ATP Written Test – Aerodynamics

Lift is proportional to the square of the airplane's velocity, if the angle of attack and other factors remain constant. For example, an airplane traveling at 200 kt. has four times the lift of the same airplane traveling at 100 kt.

At an altitude of 18,000 ft., the air has one-half the density of air at sea level. Thus, in order to maintain the same amount of lift as altitude increases, an airplane must fly at a greater true airspeed for any given angle of attack.

The reduction of the wingtip vortices due to ground effect alters the spanwise lift distribution and reduces the induced angle of attack and induced drag. The wing will require a lower angle of attack in ground effect to produce the same lift coefficient as when out of ground effect.

The pilot can control the airplane's lift, airspeed, and drag. As the angle of attack is increased, the lift increases to the stalling angle of attack, airspeed decreases, and induced drag increases with the increase in lift.

Total drag is at a minimum for the maximum lift-drag ratio (L/Dmax) at one specific angle of attack and lift coefficient. As airspeed decreases, the induced drag will increase because a greater angle of attack is required to maintain the lift coefficient. The amount of induced drag varies inversely as the square of the airspeed.

At maximum L/D, total drag is at a minimum. This occurs at a particular angle of attack and lift coefficient at which the proportion between speed and power required is the greatest. Thus, the maximum range and engine-out glide distance for a propeller-driven airplane is at maximum L/D.

The reduction of the wingtip vortices due to ground effect alters the spanwise lift distribution and reduces the induced angle of attack and induced drag. When leaving ground effect, the airplane will require an increase in angle of attack to maintain the same lift coefficient, which will increase induced drag.

Increasing airplane weight increases the induced drag (more lift is required) more than the parasite drag because a higher angle of attack is required to support the increased weight.

Even when jet aircraft do not fly at Mach 1 or greater, airflow over some portions of the upper surface of the wing may exceed Mach 1 and create a shockwave of turbulent air. This turbulent air causes a portion of the wing to partially stall, which causes a substantial increase in drag. Wing-mounted vortex generators are installed to reduce this drag caused by supersonic airflow over portions of the wing.

An airplane's indicated stall speed is affected by total airplane weight, weight distribution within the airplane, angle of bank, pitch attitude, control coordination, drag, and power settings, but NOT by angle of attack or air density.

Turbulence can cause an aircraft to stall at a significantly higher airspeed than in stable conditions. A vertical gust or wind shear can cause a sudden change in the relative wind and result in an abrupt increase in the angle of attack and, thus, result in a stall. Although a gust may not be maintained long enough for a stall to develop, the aircraft may stall while the pilot is attempting to control the flight path, particularly during an approach in gusty conditions.

Airflow separation over the wing can be delayed by various methods of replacing the slow-moving laminar airflow next to the wing's surface with faster-moving air. One method is the use of vortex generators, which are attached to the upper surface of the wing. Vortex generators make the wing surface rough, creating a turbulent flow that increases the speed of the air. Another method is to direct high-pressure air over the
top of the wing, or flap, through slots.

A load factor is the ratio of the total airload acting on the airplane to the gross weight of the airplane. For example, if an airplane has a gross weight of 2,000 lb. and during flight is subjected to aerodynamic forces which increase the total load the wing must support to 4,000 lb., the load factor is 2 (4,000 ?2,000). In this example, the airplane wing is producing lift that is equal to twice the gross weight of the airplane.

Load factor is the ratio of the total airload acting on the airplane to the gross weight of the airplane. In this problem, the total airload is 6,000 lb., and the gross weight is 2,000 lb. The load factor is 3 (6,000 / 2,000). Load factors are usually expressed in terms of "G," or pull of gravity, in this case three times the actual weight.

In any airplane at any airspeed, if a coordinated constant altitude turn is maintained, the load factor for a given degree of bank is constant, which is the resultant load of gravity and centrifugal force.

The load factor for a given airplane during a level coordinated turn is determined solely by the angle of bank.

During a level turn, both total lift and load factor increase.

At speeds below L/Dmax, an increase in pitch will slow the airplane down (i.e., the region of reverse command). At L/Dmax and above, an increase in pitch will increase climb or decrease descent.

The maximum rate-of-climb for a jet airplane will occur at a speed greater than that for L/Dmax. This occurrence is accounted for, in part, by the increase in power available with speed.

Increasing airspeed while maintaining a coordinated turn with a constant angle of bank and altitude will decrease the rate of turn. For any given bank angle, the rate of turn varies with the airspeed; the higher the speed, the slower the rate of turn. This variation compensates for added centrifugal force, allowing the load factor to remain the same.

As the angle of bank is increased, the horizontal component of lift increases, which decreases the vertical component. As this happens, weight is greater than the vertical component of lift, and lift decreases as the sink rate increases.

As airspeed is increased, the angle of attack needs to be decreased, which will cause the rate of turn to decrease. At the same time, the increase in airspeed results in an increase of the turn radius.

At slower airspeeds, an airplane can make a turn in less distance (smaller radius) and at a faster rate. Thus, to decrease the radius and increase the rate, one steepens the bank and decreases airspeed.

The lift during a turn is divided into two components, one vertical and the other horizontal. This division reduces the amount of lift which is opposing gravity. The angle of attack must be increased until the vertical component of lift is equal to the weight in order to maintain altitude.

Longitudinal stability is the quality which makes an airplane stable about its lateral axis. It involves the pitching motion as the airplane's nose moves up and down in flight. Dynamic stability is the overall tendency that the airplane displays after its equilibrium is disturbed.

