POF • • • • •
•
LFnew/LForiginal = (Vnew/Voriginal)^2 Flap extension causes a reduction in stall speed and the maximum glide distance. The Principle of Continuity states: "If the cross sectional area of a streamlined flow of subsonic air is increased, the flow velocity will decrease". The Continuity equation is: "Cross-sectional area (A) x Velocity (V) = Constant". The illustration shows the relationships between CAS (C), TAS (T) and Mach number (M) with changing Pressure Altitude. It can be seen that when in a climb above the tropopause at a constant Mach number the TAS must remain constant.
•
VS0 is the stall speed in the landing configuration
•
VS1 is the stall speed in a specified configuration
•
VS1g is the minimum speed at which the aeroplane can develop a lift force (normal to the flight path) equal to its Weight
•
The reference stall speed is VSR
•
The Induced Drag coefficient formula is: CDi = (CL)2 / Aspect Ratio
•
Tip vortices are weaker when the aircraft is close to the ground (within half a wing span). The illustration shows that when an aeroplane is out of ground effect, the induced downwash (angle) is increased and consequently the induced angle of attack is increased.
•
•
The angle of attack of an aerofoil section is defined as the angle between the undisturbed airflow and the chord line. As flaps are deployed the wing camber increases and sometimes, in addition, the wing area. So if a constant angle of attack is maintained, the Lift coefficient would increase. The illustration shows a slat (a moveable part of the leading edge, which when activated forms a slot at the wing leading edge). The slot thus formed, directs high energy air onto the wing upper surface which increases the boundary layer kinetic energy on the top of the wing and delays the stall to a higher angle of attack It can be seen from the illustration that slats do not change the camber of the wing because the Lift curve is merely extended
•
Regarding subsonic airflow in a ventury:
•
1. the dynamic pressure in the undisturbed flow and in the throat are not equal
• •
2. The total pressure in the undisturbed flow and in the throat are equal 3. Static pressure acts in all directions Stall speed (IAS) varies with weight. The Lift formula is: L = 1/2rho V2 CL S, where rho is the air density, V is the True Airspeed (TAS), CL is the Lift coefficient and S is the wing area. Together, 1/2rho V2 = Dynamic Pressure (q). Although the question asks you to consider the Lift formula, it is actually necessary to consider how the Airspeed Indicator works. We know from the Dynamic Pressure formula (q = 1/2rho V2), if we double the speed of the aircraft through the air (double the TAS), Dynamic pressure (q) will be four times greater but because of the square root gearing inside the ASI, the Indicated Airspeed will only double. If you now put four times the Dynamic Pressure into the Lift formula, it is clear that the Lift will be four times greater. If you prefer the math’s explanation, IAS is proportional to the square root of Dynamic Pressure (q). Transposed this becomes Dynamic Pressure (q) is proportional to IAS squared. This is a theory question. In flight you need to keep Lift equal to Weight. So to maintain constant Lift as airspeed (IAS) is doubled, you would need to reduce the angle of attack to 1/4 of its previous value.
•
Induced Drag is caused by tip vortices. The stronger or more effective the tip vortices, the greater the Induced Drag. Tip vortices form because of the pressure differential between the top and bottom surface of the aerofoil. The greater the Weight, the greater the Lift required to balance it, therefore the top and bottom surface pressure differential will be greater, giving stronger tip vortices and more Induced Drag. Consequently, if wing Lift is zero there will be no tip vortices and therefore zero Induced Drag.
•
The relationship between pressure, density and absolute temperature of a given mass of air can be expressed as p / (rho * T) = constant. This question relates to the 'Ideal Gas Law', which states that air density (rho) is proportional to pressure (p) and inversely proportional to absolute temperature (T). Assuming no compressibility effects, induced drag at constant IAS is affected by Aeroplane Weight (mass) will affect the amount of Lift produced, so will therefore affect Induced Drag. Where the horizontal axis crosses the vertical axis is zero Lift coefficient (CL). The swept back wing has an increased tendency to stall first at the tips due to the span-wise flow from root to tip. This reduces Lift aft of the CG and generates an aircraft nose up pitching moment. The nose-up pitching moment of some aircraft with a swept back wing is so violent and fast that no human is capable of reacting fast enough to prevent it. This is known as a Deep Stall, which can also be called a Super Stall. Any aeroplane susceptible to Deep Stall can never be allowed to stall and must therefore be fitted with a "Stall prevention device", called a Stick Pusher. With
• • • • •
•
•
increasing angle of attack the Stall Warning system (Stick shaker) will activate as normal, but if the pilot does not decrease the angle of attack, the stick pusher then activates and pushes the stick (elevator control) forward to prevent Deep Stall from happening. In two dimensional flows, span wise flow is not considered; therefore there can be no Induced Drag. Interference Drag results from the interference of the boundary layers of adjacent components, but here we have no adjacent components and therefore we don't need to consider Interference Drag. The only correct answer is Pressure Drag and Skin Friction Drag. Flap asymmetry usually happens when high lift devices are selected - one side moves and the other side stay in its original position. The illustration shows that flap asymmetry will cause a large rolling moment at any angle of attack, whereas slat asymmetry would merely cause a large difference in CLMAX. One of the biggest advantage of a slat is they do not significantly increase Drag. The given 'correct' answer therefore makes no sense. "Flap asymmetry causes a large rolling moment", is absolutely correct, but not: "a yawing moment from slat asymmetry".
•
From the illustration it can be seen that L/D MAX corresponds to the speed VMD in straight and level (1g) flight. So if speed is reduced below L/D MAX the Total Drag will increase due to increasing Induced Drag.
•
The stall speed decreases: (all other relevant factors are constant) when, during a manoeuvre, the aeroplane nose is suddenly pushed firmly downwards (e.g. as in a push over). Load Factor (n) = Lift / Weight. If the aircraft is manoeuvred by suddenly pushing the nose down the Lift will be reduced and from the above formula it can be seen that less Lift will decrease the stall speed. Because tip vortices generate Induced Drag, anything that interferes with the formation or strength of tip vortices will reduce Induced Drag. Winglets reduce the strength of tip vortices and therefore reduce Induced Drag. However, anything placed in the airflow will also generate Parasite Drag.
