hasten, lower C. delay, higher D. hasten, higher Wing Planform has the most significant effect on MCRIT. Careful design not only delays the shock stall, but also significantly reduces the severity when it occurs. If a wing has sweep back, the effective chord (parallel to the aircraft's longitudinal axis) is lengthened, but the wing's thickness remains unchanged.

Number: 6636 Question: When ice is present on the stabilizer, deflection of flaps may cause: A. The stabilizer to stall and a pitch up situation. B. The stabilizer to stall and a vertical dive. C. A roll movement due to directional instable. D. Both wings and stabilizer to stall.

Number: 6564 Question: Consider a steady flow through a stream tube at a given constant velocity. An increase in the flow's temperature will: A. Increase the mass flow. B. Increase the mass flow when the tube is divergent in the direction of the flow. C. Not affect the mass flow. D. Decrease the mass flow.

6609 - C

6636 - B

6564 - D

Number: 6579 Question: A boundary layer fence on a swept wing will improve the: A. Mach buffet characteristics. B. Low-speed stall characteristics. C. Lift coefficient of the trailing edge flap. D. Dutch roll characteristics.

Number: 6621 Question: The effects of CG position on longitudinal static stability and control response will be: A. Rearward movement of the CG will reduce stability and control response. B. Forward movement of the CG will reduce control response and increase stability. C. Rearward movement of the CG will increase stability and reduce control response. D. Forward movement of the CG will reduce stability and increase control response.

Number: 6569 Question: Assuming no flow separation and no compressibility effects the location of the centre of pressure of a positively cambered aerofoil section: A. Is at approximately 25% chord irrespective of angle of attack. B. Moves forward when the angle of attack increase. C. Is independent of angle of attack. D. Moves backward when the angle of attack increases.

6579 - B

6621 - B

6569 - B

Number: 6617 Question: If the sum of moments in flights is not zero, the aeroplane will rotate about: A. The aerodynamic centre of the wing. B. The centre of gravity. C. The centre of pressure of the wing. D. The neutral point of the aeroplane.

Number: 6619 Question: The purpose of the horizontal stabilizer is to: A. Give the aeroplane sufficient lateral stability. B. Give the aeroplane sufficient directional stability. C. Give the aeroplane enough weight in the tail. D. Give the aeroplane sufficient longitudinal stability.

Number: 6603 Question: For minimum wave drag, an aircraft should be operated at which of the following speeds? A. Mach 1.0 B. Subsonic. C. Low supersonic. D. High supersonic.

Number: 6602 Question: At higher altitudes, the stall speed (IAS): A. Decreases. B. Remains the same. C. Increases. D. Decreases until the tropopause.

6617 - B

6619 - D

6603 - B

6602 - C

Number: 6574 Question: 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 increases as the angle of attack increases. II. The strength of wing tip vortices decreases as the aspect ratio decreases. A. I is correct, II is incorrect. B. I is incorrect, II is incorrect. C. I is incorrect, II is correct. D. I is correct, II is correct.

6574 - A

Number: 6590 Question: CLMAX may be increased by the use of: A. flaps. B. slats. C. boundary layer control. D. A, B and C.

Number: 6592 Question: Speed brakes are a device used on large transport category aircraft: A. to increase drag in order to maintain a steeper gradient of descent. B. used at high speeds for turning when a yaw damper is not installed. C. for speed reduction after landing. D. and are an old version of anti block system.

6590 - D

6592 - A

Number: 6616 Question: If the aircraft is properly loaded the CG, the neutral point and the manoeuvre point will be in the order given, forward to aft: A. CG, manoeuvre point, neutral point. B. Manoeuvre point, neutral point, CG. C. Manoeuvre point, CG, neutral point. D. CG, neutral point, manoeuvre point.

Number: 6568 Question: A pitch up could be caused by: A. Forward movement of the centre of pressure. B. Lateral movement of the centre of gravity. C. A reduction in varying loads due to G. D. Forward movement of the centre of gravity.

Number: 6573 Question: 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 decreases. II. The strength of wing tip vortices increases as the aspect ratio increases. A. I is correct, II is incorrect. B. I is incorrect, II is incorrect. C. I is incorrect, II is correct. D. I is correct, II is correct.