Longitudinal stability is the quality which makes an airplane stable about its lateral axis (pitching). A longitudinally unstable airplane has a tendency to dive or climb (oscillations) progressively into a very steep dive or climb.

If the airplane's nose remains in the new position after the elevator control is pressed forward and released, the aircraft displays neutral longitudinal static stability.

When the airplane's nose moves farther from its original position, it has the undesirable characteristic of negative static stability.

Imagine that you trim the airplane for hands-off control in level flight and then momentarily give the controls a slight push to nose the airplane down. If within a brief period the nose rises to the original position and then stops, the airplane is statically stable. It is positively statically stable because it has returned to its original position after being displaced.

The airplane will stall at a higher speed with a forward CG location. The critical angle of attack is reached at a higher speed due to increased wing loading. Conversely, the airplane will stall at a lower speed with an aft CG loading. The airplane will cruise faster with an aft CG location because of reduced drag. The airplane will become less stable as the CG is moved rearward.

The rearward CG limit is largely determined by considerations of stability. When loaded to the rear of its CG limit, an airplane will become unstable about the lateral axis, i.e., longitudinally unstable. At this point when the nose is momentarily pulled up, it may alternately climb and dive, becoming steeper with each oscillation.

As a compressor stall occurs, the pressure rises and the compressor does not furnish discharge at a pressure equal to the combustion chamber pressure. As a result, a flow reversal or backfire takes place. If the stall is transient and intermittent, the indication will be the intermittent "bang" as backfire and flow reversal take place.

As a compressor stall occurs, the pressure rises and the compressor does not furnish discharge at a pressure equal to the combustion chamber pressure. As a result, a flow reversal or backfire takes place. If the stall develops and becomes steady, strong vibrations and a loud roar develop from the continuous flow reversal.

In a steady compressor stall, strong vibrations and a loud roar develop from the continuous flow reversal. The increase in compressor power required tends to reduce RPM, and the reduced airflow and increased fuel flow cause a rapid, immediate rise in exhaust gas temperature. The possibility of engine damage is immediate with the steady stall.

Since the possibility of engine damage is immediate if a compressor stall becomes steady, recovery must be accomplished immediately. The recovery from a compressor stall must entail reducing the throttle setting to reduce fuel flow, lowering angle of attack, and increasing airspeed to improve inlet conditions.

Since there is a possibility of having both subsonic and supersonic flows existing on the aircraft, it is convenient to define certain regimes of flight. Subsonic flight range occurs in Mach numbers below 0.75.

Since there is a possibility of having both subsonic and supersonic flows existing on the aircraft, it is convenient to define certain regimes of flight. Transonic flight regimes occur in Mach numbers from 0.75 to 1.20.

By definition, critical Mach number is the "free stream Mach number which produces first evidence of local sonic flow." Therefore, shock waves, buffet, airflow separation, etc., take place above critical Mach number.

The highest speed possible without supersonic flow over the wing is termed the critical Mach number.

When airflow separation occurs on the wing due to shock wave formation, there will be a loss of lift and a subsequent loss of downwash aft of the affected area. If the shock-induced separation occurs symmetrically near the wing root, a decrease in downwash behind this area is caused by a loss of lift. A decrease in downwash on the horizontal tail will create a diving moment and the aircraft will "tuck under."

On a sweptwing airplane, the wing center of pressure shifts when the airplane is shock-stalled. On a sweptwing airplane when the wingtips are shock-stalled first, the center of pressure will move forward and inward.

The sweep of a surface produces a beneficial effect in high speed flight since higher flight speeds may be obtained before components of speed perpendicular to the leading edge produce critical conditions on the wing. One of the most important advantages of a sweepback wing design is an increase in the critical Mach number.

The use of sweptwing design produces certain disadvantages which are important from the standpoint of both airplane design and flight operations. One disadvantage is that, when sweepback is combined with taper, there is an extremely powerful tendency for the wingtip section to stall prior to the wing root.

Sweptwing design contributes to lateral stability by producing stable rolling moments with sideslip. Dutch roll is a coupled lateral-directional oscillation developed when gusts cause a sweptwing-type airplane to roll in one direction while yawing in the other.

Viscous hydroplaning can occur on a thin film of water if it covers the painted or rubber-coated portion of the runway.

Viscous hydroplaning may occur at lower speeds and at lower water depths than dynamic hydroplaning.

The minimum speed in knots for dynamic hydroplaning can be calculated by the following formula:
8.73 x the square root of Tire pressure (in PSI)

Therefore, for a tire pressure of 121 PSI, the minimum speed is 96 kt.
8.73 x q121 = 8.73 x 11 = 96 kt.

The minimum speed in knots for dynamic hydroplaning can be calculated by the following formula:
8.73 x the square root of Tire pressure (in PSI)
8.73 x q70 = 8.73 x 8.36 = 73 kt.

Reverted rubber hydroplaning is the result of a locked wheel skid in which reverted rubber becomes a barrier between the tire and the runway. Steam generated by the heat of friction will hold the wheel off a smooth runway surface.

During hydroplaning, the tire(s) is (are) not in contact with the runway surface, so there is no friction, and application of the brakes would do nothing. Aerodynamic braking must be used to its fullest advantage.

Landing at a higher-than-recommended touchdown speed increases the hydroplaning potential because the wings provide lift and reduce the contact between the tires and the runway, allowing a layer of water to form between the tire and the runway.