•
•
High lift devices (leading and trailing edge) are fitted to an aeroplane to decrease the take-off and landing distance. Decreasing the take-off and landing distance is achieved by increasing the maximum lift coefficient (CLMAX) which will decrease the stall speed and therefore the minimum operating speed. When (trailing edge) flaps are deployed from 0° to 15°, there will be a comparatively large increase in CL and a small increase in CD, but with each successive increase in flap angle, the increase in CL will be less, whereas the increase in CD will be greater. For maximum efficiency during take-off, not only is an increase in CLMAX desired, but any increase in Drag needs to be minimized. Consequently, the 'optimum' flap for take-off is approximately 15°. However, for landing, maximum Drag is required, so full flap is necessary.
•
• •
Use the Lift formula for this question: (L = 1/2rho V2 CL S), where S is wing area. If the wing area is increased, Lift will increase because it is directly proportional to wing area. Sweeping the wing, either forward, or as is more usual, backwards, decreases the aerodynamic efficiency of the wing. Consequently, CLMAX is decreased. Therefore, if sweep angle is increased CLMAX will decrease. If we consider VS at CL MAX (VS = the square root of L / 1/2rho CL MAX S), it can be seen that a decrease in CLMAX will increase the stall speed.
Increasing forward sweep increases the stall speed. Wing anhedral is used by the designers to give an aeroplane the required level of Lateral Static Stability. Anhedral has no influence on stalling. • • •
•
•
The load factor is greater than 1 (one) when lift is greater than weight. Wing fences reduce the span-wise flow and help to reduce the increased tendency for a swept wing to tip stall and consequently pitch-up is reduced. Wing fences therefore improve the low speed handling characteristics of a swept wing. Use the Lift formula to consider this question: (L = 1/2rho V2 CL S), where 'rho' is air density. Altitude is the only variable in the question and any change in altitude will affect the air density. The lower the altitude the greater the air density, which will require a lower TAS to maintain a constant IAS and Lift equal Weight. The swept back wing has an increased tendency to stall first at the tips due to the span-wise flow from root to tip. This reduces Lift aft of the CG and generates an aircraft nose up pitching moment. The nose-up pitching moment of some aircraft with a swept back wing is so violent and fast that no human is capable of reacting fast enough to prevent it. This is known as a Deep Stall, which can also be called a Super Stall. Any aeroplane susceptible to Deep Stall can never be allowed to stall and must therefore be fitted with a "Stall prevention device", called a Stick Pusher. With increasing angle of attack the Stall Warning system (Stick shaker) will activate as normal, but if the pilot does not decrease the angle of attack, the stick pusher then activates and pushes the stick forward to prevent Deep Stall from happening. A wing with forward sweep does not have an increased tendency to tip stall and will not generate a nose-up pitching moment before the stall. Therefore an aeroplane with a swept forward wing is not susceptible to Deep Stall. A contributory factor to Deep Stall is a T-tail. At the extremely high angle of attack suffered in a Deep Stall the T-tail would be immersed in separated airflow, making the elevator ineffective and preventing the pilot from decreasing the angle of attack. A low mounted tail would not suffer from this problem and is therefore not a contributory factor to Deep Stall. From the illustration it can be seen that as the angle of attack decreases, the stagnation point moves up on the leading edge and the point of minimum pressure moves aft.
•
Tip vortices induce downwash behind the wing, at the tip. This induced downwash changes the average direction of the airflow over the wing. The angle of this modified airflow (The Effective Airflow) to the Relative Airflow is the Induced Angle of Attack. See illustration below.
•
Assuming constant IAS, when an aeroplane enters ground effect the effective angle of attack increases. CL = 1 / V² Air density is mass per unit volume; the unit for mass is the kg (kilogram) and the unit for volume is the m³ (cubic meter), hence density is kg/m³. The unit for force is the Newton (N).
• •
•
Wing loading is the aircraft weight (N) divided by wing area (m²), hence wing loading is N/m². Dynamic pressure is force per unit area so is also N/m².
• •
Lift is generated when a certain mass of air is accelerated in its flow direction. The dynamic pressure increases as static pressure decreases.
•
The turbulent boundary layer has more kinetic energy than the laminar boundary layer.
•
A positive camber aerofoil section at zero angle of attack will generate a small amount of lift. This is because the cross sectional area of the streamlined flow is accelerated more over the 'top' surface than the bottom. From the illustration it can be seen that the Lift curve of a positive cambered aerofoil intersects the vertical axis of the CL - alpha graph above the point of origin.
•
On a swept wing aeroplane at low airspeed, the "pitch up" phenomenon is caused by wingtip stall.
• • •
•
• • •
• • • • • • •
The angle between the aeroplane longitudinal axis and the chord line is the angle of incidence. From the illustration below, it can be seen that the aeroplane (Total) Drag in straight and level flight is lowest when the Parasite Drag is equal to the Induced Drag. Due to the formation of shock waves on the wing top and bottom surface and the subsequent rearwards movement of the shock waves as the aeroplane accelerates from subsonic to supersonic speed, the distribution of pressure on the wing changes, causing both the centre of pressure (CP) and the aerodynamic centre (AC) to move aft from 25% to the 50% chord position (mid chord). A symmetrical aerofoil section at zero degrees angle of attack will accelerate the air over the top surface and the bottom surface by the same amount, hence there will be no net lift either upwards or downwards, consequently the lift coefficient will be zero. An aeroplane maintains straight and level flight while the IAS is doubled. The change in lift coefficient will be x 0.25. The most important problem of ice accretion on a transport aeroplane during flight is reduction in CLmax. The effect of heavy (tropical) rain on the aeroplane is significant. The large amount of water on the aeroplane will increase the Weight and the accumulation of water will also disrupt the airflow. This will both decrease CL MAX and increase Drag. At the highest value of the lift/drag ratio the total drag is lowest. By definition, lift is perpendicular to the relative airflow (also known as relative wind or free stream flow) and drag is parallel to and in the same direction as the relative airflow (relative wind or free stream flow). The Parasite Drag formula is: Dp = 1/2rho V2 CDp S, therefore Dp is proportional to speed squared. 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: (i) 1/16 (ii) 1/4 Induced drag is created by the spanwise flow pattern resulting in the tip vortices. Vortex generators transfer energy from the free airflow into the boundary layer . Anhedral can be used by the designers to set the required Lateral Static Stability. Increasing Anhedral reduces Lateral Static Stability, but has no influence on stall speed. Increasing the sweep angle of the wing is used by the designers to increase the Critical Mach number. Increasing the sweep angle also makes the wing less aerodynamically efficient which will increase the stall speed. A T-tail is fitted by the designers to reduce the influence of wing downwash on the horizontal stabilizer. A T-tail also acts as an "End Plate" on the fin, which makes the fin more aerodynamically efficient. A T-tail has no influence on the stall speed.
