4-1

Forces Acting on the Aircraft

Thrust, drag, lift, and weight are forces that act upon all 

aircraft in flight. Understanding how these forces work and 

knowing how to control them with the use of power and 

flight controls are essential to flight. This chapter discusses 

the aerodynamics of flight—how design, weight, load factors, 

and gravity affect an aircraft during flight maneuvers. 

The four forces acting on an aircraft in straight-and-level, 

unaccelerated flight are thrust, drag, lift, and weight. They 

are defined as follows:

• Thrust—the forward force produced by the powerplant/ 

propeller or rotor. It opposes or overcomes the force 

of drag. As a general rule, it acts parallel to the 

longitudinal axis. However, this is not always the case, 

as explained later.

Aerodynamics 

of Flight

Chapter 4
4-2

Correct relationship of forces acting on an airplane Lift Weight

Thrust Drag

Figure 4-1. Relationship of forces acting on an airplane.

Force vectors during a stabilized climb

Thrust

Drag

Flightpath

Relative wind

Component of weight

opposed to lift 

Rearward component of weight

CG

Figure 4-2. Force vectors during a stabilized climb.

• Drag—a rearward, retarding force caused by 

disruption of airflow by the wing, rotor, fuselage, and 

other protruding objects. Drag opposes thrust, and acts 

rearward parallel to the relative wind.

• Weight—the combined load of the aircraft itself, the 

crew, the fuel, and the cargo or baggage. Weight pulls 

the aircraft downward because of the force of gravity. 

It opposes lift, and acts vertically downward through 

the aircraft’s center of gravity (CG).

• Lift—opposes the downward force of weight, is 

produced by the dynamic effect of the air acting on 

the airfoil, and acts perpendicular to the flightpath 

through the center of lift.

In steady flight, the sum of these opposing forces is always 

zero. There can be no unbalanced forces in steady, straight 

flight based upon Newton’s Third Law, which states that for 

every action or force there is an equal, but opposite, reaction 

or force. This is true whether flying level or when climbing 

or descending. 

It does not mean the four forces are equal. It means the 

opposing forces are equal to, and thereby cancel, the effects 

of each other. In Figure 4-1 the force vectors of thrust, 

drag, lift, and weight appear to be equal in value. The usual 

explanation states (without stipulating that thrust and drag 

do not equal weight and lift) that thrust equals drag and lift 

equals weight. Although basically true, this statement can 

be misleading. It should be understood that in straight, level, 

unaccelerated flight, it is true that the opposing lift/weight 

forces are equal. They are also greater than the opposing 

forces of thrust/drag that are equal only to each other. 

Therefore, in steady flight:

• The sum of all upward forces (not just lift) equals the 

sum of all downward forces (not just weight). 

• The sum of all forward forces (not just thrust) equals 

the sum of all backward forces (not just drag).

This refinement of the old “thrust equals drag; lift equals 

weight” formula explains that a portion of thrust is directed 

upward in climbs and acts as if it were lift while a portion 

of weight is directed backward and acts as if it were drag. 

[Figure 4-2]

In glides, a portion of the weight vector is directed forward, 

and, therefore, acts as thrust. In other words, any time 

the flightpath of the aircraft is not horizontal, lift, weight, 

thrust, and drag vectors must each be broken down into two 

components. 

Discussions of the preceding concepts are frequently omitted 

in aeronautical texts/handbooks/manuals. The reason is not 

that they are inconsequential, but because the main ideas 

with respect to the aerodynamic forces acting upon an 

airplane in flight can be presented in their most essential 

elements without being involved in the technicalities of the 

aerodynamicist. In point of fact, considering only level flight, 

and normal climbs and glides in a steady state, it is still true 

that lift provided by the wing or rotor is the primary upward 

force, and weight is the primary downward force.

By using the aerodynamic forces of thrust, drag, lift, and 

weight, pilots can fly a controlled, safe flight. A more detailed 

discussion of these forces follows.

Thrust

For an aircraft to move, thrust must be exerted and be greater 

than drag. The aircraft will continue to move and gain 

speed until thrust and drag are equal. In order to maintain a 

constant airspeed, thrust and drag must remain equal, just as 

lift and weight must be equal to maintain a constant altitude. 

If in level flight, the engine power is reduced, the thrust is 

lessened, and the aircraft slows down. As long as the thrust 
4-3

Level high speed Level cruise speed Level low speed

Flightpath

Relative wind

Flightpath

Relative wind

Flightpath

Relative wind

3°

6°

12°

Figure 4-3. Angle of attack at various speeds.

Figure 4-4. Some aircraft have the ability to change the direction of thrust.

is less than the drag, the aircraft continues to decelerate until 

its airspeed is insufficient to support it in the air.

