FAA Commercial Written Test – Airplanes and Aerodynamics

Raising the flaps decreases the wing camber and the angle of attack of the wing. This decreases wing lift and results in a higher stall speed.

A rectangular wing, as compared to other wing plan forms, has a tendency to stall first at the wing root, with the stall progression toward the wingtip. Because the wingtips and the ailerons stall later, the pilot can use aileron control in avoiding and recovering from the stall.

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

The angle of attack of an airfoil directly controls the distribution of pressure below and above it. When a wing is at a low but positive angle of attack, most of the lift is due to the wing's negative pressure (upper surface) and downwash. Note: Negative pressure is any pressure less than atmospheric, and positive pressure is pressure greater than atmospheric.

Frost on the surface of a wing interferes with the smooth flow of air over the wing surface; i.e., parasite drag is increased. The air flowing over the wing is thus disrupted and stalls at a lower angle of attack (a higher speed) when there is frost on the wing surface.

The angle of attack at which a wing stalls is constant regardless of weight, bank, pitch, etc.

Turbulence, in the form of vertical air currents, can cause severe load stress on a wing. It is wise, in extremely rough air, to reduce the speed to VA (design maneuvering speed). Yet, VA allows a sufficient margin of safety above stall speed, which may be increased in turbulent air due to increased load factors.

Stall speed may vary under different circumstances. Factors such as weight, load factor, power, center of gravity, altitude, temperature, and the presence of snow, ice, or frost on the wings will affect an aircraft's stall speed.

Indicated stall speed is most affected by the gross weight and how it is distributed within the airplane.

An airplane will always stall at the same angle of attack. The airplane's attitude with relation to the horizon has no significance to the stall.

In a rapid recovery from a dive, the load factor would be increased because of the rapid change in the angle of attack, since gravity and centrifugal force would prevent the airplane from immediately altering its flight path. Because the relative wind is opposite the flight path, the critical angle of attack will be reached at a higher airspeed.

The recovery from a stall in any airplane becomes progressively more difficult as the airplane's center of gravity moves aft. This difficulty is due to the decreasing stability in pitch, which results in the decrease of elevator effectiveness in lowering the nose.

Using Fig. 2, work through each of the answers to determine which is true. Power-on stalls occur at lower airspeeds in shallower banks.

The airplane stalls at 76 kt. with power on, gear and flaps up at 60° bank but stalls at 66 kt. with gear and flaps down (i.e., a difference of 10 kt.).

Because rotation is around the CG in a spin, with a rearward CG, the control arm at the rudder is sufficiently shortened that it may make spin recovery difficult, if not impossible. Intuitively, if there is too much weight near the tail, it is also hard to get the nose down to produce an angle of attack below the critical angle.

As the angle of attack is increased, up to the critical angle of attack, the greater the amount of lift is developed, and the greater the induced drag.

At an altitude of 18,000 ft. MSL, 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 be flown at a higher true airspeed for any given angle of attack.

In level flight, all lift is vertical (upwards). As bank is increased, however, a portion of the airplane's lift is transferred from a vertical component to a horizontal component. Thus, the vertical component of lift decreases and the horizontal component of lift increases as the angle of bank is increased.

In any steady-state flight, whether level flight, climbs, or descents, the sum of all forward forces is equal to the sum of all rearward forces, and the upward forces equal the downward forces.

During the transition from straight-and-level flight to a climb, a change in lift occurs as back elevator pressure is first applied, causing an increase in the angle of attack. Lift at this moment is now greater than weight and starts the airplane's climb.

As airspeed decreases, the airfoils generate less lift. Accordingly, to maintain altitude, the angle of attack must be adjusted to compensate for the decrease in lift.

Different angles of attack provide different lift coefficients (amounts of lift). Accordingly, any given angle of attack has a corresponding airspeed to provide sufficient lift to maintain altitude.

Tests show that lift and drag vary as the square of the velocity. The velocity of the air passing over the wing in flight is determined by the airspeed of the airplane. Thus, if an airplane doubles its airspeed, lift and drag will be four times greater (assuming that the angle of attack remains the same).

Total drag is at a minimum for the maximum lift/drag (L/DMAX) ratio 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 level flight. The amount of induced drag varies inversely as the square of the airspeed.