•
An "erect" spin is with the top of the aeroplane towards the inside of the helix (corkscrew shape) of the spin. The "Standard" spin recovery includes: Full rudder opposite to the direction of rotation, close the throttle, neutralize the roll control and move the pitch control forward. So keeping the aileron control neutral during recovery is the only true statement here.
•
"Just before the stall a nose down pitching moment is generated". This is caused by the rearward movement of the CP. The vane of a stall warning system with a flapper switch is activated by the change of the stagnation point. In a steady level, co-ordinated turn the Lift must be greater than the Weight, so therefore the Load Factor (n) will be greater than 1. To consider the effect on the stall speed we can refer to the formula: VS NEW = VS OLD x the square root of the Load Factor. The key to this question is the part which states, "… whilst maintaining level flight at a constant IAS …" To maintain level flight at a constant IAS, Lift must remain the same, but when flaps are extended Lift increases. Consequently, in order to maintain level flight as the flaps go down, the angle of attack must be decreased to keep Lift constant. Therefore when flaps are extended whilst maintaining straight and level flight at constant IAS the Lift coefficient (CL) must remain the same.
• •
•
• •
Trailing edge flap extension will decrease the critical angle of attack and increase the value of CLmax. It can be seen from the illustration that deploying a slat will form a slot and deploying a Krueger flap does not form a slot. It is also evident that deploying both a Krueger flap and a slat will increase the critical angle of attack.
•
It can be seen from the illustration that of the types of trailing edge flap listed, the Fowler flap is the most effective (highest CLMAX).
•
A slotted flap will increase the CLmax by increasing the camber of the aerofoil and re-energizing the airflow.
•
The purpose of an auto-slat system is to extend automatically when a certain value of angle of attack is exceeded. The type of stall that has the largest associated angle of attack is a deep stall. As an aeroplane accelerates through the transonic region the shock waves on the wing move rearwards and change the pressure distribution on the wing, which makes the Aerodynamic Centre (AC) move rearwards from 25% aft towards 50% aft. The top illustration shows the Longitudinal Static Stability in subsonic flow, with the AC at 25% aft. The bottom illustration shows the AC at 50% aft when the flow becomes supersonic.
• •
It can be seen that in supersonic flow, when the AC has moved aft to the 50% chord position, the destabilizing wing moment is reduced and consequently, Static Longitudinal Stability is increased.
• • • • • •
The shockwaves associated with supersonic flight cause the pressure patterns to appear rectangular in form unlike the pressure patterns for subsonic flight which appear smooth and slightly humped when plotted on a graph. Increasing air density will have the following effect on the drag of a body in an air stream (angle of attack and TAS are constant) the drag increases. "A line connecting the leading- and trailing edge midway between the upper and lower surface of a aerofoil". This definition is applicable for the camber line. Power is the rate of doing work (how quickly work is done). Work is Force (N) x Distance (metre), so Power is Force (N) x Distance (metre) divided by time (seconds) or N. m/s Compared with the clean configuration, the angle of attack at CLmax with trailing edge flaps extended is smaller. The illustration shows an Expansion Wave. One of the characteristics of supersonic flow is that it can follow a convex corner because it expands upon reaching the corner. Consequently, velocity increases and the other parameters, pressure, density and temperature all decrease. Local speed of sound (a) is proportional to air temperature. So as the temperature decreases through an expansion wave, the local speed of sound will decrease.
• • • • • • •
• • •
The flow on the upper surface of the wing has a component in wing root direction. Spoiler extension increases the stall speed, the minimum rate of descent and the minimum angle of descent. During a climbing turn to the right the angle of attack of both wings is the same. Pstat + ½rhoV²= constant. 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: 60°. During a steady horizontal turn, the stall speed increases with the square root of the load factor. The illustration shows the correct orders of increasing critical angle of attack are: flaps only extended, clean wing, slats only extended.
Floating due to ground effect during an approach to land will occur when the height is less than halve of the length of the wing span above the surface. An aerofoil is cambered when the line, which connects the centres of all inscribed circles, is curved. From the illustration it can be seen that while flying in the speed unstable region, any reduction in speed will give an increase in Drag, which will lead to a greater reduction in speed.
• •
• •
• • • • •
An aeroplane in straight and level flight is subjected to a strong vertical gust. The point on the wing, where the instantaneous variation in wing lift effectively acts is known as the: aerodynamic centre of the wing. The Boundary Layer is a layer of air flowing over a surface, the lowest layer of molecules having zero velocity relative the surface, but increasing in speed with increasing distance from the surface, until the free stream flow velocity is reached. Therefore the Boundary Layer has a lower average velocity than that of the free stream. The polar curve of an aerofoil section is a graphic relationship between lift coefficient Cl and drag coefficient Cd. The Boundary Layer is a layer of air flowing over a surface, the lowest layer of molecules having zero velocity relative the surface, but increasing in speed with increasing distance from the surface until the free stream flow velocity is reached. There are two types of boundary layer, the laminar boundary layer and the turbulent boundary layer. At the leading edge the boundary layer is initially laminar (no intermixing of adjacent layers) and will remain laminar as it moves rearwards until it becomes turbulent at the Transition Point. The turbulent flow will continue towards the trailing edge and just before it reaches the trailing edge it will start to separate from the surface, this is called The Separation Point. Because the laminar boundary layer is smooth with no intermixing its kinetic energy is comparatively low and is therefore easily separated. Because the turbulent boundary layer has lots of intermixing between the layers of air molecules it will therefore contains more kinetic energy than a laminar boundary layer. Consequently, compared to a laminar boundary layer, a turbulent boundary layer is better able to resist a positive pressure gradient before it separates. A positively cambered aerofoil will generate zero lift: at a negative angle of attack. Wing Loading = Weight / Wing Area. Therefore for a given Wing Area, the greater the Weight, the higher the Wing Loading. Stall speeds are determined with the CG at the forward limit. Minimum control speeds are determined with the CG at the aft limit. Regulations state that the stall warning must begin at a speed (VSW) exceeding the speed at which the stall is identified by not less than 5 kt or 5% CAS, whichever is greater. That would be 1.05 VS, so the only correct answer is "greater than VS".