Likewise, if the engine power is increased, thrust becomes 

greater than drag and the airspeed increases. As long as 

the thrust continues to be greater than the drag, the aircraft 

continues to accelerate. When drag equals thrust, the aircraft 

flies at a constant airspeed.

Straight-and-level flight may be sustained at a wide range 

of speeds. The pilot coordinates angle of attack (AOA)—the 

acute angle between the chord line of the airfoil and the 

direction of the relative wind—and thrust in all speed regimes 

if the aircraft is to be held in level flight. Roughly, these 

regimes can be grouped in three categories: low-speed flight, 

cruising flight, and high-speed flight.

When the airspeed is low, the AOA must be relatively high 

if the balance between lift and weight is to be maintained. 

[Figure 4-3] If thrust decreases and airspeed decreases, lift 

becomes less than weight and the aircraft starts to descend. 

To maintain level flight, the pilot can increase the AOA 

an amount which will generate a lift force again equal to 

the weight of the aircraft. While the aircraft will be flying 

more slowly, it will still maintain level flight if the pilot has 

properly coordinated thrust and AOA. 

Straight-and-level flight in the slow-speed regime provides 

some interesting conditions relative to the equilibrium of forces 

because with the aircraft in a nose-high attitude, there is a 

vertical component of thrust that helps support it. For one thing, 

wing loading tends to be less than would be expected. Most 

pilots are aware that an airplane will stall, other conditions 

being equal, at a slower speed with the power on than with the 

power off. (Induced airflow over the wings from the propeller 

also contributes to this.) However, if analysis is restricted to 

the four forces as they are usually defined during slow-speed 

flight the thrust is equal to drag, and lift is equal to weight.

During straight-and-level flight when thrust is increased and 

the airspeed increases, the AOA must be decreased. That is, 

if changes have been coordinated, the aircraft will remain in 

level flight, but at a higher speed when the proper relationship 

between thrust and AOA is established.

If the AOA were not coordinated (decreased) with an 

increase of thrust, the aircraft would climb. But decreasing 

the AOA modifies the lift, keeping it equal to the weight, 

and the aircraft remains in level flight. Level flight at even 

slightly negative AOA is possible at very high speed. It is 

evident then, that level flight can be performed with any 

AOA between stalling angle and the relatively small negative 

angles found at high speed.

Some aircraft have the ability to change the direction of the 

thrust rather than changing the AOA. This is accomplished 

either by pivoting the engines or by vectoring the exhaust 

gases. [Figure 4-4]
4-4

Form drag

FLAT PLATE

SPHERE

SPHERE WITH

A FAIRING

SPHERE INSIDE

A HOUSING

Figure 4-5. Form drag.

Figure 4-6. A wing root can cause interference drag.

Drag

Drag is the force that resists movement of an aircraft through 

the air. There are two basic types: parasite drag and induced 

drag. The first is called parasite because it in no way functions 

to aid flight, while the second, induced drag, is a result of an 

airfoil developing lift.

Parasite Drag

Parasite drag is comprised of all the forces that work to slow 

an aircraft’s movement. As the term parasite implies, it is 

the drag that is not associated with the production of lift. 

This includes the displacement of the air by the aircraft, 

turbulence generated in the airstream, or a hindrance of air 

moving over the surface of the aircraft and airfoil. There are 

three types of parasite drag: form drag, interference drag, 

and skin friction. 

Form Drag 

Form drag is the portion of parasite drag generated by the 

aircraft due to its shape and airflow around it. Examples 

include the engine cowlings, antennas, and the aerodynamic 

shape of other components. When the air has to separate 

to move around a moving aircraft and its components, it 

eventually rejoins after passing the body. How quickly 

and smoothly it rejoins is representative of the resistance 

that it creates which requires additional force to overcome. 

[Figure 4-5] 

Notice how the flat plate in Figure 4-5 causes the air to swirl 

around the edges until it eventually rejoins downstream. Form 

drag is the easiest to reduce when designing an aircraft. The 

solution is to streamline as many of the parts as possible.

Interference Drag

Interference drag comes from the intersection of airstreams 

that creates eddy currents, turbulence, or restricts smooth 

airflow. For example, the intersection of the wing and the 

fuselage at the wing root has significant interference drag. 

Air flowing around the fuselage collides with air flowing 

over the wing, merging into a current of air different from 

the two original currents. The most interference drag is 

observed when two surfaces meet at perpendicular angles. 

Fairings are used to reduce this tendency. If a jet fighter 

carries two identical wing tanks, the overall drag is greater 

than the sum of the individual tanks because both of these 

create and generate interference drag. Fairings and distance 

between lifting surfaces and external components (such as 

radar antennas hung from wings) reduce interference drag. 