6616 - D

6568 - A

6573 - A

Number: 6581 Question: Regarding deep stall characteristics, identify whether the following statements are correct or incorrect: I. A wing with forward sweep and a low horizontal tail makes an aeroplane prone to deep stall. II. A stick shaker system is fitted to an aeroplane to resolve deep stall problems. A. I is correct, II is correct. B. I is incorrect, II is incorrect. C. I is incorrect, II is correct. D. I is correct, II is incorrect.

6581 - B

Number: 6624 Question: The C.G. position of an aeroplane is forward of the neutral point in a fixed location. Speed changes cause a departure from the trimmed position. Which of the following statements about the stick force stability is correct? A. Aeroplane nose up trim decreases the stick force stability. B. Increase of speed generates pull forces. C. Stick force stability is not affected by trim. D. Increasing 10kt trimmed at low speed has more effect on the stick force than increasing 10kt trimmed at high speed.

Number: 6583 Question: Flap selection at constant IAS in straight and level flight will increase the: A. Maximum lift coefficient (CLmax) and the drag. B. Lift coefficient and the drag. C. Stall speed. D. Lift and the drag.

Number: 935 Question: Which of the following statements is correct? A. Dynamic stability means that after being displaced from original equilibrium condition, the aeroplane will return to that condition without oscillation. B. A dynamically stable aeroplane would be almost impossible to fly manually. C. Dynamic stability is possible only when the aeroplane is statically stable about the relevant axis. D. Static stability means that the aeroplane is also dynamically satble about the relevant axis.

6624 - D

6583 - A

935 - C

Number: 6604 Question: As an aircraft accelerates through the transonic speed range: A. The coefficient of drag decreases then increases. B. The coefficient of drag increases then decreases. C. The coefficient of drag increases. D. The coefficient of drag decreases.

Number: 6625 Question: A Mach trimmer: A. Increases the stick force per g at high Mach numbers. B. Corrects insufficient stick force stability at high Mach numbers. C. Has no effect on the shape of the elevator position versus speed (IAS) curve for a fully hydraulic controlled aeroplane. D. Is necessary for compensation of the autopilot at high Mach Numbers.

Number: 6631 Question: Pitch is movement around the: A. Lateral axis. B. Vertical axis. C. Longitudinal axis. D. Yaw axis.

6604 - B

6625 - B

6631 - A

Number: 6565 Question: The angle between the direction of the undisturbed airflow (relative wind) and the chord line of an aerofoil is the: A. glide path angle. B. climb path angle. C. angle of attack. D. same as the angle between chord line and fuselage axis.

Number: 7220 Question: The aspect ratio of the wing: A. Is the ratio between the tip chord and the wing span. B. Is the ratio between the wing span and the root chord. C. Is the ratio between chord and root chord. D. Is the ratio between the wing span and the mean geometric chord.

6565 - C

7220 - D

Number: 6615 Question: The axes of an aircraft by definition must all pass through the: A. Centre of gravity. B. Flight deck. C. Centre of pressure. D. Aircraft datum.

Number: 6600 Question: When an aircraft is flying at speeds above Mach 1, pressure disturbances from the aircraft will be felt only: A. in front of the oblique shockwave. MACH CONE Only the region behind the oblique shock wave is affected by B. in front of the normal shockwave. disturbances and is sometimes referred to as the zone of action. C. within the Mach cone. The region ahead of the oblique shock wave is not affected by the D. in front of the Mach cone. disturbances and is called the zone of silence.

In three dimensions, the disturbances emanating from the moving body expand outward as spheres and not circles. When the speed is above Mach 1, these spheres are enclosed within a cone, called the Mach cone and it is within the Mach cone that disturbances are felt.

6615 - A

6600 - C

Number: 6608 Question: Compared to a straight wing of the same airfoil section a wing swept at 30 should theoretically have an Mcrit … times Mcrit for the straight wing, but will, in practice gain … that increase: A. 1.414; twice B. cosine 30; twice C. sine 30; half D. 1.154; half

Number: 6634 Question: When the control column is moved forward and to the right: A. The elevator goes down, the starboard aileron moves down and the port aileron moves up. B. The elevator goes up, the starboard aileron moves down and the port aileron moves up. C. The elevator goes down, the starboard aileron moves up and the port aileron moves down. D. The elevator goes up, the starboard aileron moves up and the port aileron moves down.