• • •
• • • • •
•
The Mean Aerodynamic Chord (MAC) for a given wing of any plan form is the chord of a rectangular wing with same moment and lift. Behind the transition point in a boundary layer the mean speed and friction drag increases. The formula we need when calculating Lift (L) or Load Factor (n) in a turn is: L or n = 1 / COS PHI; where PHI = bank angle. And when calculating the affect of bank angle on stall speed, the formula becomes VS NEW = VS OLD x the square root of 1 / COS PHI or to put it another way, VS NEW = VS OLD x the square root of n. Where PHI is the bank angle and n = Load Factor. Therefore when a pilot makes a turn in horizontal flight, the stall speed increases with the square root of the Load Factor. The true airspeed (TAS) is lower than the indicated airspeed (IAS) at ISA conditions and altitudes below sea level. The lift force, acting on an aerofoil: (no flow separation) is mainly caused by suction on the upper side of the aerofoil. The relative thickness of an aerofoil is expressed in % chord. Aspect ratio of a wing is the ratio between wing span squared and wing area. The wing of an aeroplane will never stall at low subsonic speeds as long as.... the angle of attack is smaller than the value at which the stall occurs.
The illustration shows an Expansion Wave. One of the characteristics of supersonic flow is that it can follow a convex corner because it expands upon reaching the corner. Consequently, velocity increases and the other parameters, pressure, density and temperature all decrease. Consequently, the density in front of an expansion wave is higher than behind it and the pressure in front of an expansion wave is also higher than behind it.
•
Static stability is the initial (split second) reaction of the aeroplane immediately following the removal of a disturbing force. Dynamic stability is what happens after the initial reaction. If an aeroplane is statically stable it will start back towards its previous state of equilibrium and if an aeroplane is statically unstable it will start to move further away from its previous state of equilibrium.
Therefore, Dynamic stability is possible only when the aeroplane is statically stable. •
It must be fully understood that this is a theory question. In practice you would not even taxi with ice on the wing leading edge, let alone attempt a take-off. About 80% of Lift is generated by the top surface of the wing. And of the top wing surface, about 20% aft of the leading edge is where the majority of the flow acceleration takes place. So anything which interferes with the airflow over the leading 20% of the wing top surface will have a major influence on Lift generation. The question gives us ice located on the wing leading edge. During the take-off run the aeroplane is merely accelerating down the runway and ice located on the wing leading edge will only increase the Drag slightly. During the climb with all engines operating, the aeroplane will be at a small angle of attack (about 4°) and ice on the wing leading edge will increase the Drag and decrease CL by a moderate amount. During the last part of rotation, the angle of attack will be quite high and any ice on the leading edge will have a significant negative influence on Lift production and in fact the wing will probably fully stall.
•
The following factors increase stall speed: an increase in load factor, a forward cg shift, and decrease in thrust.
•
The illustration shows a (Normal) shock wave. As the airflow passes through a shock wave the velocity decreases and the pressure, the density and the temperature all increase.
•
At what speed does the front of a shock wave move across the earth's surface? The ground speed of the aeroplane. Low speed pitch up is caused by the outward drift of the boundary layer on a swept-back wing. The Centre of Pressure (CP) is the point on an aerofoil through which the Lift acts. We know that the location of the CP is a product of the average of the top surface Lift pressure and the bottom surface Lift pressure. We are also aware that as the angle of attack is increased, the Lift pressure distribution changes and the CP moves forward. When studying Longitudinal Static Stability and certain aspects of Aircraft Limitations, use of the moving CP would be confusing, so in
• •
those circumstances a different reference point is used, called the Aerodynamic Centre (AC). Consider illustration (A) below: The Aerodynamic Centre (AC) is located 25% aft [of the leading edge]. It can be seen that at angle of attack (alpha 1), Lift (L1) acts through distance (d1) from the Aerodynamic Centre (AC). This generates a nose down pitching moment (M) about the AC. If you now consider illustration (B): it can be seen that an upward vertical gust has increased the angle of attack to (alpha 2), which increases the Lift to (L2) and moved the CP forward, decreasing the distance of the CP from the AC to (d2). But because (L2) has increased by the same amount that (d2) has decreased, the nose down pitching moment (M) about the AC has remained the same. Now consider illustration (C): because pitching moment (M) has stayed the same with an increase in angle of attack, the change in Lift (delta L), can be considered to act at the AC, as illustrated in illustration (D). Using the Aerodynamic Centre (AC) as the point through which the change in Lift acts greatly simplifies the visualization of the effect of changes in angle of attack due to gusts, both in the study of Longitudinal Static Stability and the affect of a gust stressing the aeroplane when studying Aircraft Limitations. Facts to remember about the Aerodynamic Centre: 1, The AC is located 25% aft [of the leading edge] in subsonic flow. 2. The pitching moment about the AC is always nose down. 3. The pitching moment about the AC does not change if the angle of attack changes. 4. The AC is the point through which the change in Lift acts [due to a change in angle of attack].
•
Wing sweep angle is the angle between the quarter-chord line of the wing and the lateral axis.
• • •
The mean geometric chord of a wing is the wing area divided by the wing span. Taper ratio of a wing is the ratio between tip chord and root chord. Wing twist (geometric and aerodynamic) is used to: 1. improve stall characteristics. 2. Reduce induced drag. Bernoulli's equation is Pstat + 1/2 * rho *TAS² = constant. If ice is present on the leading edge of the wings, it may increase the landing distance due to a higher Vth with 40-50%. The load factor is less than 1 (one) during a steady wings level descent. The load factor is less than 1 (one) When lift is less than weight. The load factor is less than 1 (one) during a steady wings level climb. The top shock will be at its weakest at Mcrit, it is only just forming The shockwave moves aft as M increases 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 As altitude increased, the stall speed (IAS) initially remains constant and at higher altitudes increases. Shock induced separation can occur behind a strong normal shock wave, independent of angle of attack. The subsonic speed range ends at Mcrit.
• • • • • • • • • • • •
•
There are three aeroplane speed ranges. The slowest speed range is 'Subsonic'; defined as the speed range when all of the flow speeds relative to the surface of the aeroplane are less than the speed of sound (ML < 1.0). The 'subsonic' speed region is between zero and the Critical Mach number (MCRIT). The next highest speed range is 'Transonic'; defined as the speed range where some of the flow speeds relative to the surface of the aeroplane are less than the speed of sound (ML < 1.0) and some of the flow speeds relative to the surface of the aeroplane are greater than the speed of sound (ML > 1.0). The 'Transonic' speed region is between the Critical Mach number (MCRIT) and approximately M1.2 The highest speed region in which we are interested is 'Supersonic'; defined as the speed range where all of the flow speeds relative to the surface of the aeroplane are greater than the speed of sound (ML > 1.0). The 'Supersonic' speed region is between approximately M1.2 and M5.0
•
The critical Mach number of an aeroplane is the Mach number above which, locally, supersonic flow exists somewhere over the aeroplane. The critical Mach number of an aeroplane can be increased by sweepback of the wings. Wave Drag is mostly caused by the airflow being heated as it passes through the shock waves, but a small proportion of Wave Drag is due to shock wave induced separation.