If the airplane is operated in steady flight at L/DMAX, the total drag is at a minimum. Many important items of airplane performance are obtained in flight at L/D MAX. For a propeller-driven airplane, these items include maximum range and maximum power-off glide range.

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

An airplane's lift is produced by a pressure differential resulting from relatively lower (i.e., less than atmospheric) pressure above the wing and higher (i.e., greater than atmospheric) pressure below the wing's surface.

Use Fig. 3 to determine the L/D ratio for a given angle of attack. At the bottom of the chart, locate 10 (i.e., 10° angle of attack) and move vertically up to the L/D curve (the third curve as you move up). Then move right to the margin to determine the L/D ratio of 11:1 (i.e., 1-ft. loss of altitude for every 11 ft. of horizontal distance traveled). Thus, at a distance of 5,280 ft. (1 SM), the airplane will lose 480 ft. (5,280 ?11) of altitude.
Answer (B) is incorrect because the airplane would lose 240 ft. in ? not 1, SM at an angle of attack of 10?

Use Fig. 3 to determine the L/D ratio for a given angle of attack. At the bottom of the chart, locate 8 (i.e., 8° angle of attack) and move vertically up to the L/D curve (the third curve as you move up). Then move right to the margin to determine the L/D ratio of 12:1 (i.e., 1-ft. loss of altitude for every 12 ft. horizontal distance traveled). Thus, at a distance of 15,840 ft. (5,280 ft./SM x 3 SM), the airplane will lose 1,320 ft. (15,840 ? 12) of altitude.

Enter the bottom of the chart in Fig. 3 at 2° angle of attack and move vertically up to the L/D curve. From this point move right horizontally to the point where the L/D curve intersects. Then move vertically down to the bottom of the chart to determine a 16.5° angle of attack. Thus, the L/D ratio is approximately the same at both 2° and 16.5° angle of attack.

Point A (Fig. 1) is at the minimum point on the total drag curve. By definition, this is the point of maximum L/D ratio. Note that airspeed is on the horizontal axis and drag is on the vertical axis.

Point B (Fig. 1) is the intersection of the parasite and induced drag curves, which is the point where the total drag is at its minimum (also known as the point of maximum L/D ratio). L/D MAX is the airspeed at which the pilot of either a jet or a propeller-driven airplane can expect to obtain that airplane's
maximum glide range.

Lift opposes the downward force of weight, is produced by the dynamic effect of the air acting on the wing, and acts perpendicular to the relative wind through the wing's center of lift.

Lift acts perpendicular to the relative wind, which is opposite the flight path. Drag acts parallel to the flight path.

In steady-state level flight, the sum of the opposing forces is equal to zero.

An airplane leaving ground effect (a height greater than the wingspan) will
1. Require an increase in angle of attack to maintain the same lift coefficient,
2. Experience an increase in induced drag and thrust required,
3. Experience a decrease in stability and a nose-up change in moment, and
4. Produce a reduction in static source pressure and increase in indicated airspeed.

In ground effect, induced drag decreases due to a reduction in wingtip vortices (caused by a reduction in the wing's downwash), which alters the spanwise lift distribution and reduces the induced angle of attack. Thus, the wing will require a lower angle of attack in ground effect to produce the same lift as when out of ground effect.

In ground effect, induced drag decreases due to a reduction in wingtip vortices (caused by a reduction in the wing's downwash), which alters the spanwise lift distribution and reduces the induced angle of attack. Thus, if an airplane is brought into ground effect with a constant angle of attack, an increase in lift will result.

When an airplane's attitude is momentarily displaced and it remains at its new attitude, it is said to have neutral longitudinal static stability. Longitudinal stability is the quality which makes an airplane stable about its lateral axis (pitch).

Longitudinal stability is the quality which makes an airplane stable about its lateral (i.e., pitch) axis. This motion is controlled by the elevators.

Dynamic stability is the overall tendency that the airplane displays after its equilibrium is disturbed. Negative dynamic stability (dynamic instability) is a property which causes oscillations set up by a statically stable airplane to become progressively greater. Longitudinal instability refers to pitch oscillations.

When an airplane's elevator control is pressed forward and released and its attitude initially tends to return to its original position, the airplane displays positive static stability.

As the CG is moved rearward, it may move behind the center of lift, in which case the airplane is said to have negative stability about its lateral axis. Recall that the CG should be forward of the center of lift and that the tail surface is designed to have negative lift.