Number: 6612 Question: In the transonic range CLmax will … and the 1g stalling speed will … A. increase, decrease. B. decrease, increase. C. decrease, decrease. D. increase, increase.

6608 - D

6634 - C

6612 - B

Number: 6599 Question: On a typical transonic airfoil the transonic rearward shift of the CP occurs at about: A. M 0.75 to M 0.98 B. M 0.75 to M 0.89 C. M 0.89 to M 0.98 D. M 0.91 to M 1.4

Number: 15822 Question: A typical curve representing propeller efficiency of a fixed pitch propeller versus TAS at constant RPM is: A. diagram 1. B. diagram 2. C. diagram 4. D. diagram 3.

Maximum efficiency occurs at one airspeed and one particular blade angle of attack.

The highest efficiency obtained by a propeller is 85% to 88%. The blade angle is usually set so that the speed for maximum efficiency is close to the cruising speed. At any other airspeed, the efficiency is relatively low and only a small proportion of the power being delivered by the engine is used to propel the aircraft.

6599 - C

15822 - A

Number: 6629 Question: What happens to lateral stability when flaps are extended? A. Lateral stability is increased as the centre of pressure moves inboard. B. Lateral stability is unaffected, as the wings are symmetrical. C. Lateral stability is decreased. D. Lateral stability is increased as lift is increased.

Number: 6638 Question: An aircraft is approaching to land with its CG at the forward limit. It will be … to flare and VREF will be … than normal. A. Easy; lower. B. Difficult; higher. C. Difficult; lower. D. Easy; higher.

Number: 6593 Question: Which of the following is the most important result/problem caused by ice formation? A. Increased weight. B. Increased drag. C. Reduction in CLmax. D. Blockage of the controls.

6629 - C

6638 - B

6593 - C

Number: 6605 Question: The critical speed where the speed is too low and too high at the same time is called: A. Coffin corner. B. High speed buffet. The point at which the two stalls coincide is often referred to as the coffin corner. C. Mcrit. D. Supersonic.

Number: 6587 Question: flaps are used in order to: A. Increase max L/D. B. Reducing drag. C. Increase max lift coefficient by increasing max angle of attack. D. Decrease stalling speed and reduce max angle of attack thereby achieving a more nose down altitude near and at stalling speed.

6605 - A

6587 - D

Number: 6566 Question: When airflow over a wing becomes supersonic, the pressure pattern on the top surface will become? A. Triangular. B. Irregular. C. The same as subsonic. D. Rectangular.

Number: 16338 Question: Which of these statements about stall speed is correct? A. Increasing forward sweep decreases stall speed. B. Decreasing wing anhedral decreases stall speed. C. Increasing forward sweep increases stall speed. D. Increasing wing anhedral decreases stall speed.

6566 - B

16338 - C

Number: 15828 Question: A typical curve representing propeller efficiency of a fixed pitch propeller versus TAS at constant RPM is: A. diagram 3. B. diagram 2. C. diagram 4. D. diagram 1.

Maximum efficiency occurs at one airspeed and one particular blade angle of attack.

The highest efficiency obtained by a propeller is 85% to 88%. The blade angle is usually set so that the speed for maximum efficiency is close to the cruising speed. At any other airspeed, the efficiency is relatively low and only a small proportion of the power being delivered by the engine is used to propel the aircraft.

15828 - A

Number: 6637 Question: What is the effect on the aeroplane's static longitudinal stability of a shift of the centre of gravity to a more aft location and on the required control deflection for a certain pitch up or down? A. The static longitudinal stability is smaller and the required control deflection is larger. B. The static longitudinal stability is larger and the required control deflection is smaller. C. The static longitudinal stability is larger and the required control deflection is larger. D. The static longitudinal stability is smaller and the required control deflection is smaller.

Number: 6618 Question: If an airplane has poor longitudinal stability in flight, what can be done to increase the stability? A. Increase stabiliser surface area. B. Install a yaw damper. C. Reduce in keel surface area. D. Increase elevator range of movement.