• •
•
Swept back wings are fitted to an aeroplane to increase its critical Mach number (MCRIT). Therefore, decreasing wing sweepback will decrease MCRIT. The Drag Divergence Mach Number is the Mach number at which the aerodynamic Drag on an aerofoil or an airframe begins to increase rapidly as the Mach number continues to increase. This increase in Drag is a result of shock wave formation (Wave Drag), and can cause the Drag coefficient to rise to more than ten times its low speed value. The Drag Divergence Mach Number is usually close to, and always greater than, the Critical Mach number. Therefore if the Critical Mach number decreases, the Drag Divergence Mach Number will decrease.
•
An oblique shock wave has similar affects on the airflow passing through it as does a normal shock wave, but there are a few small variations. 1. An oblique shock wave is so called because it is at an angle of greater than 90° to the direction of the upstream flow. 2. There IS a flow direction change through an oblique shock wave. 3. The airflow enters an oblique shock wave at supersonic speed and emerges at a lower, but still supersonic speed. 4. There is an increase in Static Pressure, Density and Compression through an oblique shock wave.
5. There is a decrease in the energy of the airflow and a decrease in Total Pressure through an oblique shock wave. 6. The temperature of the air increases through an oblique shock wave. For this question we need to remember: because the temperature increases through an oblique shock wave the temperature behind is higher than in front of it. And because the Static Pressure increases through an oblique shock wave, the Static Pressure behind is higher than in front of it. •
•
• • •
•
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 stabilizer. The illustration shows that, in comparison to a conventional aerofoil section, typical shape characteristics of a supercritical aerofoil section are: a larger nose radius, flatter upper surface and negative, as well as positive, camber.
During a descent at a constant Mach number (assume zero thrust and standard atmospheric conditions) the angle of attack will decrease. If a symmetrical aerofoil is accelerated from subsonic to supersonic speed, the aerodynamic centre will move aft to the mid chord. Reducing the thickness/chord ratio of a wing increases MCRIT in the same way as does a swept wing. Therefore there will be a reduction in the variations in Drag coefficient, a delay in the onset of shock wave formation and a reduction in the variations in Lift coefficient. A normal shock wave has several distinguishing features, chief among them being: 1. A Normal shock wave is so called because it is normal (perpendicular) to the direction of the upstream flow. 2. There is no flow direction change through a normal shock wave. 3. The airflow enters a normal shock wave at supersonic speed and emerges at subsonic speed. 4. There is a greater increase in Static Pressure, Density and compression through a normal shock wave, compared to an oblique shock wave. 5. There is a greater decrease in the energy of the airflow and lower Total Pressure through a normal shock wave, compared to an oblique shock wave. 6. The temperature of the air increases through a normal shock wave.
• • • • • • • • • • •
7. The least energy loss through a normal shock wave is when the Local Mach number (ML) is just above ML1.0 Compressibility effects depend on Mach number. The Mach number is the ratio between the TAS of the aeroplane and speed of sound of the undisturbed flow. A "Shock Stall" is due to airflow separation from the formation of a shock wave. Shock waves form at high speed and therefore a small angle of attack. When an aeroplane is flying faster than M1.0, any pressure changes taking place in the air flow over the surface will only affect parts of the aeroplane within the Mach Cone. (MCRIT) is: "The aircraft Mach number at which the local velocity first reaches M1.0". The average modern high speed jet transport has an MCRIT of about M0.8 The loss of total pressure in a shock wave is due to the fact that kinetic energy in the flow is converted into heat energy. The maximum acceptable cruising altitude is limited by a minimum acceptable load factor because exceeding that altitude Turbulence may induce mach buffet. The speed of sound varies with the square root of the absolute temperature. The lower the temperature, the lower the speed of sound; and vice versa. The 'sonic boom' (or sonic bang) of an aeroplane flying at supersonic speed is a result of the shock waves the aeroplane is generating striking the ear drums of a listener (usually on the ground). (M = TAS / a). Vortex generators mounted on the upper wing surface will decrease the shock wave induced separation.
•
As an aeroplane accelerates above its critical Mach number (MCRIT) a shock wave will form on the wing top surface. As the aeroplane continues to accelerate, this shock wave will get thicker, increase in length and move backwards. At about M0.9 another shock wave will form on the wing bottom surface, getting thicker, longer and moving backwards with further increases in Mach number. At about M0.98 these two shock waves will reach the trailing edge of the wing. As soon as M1.0 is exceeded, another shock wave will form a small distance in front of the wing. This is called the Bow Wave because it is similar in appearance to the bow wave in front of a boat as it moves through water. With further acceleration, the Bow Wave will move closer to the wing leading edge and eventually will get no closer at about M1.3 to M1.4. This is called: the speed of attachment.
•
The sin of the Mach (cone) angle (µ) = a / TAS. In other words, sin µ = 1 / M. Therefore, the greater the Mach number (M), the smaller the Mach angle (µ) The Mach trim system will adjust the stabilizer, depending on the Mach number. When air has passed through a shock wave the local speed of sound is increased.
• •
• • • • • • • • • • • • • • • • • • • • •
The speed range between high and low speed buffet increases during a descent at a constant IAS. High speed buffet is induced by boundary layer separation due to shock waves. Vortex generators on the upper side of the wing decrease wave drag. Mach buffet occurs at the Mach number at which shock wave induced boundary layer separation occurs. A transonic Mach number is a Mach number at which both subsonic and supersonic local speeds occur. Speed of sound increases with temperature increase. In transonic flight the ailerons will be less effective than in subsonic flight because aileron deflection only partly affects the pressure distribution around the wing. As an aircraft accelerates through the transonic speed range the coefficient of drag increases then decreases. Shock stall occurs when the lift coefficient, as a function of Mach number, reaches its maximum value. When an aeroplane is flying faster than M1.0, any pressure changes taking place in the air flow over the surface will only affect parts of the aeroplane within the Mach Cone. On a typical transonic airfoil the transonic rearward shift of the CP occurs at about M 0.89 to M 0.98. No noticeable shock waves will form over any wing section when flying just above MCRIT. "Coffin Corner" is the colloquial name for the pressure altitude where the speed for low speed buffet is the same as the speed for high speed buffet. The 'Technical' name for which is "The Aerodynamic Ceiling". Dutch Roll is stability related characteristic and occurs when the Lateral Static Stability is greater than the Directional Static Stability. Shock waves have no influence on Dutch Roll. The local speed of sound is dependent on temperature only. The lower the temperature, the lower the local speed of sound; and vice versa. Stall speed does not vary with Density Altitude. However, at very high altitudes the stall speed will start to increase due to increasing Mach number (compressibility effects). The Mach-trim function is installed on most commercial jets in order to minimize the adverse effects of changes in the position of centre of pressure. What happens to lateral stability when flaps are extended? Lateral stability is decreased. A downward adjustment of a trim tab in the longitudinal control system has the following effect: the stick position stability remains constant. During a phugoid altitude varies significantly, but during a short period oscillation it remains approximately constant. An increase of 10kt from the trimmed position at high speed has less effect on the stick force than an increase of 10kt from the trimmed position at low speed.