To compensate for the added lift that would result if the airspeed were increased during a turn, the angle of attack must be decreased, or the angle of bank increased, to maintain a constant altitude.

For any given angle of bank, the rate of turn varies with the airspeed. For example, if the angle of bank is held constant and the airspeed is increased, the rate of turn will decrease and vice versa. Because of this, there is no change in centrifugal force while holding a constant angle of bank; thus, the load factor remains constant.

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.

At a constant altitude in a coordinated turn, for each angle of bank each airspeed has a specific rate and radius of turn, and they do not vary.

When in a constant bank in a coordinated turn, an increase in airspeed will decrease the rate of turn. Because the bank is held constant, there will be no change in load factor.

As you enter a turn, lift is divided into horizontal and vertical components. This division reduces the amount of lift which is opposing weight, and thus the airplane loses altitude unless additional lift is created. This is done by increasing back elevator pressure to increase the angle of attack until the vertical component of lift is equal to the weight to maintain altitude without a change in thrust.

As you enter a turn, lift is divided into horizontal and vertical components. This division reduces the amount of lift which is opposing weight, and thus the airplane loses altitude unless additional lift is created. This is done by increasing back elevator pressure to increase the angle of attack until the vertical component of lift is equal to the weight to maintain altitude without a change in thrust.

A standard rate turn is one during which the heading changes at a rate of 3°sec. Thus, a 360° turn would take 2 min. (360°/3°sec. = 120 sec., or 2 min.).

A load factor is the ratio of the total airload acting on the airplane to the gross weight of the airplane. For example, if the airload imposed on the wing is twice the actual weight of the airplane, the load factor is said to be 2 G's, and the stall speed increases.

Since the load factor is the ratio between the total airload imposed on the wing and the gross weight of the airplane, the load factor is the lift generated by the wings divided by the total weight of the airplane. For example, an airplane weighing 2,000 lb. and having a load factor of 2.0 would require 4,000 lb. of lift by the wings. Thus, the load factor of 2.0 is equal to the 4,000 lb. of wing lift divided by the gross weight of 2,000 lb.

In any airplane at any airspeed, if a constant altitude is maintained during the turn, the load factor for a given degree of bank is the same, which is the resultant of weight and centrifugal force. Because of the increased load factor in a turn, the stall speed is also increased in proportion to the square root of the load factor.

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

In a constant-altitude, 60° bank turn, the wings are loaded at 2 G's. Therefore, the total load of a 2,000-lb. airplane is 4,000 lb. (2,000 x 2).

Since the only determinant of load factor in level coordinated turns is the amount of bank, a change in airspeed does not change the load factor. When airspeed is increased, however, the rate of turn decreases and the radius of turn increases.

Since 90 lb. is less than the amount of placarded weight (100 lb.), there is no problem with the weight. The positive load factor of 3.5 G's is within the normal operational limit of 3.8 G's of normal category airplanes. The placarded weight does not have to be divided by the design load factor of the airplane. When 100 lb. was set as a baggage limit in this particular case, the designers recognized that, since this is a normal category airplane, it may be subjected to 3.8 G's, i.e., 380 lb. The baggage weight of 90 lb. is multiplied by 3.5 G's to get a load of 315 lb.

In Fig. 4, the relationship between bank angle degrees on the horizontal axis is related to both load factor or "G" units on the vertical scale and percent increase in stall speed on the vertical axis. There are two curves on the graph. Each curve relates to the vertical scale or the vertical axis. At a 60° bank, find 60° on the horizontal axis, go up to the load factor curve, and then horizontally left to the far left vertical scale to determine approximately 2 G's. At 80° there are  approximately 6 G's. Thus, the increase in load factor is 4 G's (6 - 2) when the angle of bank is increased from 60° to 80°

Use Fig. 4 to determine the percentage increase in stall speed under a load factor of 2 G's. First, find 2 G's on the far left vertical scale and move horizontally to the right to the load factor curve, which intersects at about a 60° bank. Then move vertically up from that point to the intersection of the stall speed increase curve. Next move left horizontally to the vertical axis to determine a 40% increase in stall speed. If the unaccelerated stall speed is 60 kt., the accelerated stall speed is 84 kt. (60 kt. x 140%).