66337 - D

6618 - A

Number: 6626 Question: Longitudinal dynamic oscillation takes two forms. One of these, long period oscillation, involves slow changes in: A. Height and load factor. B. Height and speed. C. Speed and load factor. D. Pitch and load factor. Long Period Oscillation (Phugoid) This involves very long periods of oscillation (20 - 100 sec) with noticeable variations in pitch attitude, altitude, and airspeed, whilst the angle of attack remains nearly constant (i .e. if the aircraft experiences a horizontal gust, its airspeed momentarily changes, but its angle of attack remains virtually constant). Any change in airspeed is accompanied by a change in drag. Short Period Oscillation This involves very short periods of oscillation, typically 1-2 sec, when an aircraft is subjected to a vertical gust. The disturbance causes the aircraft to rotate about its lateral axis, and varies its angle of attack, whilst the airspeed remains virtually constant. The change in angle of attack also varies the lift, resulting in a pitching moment. If the aircraft is statically longitudinally stable, any disturbance in pitch sets up an oscillatory motion about the aircraft's lateral axis, where oscillation is dynamically stable or unstable.

ALTITUDE During a phugoid altitude varies significantly, but during a short period oscillation it remains approximately constant.

Phugoid A phugoid or fugoid (play /ˈfjuːɡɔɪd/) is an aircraft motion where the vehicle pitches up and climbs, and then pitches down and descends, accompanied by speeding up and slowing down as it goes "uphill" and "downhill." This is one of the basic flight dynamics modes of an aircraft (others include short period, dutch roll, and spiral divergence), and a classic

SPEED During a phugoid the speed varies significantly, whereas during a short period oscillation it does not.

long period oscillation

short period oscillation

SPEED

VARIES

CONSTANT

ANGLE OF ATTACK

CONSTANT

VARIES

ALTITUDE

VARIES

CONSTANT

6626 - B

Number: 6597 Question: transonic speed is: A. A speed at which locally around the aeroplane both supersonic and subsonic speeds exist. B. The speed range between Mcrit and Mmo. C. A speed at which locally an oblique shock wave has developed in the flow along the aeroplane. D. A speed at which compressibility effect are first noticeable. In subsonic flight the total airflow around an aircraft is travelling at a speed less than the speed of sound. This includes speeds of approximately Mach 0.75 or less. Transonic flight occurs at speeds between Mach 0.75 and Mach 1.2, where the airflow around an aircraft is partly subsonic, and partly supersonic. Supersonic flight occurs at speeds between Mach 1.2 and Mach 5.0, where the total airflow around an aircraft is travelling at a speed greater than the speed of sound.

Number: 6589 Question: Which of the following occurs when trailing edge flaps are extended? A. The critical angle of attack is constant, but CLMAX increases. B. The critical angle of attack remains constant and stall speed increases. C. CLMAX increases and the critical angle of attack increases. D. The critical angle of attack decreases and CLMAX increases.

6597 - A

6589 - D

Number: 16431 Question: Which statement is correct regarding the pitching moment coefficient Cm versus angle of attack diagram? A. Line 1 shows an aeroplane with increasing static longitudinal instability at very high angles of attack. B. Line 4 shows an aeroplane with a greater static longitudinal stability at low angles of attack than that shown in line 3. C. Line 4 shows an aeroplane with increasing static longitudinal stability at very high angles of attack. D. Line 3 shows an aeroplane with increasing static longitudinal stability at high angles of attack.

16431 - C

Number: 16433 Question: Which statement is correct regarding the pitching moment coefficient Cm versus angle of attack diagram? A. Line 3 shows an aeroplane with reducing static longitudinal stability at high angles of attack. B. Line 4 shows an aeroplane with greater static longitudinal stability at low angles of attack than that in shown line 3. C. Line 1 shows an aeroplane with increasing static longitudinally instability at very high angles of attack. D. Line 4 shows an aeroplane with reducing static longitudinal stability at very high angles of attack.

16433 - A

Number: 16777 Question: Which of these statements about the pitching moment coefficient versus angle of attack lines in the annex is correct? A. The CG position is further forward at line 1 when compared with line 4. B. Static longitudinal stability is greater at line 3 when compared with line 4 at low and moderate angles of attack. . C. In its curved part at high angles of attack line 2 illustrates a decreasing static longitudinal stability. D. The CG position is further aft at line 1 when compared with line 4.

16777 - D

Number: 6588 Question: When deploying the flaps the effective angle of attack: A. May increase or decrease depending on the aircraft type. B. Decreases. C. Increases. D. Remains the same. to deploy - anwenden, einsetzen

Number: 6623 Question: The effect of Mach trim on stick forces for power operated controls: A. A Mach trim system is not required if an aircraft has power operated controls. B. Is to decrease the stick force gradient to prevent the possibility of high speed stall. C. Is to decrease the stick force gradient to ensure the pilot can manoeuvre the aircraft adequately when flying at a high transonic Mach number. D. Is to maintain the required stick force gradient.