• • • •
• • • • • •
Static lateral stability will be decreased by: the use of a low, rather than high, wing mounting. Dutch roll will be corrected by a yaw damper. Static lateral stability will be decreased by: reducing wing sweepback. This information provided is of a Cmcg diagram, showing moment about the CG against alpha In Part 1 the aircraft wants to pitch up, but as alpha increases the pitch up force gets less and less i.e. it is self correcting and thus positively stable. Where the trace crosses the horizontal line Cmcg is zero, the aircraft neither wants to go nose up nor nose down. In fact it is "in trim" in both places. At Point 2 just briefly, changing alpha has no effect on the nose up or down moment so the aircraft is neutrally longitudinally stable. In Part 3, unfortunately for the pilot, any increase in alpha will increase the nose up moment. The aircraft is negatively stable - unstable. In fact this is typical of the entry into an irrecoverable deep stall. The slope of the line shows stability or not, above or below the horizontal axis shows where the nose will want to go if you let go of the stick
Static lateral stability will be increased by: increasing wing sweepback. What is predominantly used to set the Tail Horizontal Stabilizer for Take-off? CG. Static directional stability is the: tendency of an aeroplane to recover from a skid with the rudder free. Forward movement of the CG will reduce control response and increase stability. How can the designer of an aeroplane with straight wings increase the static lateral stability? By increasing the aspect ratio of the vertical stabilizer, whilst maintaining a constant area. The pitching moment versus angle of attack line in the diagram, which corresponds to a CG located at the neutral point of a given aeroplane at low and moderate angles of attack is: line 2. The Neutral Point is the position of the CG that gives the aircraft neutral longitudinal static stability.
• •
• • • • • •
• • • • • • •
The effect of a high wing with zero dihedral is Positive dihedral effect. 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? 450 N. What is the recommended action following failure of the yaw damper(s) of a jet aeroplane, flying at normal cruise altitude and speed prior to encountering Dutch roll problems? Reduce altitude and Mach number. As the stability of an aeroplane increases its manoeuvrability decreases. Excessive static lateral stability is an undesirable characteristic for a transport aeroplane because it would impose excessive demands on roll control during a sideslip. An aeroplane that tends to return to its pre-disturbed equilibrium position after the disturbance has been removed is said to have positive static stability. Static lateral stability will be increased by reducing wing anhedral. The contribution of the wing to the static longitudinal stability of an aeroplane depends on CG location relative to the wing aerodynamic centre. An aft CG shift decreases static longitudinal stability. Static lateral stability will be decreased by increasing wing anhedral. For an aeroplane to possess dynamic stability, it needs static stability and sufficient damping. An aeroplane is sensitive to Dutch roll when static lateral stability is much more pronounced than static directional stability. The purpose of a dorsal fin is to maintain static directional stability at large sideslip angles. The effect of a positive wing sweep on static directional stability is as follows stabilizing effect. If the total sum of moments about one of its axes is not zero, an aeroplane would experience an angular acceleration about that axis.
• • • • • • • • • • • • • • • • • • • • • •
An aeroplane has static directional stability; in a side slip to the right, initially the nose of the aeroplane tends to move to the right. For a normal stable aeroplane, the centre of gravity is located with a sufficient minimum margin ahead of the neutral point of the aeroplane. The maximum aft position of the centre of gravity is, amongst others, limited by the required minimum value of the stick force per g. The manoeuvrability of an aeroplane is best when the cg is on the aft cg limit. An aeroplane with an excessive static directional stability in relation to its static lateral stability will be prone to spiral dive (spiral instability). A Mach trimmer corrects the change in stick force stability of a swept wing aeroplane above a certain Mach number. Positive static lateral stability is the tendency of an aeroplane to roll to the left in the case of a sideslip (with the aeroplane nose pointing to the left of the incoming flow). If the static lateral stability of an aeroplane is increased, whilst its static directional stability remains constant its sensitivity to Dutch roll increases. Lateral static stability is determined by Aircraft response to sideslip. The system is concerned with positive 'g' but if the pilot pushed the yoke forward and experienced negative 'g' it would increase the stick force the pilot felt in pushing the yoke forward. If an aeroplane exhibits insufficient stick force per g, this problem can be resolved by installing a bobweight in the control system which pulls the stick forwards. In a steady sideslip you are holding on aileron to balance the tendency of the aircraft to roll back level - sideslip stability. Increasing dihedral increases sideslip stability and you will then have to hold on more aileron to hold your attitude. Longitudinal stability is directly influenced by centre of gravity position. If an airplane has poor longitudinal stability in flight, what can be done to increase the stability? Increase stabilizer surface area. The aft CG limit can be determined by the minimum acceptable static longitudinal stability. A forward CG shift decreases longitudinal manoeuvrability. The dihedral construction of an aircraft wing provides Lateral stability about the longitudinal axis. Positive static longitudinal stability means that a nose-down moment occurs after encountering an up-gust. After an aeroplane has been trimmed the stick position stability will be unchanged. Stick force per g is dependent on cg location. An aeroplane's sideslip angle is defined as the angle between the speed vector and the plane of symmetry. The effect of the wing downwash on the static longitudinal stability of an aeroplane is negative.
• • • • • • • • • • • • • • • • • • • • •
The stick force per g must have both an upper and lower limit in order to ensure acceptable control characteristics. Upward deflection of a trim tab in the longitudinal control results in the stick position stability remaining constant. A bob weight and a down spring have the same effect on the stick force stability. What is the effect of elevator trim tab adjustment on the static longitudinal stability of an aeroplane? No effect. Which part of an aeroplane provides the greatest positive contribution to static longitudinal stability? The horizontal tailplane. Static lateral stability should not be too large, because too much aileron deflection would be required in a crosswind landing. When the CG is close to the forward limit: Very high stick forces are required in pitch because the aircraft is very stable. The effect of Mach trim on stick forces for power operated controls: Is to maintain the required stick force gradient. If the aircraft is properly loaded the CG, the neutral point and the manoeuvre point will be in the order given, forward to aft: CG, neutral point, manoeuvre point. A negative contribution to the static longitudinal stability of conventional jet transport aeroplanes is provided by: the fuselage. Any of the design features that increases the static lateral stability, by increasing the lift on the low wing, will add to the static directional stability. A straight wing with no dihedral will not have much of a change in lift on the low wing during a sideslip, so it will not have the increase in drag to yaw it into the relative airflow. Control surface flutter can be eliminated by: mass balancing of the control surface. The positive manoeuvring limit load factor for a light aeroplane in the utility category in the clean configuration is: 4.4. The relationship between the stall speed VS and VA (EAS) for a large transport aeroplane can be expressed in the following formula: VA >= VS * SQRT(2.5). For most jet transport aeroplanes, the maximum operating limit speed, VMO: is replaced by MMO at higher altitudes. The stall speed line in the manoeuvring load diagram runs through a point where the: speed = VS, load factor = +1. What can happen to the aeroplane structure flying at a speed just exceeding VA? It may suffer permanent deformation if the elevator is fully deflected upwards. Aircraft designers could build for a higher safe g limit if they wanted to, but cost and weight considerations usually mean that 2.5g is both the minimum and the maximum. Load factor is increased by: upward gusts. The extreme right limitation for both gust and manoeuvre diagrams is created by the speed: VD.
• •
• • • • • • • • • • • • • • • • • • •
The gust load factor due to a vertical upgust increases when: the gradient of the CL-alpha graph increases. Gust Load Factors vary depending on altitude, mass/weight, speed and the slope of the CL-alpha curve. With increasing altitude and increasing mass/weight the gust load factor will decrease. Wing loading is the ratio of all-up weight/wing area, so like mass/weight this will also result in a reduced gust load factor. Conversely if speed is increased then the effect of an upgust will be more severe and lead to an increased gust load factor. Which factor should be taken into account when determining VA? The limit load factor. Aileron flutter can be caused by: cyclic deformations generated by aerodynamic, inertial and elastic loads on the wing. 1. Increasing the aspect-ratio of the wing will increase the gust load factor.2. Increasing the speed will increase the gust load factor. I. Aero-elastic coupling affects flutter characteristics. II. The risk of flutter increases as IAS increases. Aileron reversal can be caused by: Twisting of the wing above reversal speed. All gust lines in the gust load diagram originate from a point where the speed = 0, load factor = +1 If climbing at VMO, it is possible to exceed MMO. VMO: should be not greater than VC. Which of these statements concerning flight in turbulence is correct? VRA is the recommended turbulence penetration air speed. The manoeuvring speed VA, expressed as indicated airspeed, of a transport aeroplane: depends on aeroplane mass and pressure altitude. The gust limit load factor can be higher than the manoeuvring limit load factor. Assuming ISA conditions, which statement with respect to the climb is correct? At constant IAS the Mach number increases. The positive manoeuvring limit load factor for a large transport aeroplane with flaps extended is: 2.0. The stall speed lines in the manoeuvring load diagram originate from a point where the: speed = 0, load factor = 0. How can wing flutter be prevented? By locating mass in front of the torsion axis of the wing. VA is: the maximum speed at which maximum elevator deflection up is allowed. The load factor in turbulence may fluctuate above and below 1 and can even become negative. Mass-balancing of control surfaces is used to: prevent flutter of control surfaces. For an aeroplane with one fixed value of VA the following applies. VA is: the speed at which the aeroplane stalls at the manoeuvring limit load factor at MTOW.
• • • • • • • • • • • • • • • • • • • • •
The significance of VA for jet transport aeroplanes is reduced at high cruising altitudes because: buffet onset limitations normally become limiting. Now on entering turbulence, a gust will change the angle of attack; the change in CL for the swept wing will be less than for the straight wing. Swept wings are less sensitive to gusts. Control surface flutter: Is a destructive vibration that must be damped out within the flight envelope. Flight in severe turbulence may lead to a stall and/or structural limitations being exceeded. What is the primary input for an artificial feel system? IAS. The elevator deflection required for a given manoeuvre will be: larger at low IAS when compared to high IAS, larger for a forward CG position when compared to an aft position. Yaw is followed by roll because the: 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. Outboard ailerons (if present) are normally used: in low speed flights only. The forward CG limit is mainly determined by the amount of pitch control available from the elevator. The centre of gravity moving aft will: increase the elevator up effectiveness. When the cg position is moved forward, the elevator deflection for a manoeuvre with a given load factor greater than 1 will be: larger. Stick forces, provided by an elevator feel system, depend on: elevator deflection, dynamic pressure. Which kind of ''tab'' is commonly used in case of manual reversion of fully powered flight controls? Servo tab. Which statement is correct about a spring tab? At high IAS it behaves like a servo tab. 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? (1) Reduces (2) reduces. Which three aerodynamic means decrease manoeuvring stick forces? Servo tab horn balance - spring tab. An example of differential aileron deflection during initiation of left turn is: Left aileron: 5° up. Right aileron: 2° down. An aeroplane is provided with spoilers and both inboard and outboard ailerons. Roll control during cruise is provided by: inboard ailerons and roll spoilers. In straight and level flight, as speed is reduced: the elevator is deflected further upwards and the trim tab further downwards. When roll spoilers are extended, the part of the wing on which they are mounted: experiences a reduction in lift, which generates the desired rolling moment. In addition there is a local increase in drag, which suppresses adverse yaw. When power assisted controls are used for pitch control: a part of the aerodynamic forces is still felt on the column.