MACH TRIM To guard against nose tuck under, frequent pitch trim changes are required. This is carried out by a variable incidence tailplane, which is automatically positioned by a Mach trim system. This system is designed to aid aircraft longitudinal stability and ensures that the forward stick forces increase proportionally with increasing Mach number. It is operational at high Mach numbers in the transonic speed range.

6588 - C

6623 - D

Number: 16610 Question: Static lateral stability will be increased by: A. reducing wing sweepback. B. reducing wing anhedral. C. the use of a low, rather than high, wing mounting. D. reducing the size of the vertical tail.

16610 - B

Number: 6577 Question: The boundary layer is considered to be turbulent: A. Just in front of the transition point. B. Just aft of the separation point. The boundary layer is the layer of air between the surface and the free stream velocity in which local C. Just in front of the centre of pressure. retardation takes place. Like the main airflow, the D. Between the transition and separation points. boundary layer flow can be either laminar or turbulent in nature.

Number: 6560 Question: The units of density of the air (I) and the force (II) are: A. (I) kg / m³, (II) N. B. (I) N / m³, (II) N. C. (I) kg / m², (II) kg. D. (I) N / kg, (II) kg.

Number: 6627 Question: Lateral static stability is determined by: A. CG position. B. Aspect ratio. C. Wingspan. D. Aircraft response to sideslip. LATERAL STATIC STABILITY The static lateral stability of an ai rcraft is its natural, or inbuilt tendency to recover from a disturbance in roll. A disturbance in roll causes one wing to rise and the other to drop. The wings naturally damp out the motion in roll and the aircraft assumes a banked attitude.

6577 - D

6560 - A

6627 - D

Number: 6584 Question: Which statement is correct? A. Spoiler extension decrease the stall speed and the minimum rate of descent, but increase the minimum descent angle. B. Extension of flaps has no influence on the minimum rate of descent, as only the TAS has to be taken into account. C. Extension of flaps will increase (CL/CD)max, causing the minimum rate of descent to decrease. D. Extension of flaps causes a reduction of the stall speed, the maximum glide distance also reduces.

Number: 6596 Question: During a descent at a constant Mach number (assume zero thrust and standard atmospheric conditions): A. The descent angle will decrease. B. The TAS will decrease. C. The angle of attack will decrease. D. The pitch angle will increase.

6584 - D

6596 - C

Number: 6567 Question: Lowering the inboard flaps causes the wing Centre of Pressure: A. to move inward and forward. B. to move forward. C. to move inboard towards the wing root. D. to move outboard towards the wing tips.

Number: 6582 Question: Regarding deep stall characteristics, identify whether the following statements are correct or incorrect: I. The combination of a wing with sweepback and a T-tail make an aeroplane prone to deep stall. II. A stick shaker system is fitted to an aeroplane to resolve deep stall problems A. I is correct, II is correct. B. I is incorrect, II is correct. C. I is incorrect, II is incorrect. D. I is correct, II is incorrect.

6567 - C

6582 - D

Number: 6628 Question: The dihedral construction of an aircraft wing provides: A. Longitudinal stability about the lateral axis. B. Lateral stability about the normal axis. C. Directional stability about the lateral axis. D. Lateral stability about the longitudinal axis.

Number: 6580 Question: The input to a stick shaker comes from: A. Angle of attack, and sometimes the rate of change in angle of attack. B. The angle of incidence. C. The angle of attack only. D. The airspeed, and sometimes the rate of change in airspeed.

Number: 6595 Question: How does temperature influence the speed of sound? A. Speed of sound decreases with temperature increase. B. Speed of sound increases with temperature increase. C. Speed of sound is not influenced by temperature. D. Speed of sound remains constant.

6628 - D

6580 - A

6595 - B

Number: 16609 Question: Static lateral stability will be increased by: A. increasing wing anhedral. B. reducing the size of the vertical tail. C. the use of a low, rather than high, wing mounting. D. increasing wing sweepback.

Number: 6635 Question: A primary stop is mounted on an elevator control system in order to: A. Maintain constant control cable tension. B. Restrict the range of movement of the elevator. C. Restrict the range of movement of the control column. D. Prevent overloading of control cables.