• • • • • • • • • • •
• • • • • • • •
The pitch angle is defined as the angle between the: longitudinal axis and the horizontal plane. Aileron deflection causes a rotation around the longitudinal axis by: changing the wing camber and the two wings therefore produce different lift values resulting in a moment about the longitudinal axis. Aeroplane manoeuvrability decreases for a given control surface deflection when: IAS decreases. For a given elevator deflection, aeroplane longitudinal manoeuvrability increases when: the CG moves aft. Differential aileron deflection: equals the drag of the right and left aileron. Examples of aerodynamic balancing of control surfaces are: servo tab, spring tab, seal between the wing trailing edge and the leading edge of control surface. An advantage of locating the engines at the rear of the fuselage, in comparison to a location beneath the wing, is : less influence of thrust changes on longitudinal control. An aeroplane has a servo tab controlled elevator. What will happen if the elevator jams during flight? Pitch control sense is reversed. A horn balance in a control system has the following purpose: to decrease stick forces. What is the position of the elevator in relation to the trimmable horizontal stabiliser of a power assisted aeroplane that is in trim? The position depends on speed, the position of slats and flaps and the position of the centre of gravity. 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: rotation will require a higher than normal stick force. 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: in front of the hinge. In a differential aileron control system the control surfaces have a larger upward than downward maximum deflection. How does positive camber of an aerofoil affect static longitudinal stability? It has no effect, because camber of the aerofoil produces a constant pitch down moment coefficient, independent of angle of attack. An aeroplane's bank angle is defined as the angle between its: lateral axis and the horizontal plane. Low speed pitch-up can be caused by a significant thrust: increase with podded engines located beneath a low-mounted wing. Artificial feel is required: with fully powered flight controls. One advantage of mounting the horizontal tailplane on top of the vertical fin is: to improve the aerodynamic efficiency of the vertical fin. A primary stop is mounted on an elevator control system in order to: Restrict the range of movement of the elevator.
• • • • • • • • • • • • • • • •
•
When ice is present on the stabilizer, deflection of flaps may cause: The stabilizer to stall and a vertical dive. The following is true concerning a balance tab. It is: a form of aerodynamic balance. The inputs to the Q feel unit are from: Pitot and static pressures. The reasons for having a trim system on powered assisted flying controls is: Enables the stick force to be reduced to zero. Deflecting the elevator up, when the trim tab is in neutral, will cause the tab to: Remain in line with the elevator. Some airplanes have spring tabs mounted into the control system. This is to provide: A reduction in the pilots’ effort to move the controls against high air loads. The 'slipstream effect' of a propeller is most prominent at: low airspeeds with high power setting. Fixed-pitch propellers are usually designed for maximum efficiency at: cruising speed 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: decrease and the rate of descent will decrease. 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: increase and the rate of descent will decrease. A windmilling propeller: produces drag instead of thrust. The difference between a propeller's blade angle and its angle of attack is called: the helix angle. The reference section of a propeller blade with radius R is usually taken at a distance from the propeller axis equal to: 0.75 R. For a fixed-pitch propeller designed for cruise, the angle of attack of each blade, measured at the reference section: is optimum when the aircraft is in a stabilized cruising flight. Constant-speed propellers provide a better performance than fixed-pitch propellers because they: produce an almost maximum efficiency over a wider speed range. Propeller efficiency may be defined as the ratio between: usable (power available) power of the propeller and shaft power.
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 efficiency. The designer has to decide what he wants, you cannot have both. The angle of attack of a fixed pitch propeller blade increases when: RPM increases and forward velocity decreases Increasing speed reduces the angle of attack on the prop blades. If you had a fixed pitch prop this would mean that you lost thrust, and the faster you went the less thrust you would have. But you have a constant speed prop, and as the angle of attack reduces the torque drag reduces and the prop RPM begins to go up. The CSU senses this, and increases the blade angle to get the RPM stabilized again. So you settle down at the higher speed with increased blade angle.
• • • • • • •
Think Hurricane Mk I, fixed pitch wooden prop with a horrendously coarse pitch for high speed. For any propeller: thrust is the component of the total aerodynamic force on the propeller parallel to the rotational axis. If S is the frontal area of the propeller disc, propeller solidity is the ratio of: the total frontal area of all the blades to S. Increasing the camber on propeller blades will, if all else is the same: Increase the power absorption capability. The number of blades in a propeller would be increased: To increase power absorption capability. Counter rotating propellers have the effect of: Canceling out the torque and gyroscopic effects. The first action in the event of propeller runaway (overspeed conditions), should be to: Close the throttle. How will the area ratio of a propeller be calculated? Area of all propeller blades to the total circular surface.
• • • • • • • • • • • • • • • • • • • • • •
During a straight, steady climb and with the thrust force parallel to the flight path: lift is the same as during a descent at the same angle and mass. For shallow flight path angles in straight and steady flight, the following formula can be used: sin gamma = T/W - CD/CL. Turning motion in a steady, level co-ordinated turn is created by: the centripetal force. In a co-ordinated horizontal turn, the magnitude of the centripetal force at 45 degrees of bank: is equal to the weight of the aeroplane. What factors determine the distance traveled over the ground of an aeroplane in a glide? The wind and the lift/drag ratio, which changes with angle of attack. The speed VMCL can be limited by the available maximum roll rate. For a given aeroplane which two main variables determine the value of VMCG? Airport elevation and temperature. The airload on the horizontal tailplane (tailload) of an aeroplane in straight and level cruise flight: is in general directed downwards and will become less negative when the c.g. moves aft. The descent angle of a given aeroplane in a steady wings level glide has a fixed value for a certain combination of: (ignore compressibility effects and assume zero thrust) configuration and angle of attack. In a steady straight climb at climb angle 'gamma', the lift of an aeroplane with weight W is approximately: W * cos (gamma) When an aeroplane performs a straight steady climb with a 20% climb gradient, the load factor is equal to: 0.98. Which of the following parameters can be read from the parabolic polar diagram of an aeroplane? The minimum glide angle and the parasite drag coefficient. In a straight, steady climb the thrust must be: greater than the drag because it must also balance a component of weight. The bank angle in a rate-one turn depends on: TAS. Why is VMCG determined with the nosewheel steering disconnected? Because the value of VMCG must also be applicable on wet and/or slippery runways. The lift to drag ratio determines the: horizontal glide distance from a given altitude at zero wind and zero thrust. How does VMCG change with increasing field elevation and temperature? Decreases, because the engine thrust decreases. An aeroplane's flight path angle is defined as the angle between its: speed vector and the horizontal plane. In a steady, horizontal, co-ordinated turn: thrust equals drag, because there is equilibrium of forces along the direction of flight. An aircraft in flight is affected by loads. These may be classified as: Compressive, tensile, shear and torsional. In order to climb with the speed for maximum climb rate, the aircraft should be flown with the IAS at which: The power excess is maximal. In a steady banked turn the lift will: Equal the resultant of weight and centrifugal force.
•
Which are the two most important parameters to determine the value of VMCG? Engine thrust and rudder deflection.