Number: 6607 Question: The maximum cruise altitude can be limited by a 1.3g load factor because when exceeding that altitude. A. Use of normal manoeuvring bank angles may cause the limit load factor to be exceeded. B. Turbulence may induce high speed or low speed buffet. C. High speed buffet will occur immediately after exceeding this maximum altitude. D. Turbulence may cause the limit load factor to be exceeded.

16609 - D

6635 - B

6607 - B

Number: 16411 Question: The elevator deflection required for a given manoeuvre will be: A. the same at all speeds. B. larger at low IAS when compared to high IAS. C. the same for all CG positions. D. larger for an aft CG position when compared to a forward position.

Number: 6633 Question: When pulling out of a dive the angle of attack: A. Remains the same. B. Decreases. C. Increases. D. Cannot be increased at all due to structural considerations.

Number: 6571 Question: When an aircraft with a typical aerofoil is in level flight at low speed and high angle of attack, the normal axis is: A. Nearly vertical. B. Vertical. C. Horizontal from side to side. D. Horizontal from front to rear.

Number: 6632 Question: Which statement is correct? 1) The angle of attack of a positively cambered aerofoil has a negative value when the lift coefficient equals zero. 2) There is a nose down pitching moment about a positively cambered aerofoil when the lift coefficient equals zero. A. 1 is incorrect, 2 is correct. B. 1 is incorrect, 2 is incorrect. C. 1 is correct, 2 is correct. D. 1 is correct, 2 is incorrect.

16411 - B

6633 - C

6571 - A

6632 - C

Number: 6586 Question: When a trailing edge flap is lowered fully: A. the C of P moves to the rear and lift/drag ratio is increased. B. the C of P moves to the rear and lift/drag ratio is decreased. C. the C of P moves to the rear and lift/drag ratio is unaffected. D. the C of P moves forwards and lift/drag ratio is decreased.

Number: 6563 Question: Which of these statements about a stationary subsonic flow are correct or incorrect? I. The static pressure decreases as the streamlines converge. II. The velocity increases as the streamlines converge. A. I is incorrect, II is correct. B. I is correct, II is incorrect. C. I is correct, II is correct. D. I is incorrect, II is incorrect.

6586 - B

6563 - C

Number: 6622 Question: When the CG is close to the forward limit: A. Very high stick forces are required in pitch because the aircraft is very stable. B. Longitudinal stability is reduced. C. Very small forces are required on the control column to produce pitch. D. The stalling speed is reduced.

Number: 6640 Question: If an increase in power tends to make the nose of aeroplane rise, this is the result of the: A. Line of thrust being above the CG. B. Centre of lift and CB being collocated. C. Line of thrust being below the CG. D. Centre of lift being ahead of the CG.

Number: 6591 Question: Slats: A. De-energise the boundary layer thereby increasing the stalling angle of attack. B. De-energise the boundary layer, thereby decreasing the stalling angle of attack. C. Re-energise the boundary later thereby increasing the stalling angle of attack. D. Re-energise the boundary layer thereby decreasing the stalling angle of attack.

6622 - A

6640 - C

6591 - C

07-00 NEW QUESTIONS

PRINCIPLES OF FLIGHT

07-07 New Questions (xxx Questions)

Number: 10136 Question: An aeroplane performs a right turn, the slip indicator is left of neutral. One way to coordinate the turn is to apply: A B C D

a higher turn-rate. more left rudder. more right rudder. less right bank.

Number: 15108 Question: In a skidding turn (the nose pointing inwards), compared with a co-ordinated turn, the bank angle (i) and the "ball" or slip indicator (ii) are respectively: A B C D

(i) too large, (ii) displaced towards the high wing. (i) too large, (ii) displaced towards the low wing. (i) too small, (ii) displaced towards the high wing. (i) too small, (ii) displaced towards the low wing.

Number: 14968 Question: In a steady, horizontal, co-ordinated turn: A B C D

thrust equals drag, because drag is the same as in straight and level flight. thrust is greater than drag, because the excess thrust also supplies the centripetal force. thrust equals drag, because there is equilibrium of forces along the direction of flight. thrust is greater than drag, because the centrifugal force reduces the aeroplane speed.

10136 - B

15108 - C

14968 - C