[Figure 4-6]

Skin Friction Drag

Skin friction drag is the aerodynamic resistance due to the 

contact of moving air with the surface of an aircraft. Every 

surface, no matter how apparently smooth, has a rough, 

ragged surface when viewed under a microscope. The air 

molecules, which come in direct contact with the surface of 

the wing, are virtually motionless. Each layer of molecules 

above the surface moves slightly faster until the molecules 

are moving at the velocity of the air moving around the 

aircraft. This speed is called the free-stream velocity. The 

area between the wing and the free-stream velocity level is 

about as wide as a playing card and is called the boundary 

layer. At the top of the boundary layer, the molecules increase 

velocity and move at the same speed as the molecules 

outside the boundary layer. The actual speed at which the 

molecules move depends upon the shape of the wing, the 

viscosity (stickiness) of the air through which the wing or 

airfoil is moving, and its compressibility (how much it can 

be compacted).
4-5

Figure 4-7. Wingtip vortex from a crop duster.

The airflow outside of the boundary layer reacts to the 

shape of the edge of the boundary layer just as it would 

to the physical surface of an object. The boundary layer 

gives any object an “effective” shape that is usually slightly 

different from the physical shape. The boundary layer may 

also separate from the body, thus creating an effective shape 

much different from the physical shape of the object. This 

change in the physical shape of the boundary layer causes a 

dramatic decrease in lift and an increase in drag. When this 

happens, the airfoil has stalled.

In order to reduce the effect of skin friction drag, aircraft 

designers utilize flush mount rivets and remove any 

irregularities which may protrude above the wing surface. 

In addition, a smooth and glossy finish aids in transition of 

air across the surface of the wing. Since dirt on an aircraft 

disrupts the free flow of air and increases drag, keep the 

surfaces of an aircraft clean and waxed.

Induced Drag

The second basic type of drag is induced drag. It is an 

established physical fact that no system that does work in the 

mechanical sense can be 100 percent efficient. This means 

that whatever the nature of the system, the required work 

is obtained at the expense of certain additional work that is 

dissipated or lost in the system. The more efficient the system, 

the smaller this loss. 

In level flight the aerodynamic properties of a wing or rotor 

produce a required lift, but this can be obtained only at the 

expense of a certain penalty. The name given to this penalty 

is induced drag. Induced drag is inherent whenever an airfoil 

is producing lift and, in fact, this type of drag is inseparable 

from the production of lift. Consequently, it is always present 

if lift is produced. 

An airfoil (wing or rotor blade) produces the lift force by 

making use of the energy of the free airstream. Whenever 

an airfoil is producing lift, the pressure on the lower surface 

of it is greater than that on the upper surface (Bernoulli’s 

Principle). As a result, the air tends to flow from the high 

pressure area below the tip upward to the low pressure area 

on the upper surface. In the vicinity of the tips, there is a 

tendency for these pressures to equalize, resulting in a lateral 

flow outward from the underside to the upper surface. This 

lateral flow imparts a rotational velocity to the air at the tips, 

creating vortices, which trail behind the airfoil. 

When the aircraft is viewed from the tail, these vortices 

circulate counterclockwise about the right tip and clockwise 

about the left tip. [Figure 4-7] Bearing in mind the direction 

of rotation of these vortices, it can be seen that they induce 

an upward flow of air beyond the tip, and a downwash flow 

behind the wing’s trailing edge. This induced downwash has 

nothing in common with the downwash that is necessary to 

produce lift. It is, in fact, the source of induced drag. The 

greater the size and strength of the vortices and consequent 

downwash component on the net airflow over the airfoil, the 

greater the induced drag effect becomes. This downwash over 

the top of the airfoil at the tip has the same effect as bending 

the lift vector rearward; therefore, the lift is slightly aft of 

perpendicular to the relative wind, creating a rearward lift 

component. This is induced drag.

In order to create a greater negative pressure on the top of an 

airfoil, the airfoil can be inclined to a higher AOA. If the AOA 

of a symmetrical airfoil were zero, there would be no pressure 

differential, and consequently, no downwash component and 

no induced drag. In any case, as AOA increases, induced 

drag increases proportionally. To state this another way—the 

lower the airspeed the greater the AOA required to produce 

lift equal to the aircraft’s weight and, therefore, the greater 

induced drag. The amount of induced drag varies inversely 

with the square of the airspeed.

Conversely, parasite drag increases as the square of the 

airspeed. Thus, as airspeed decreases to near the stalling 

speed, the total drag becomes greater, due mainly to the sharp 

rise in induced drag. Similarly, as the airspeed reaches the 

terminal velocity of the aircraft, the total drag again increases 

rapidly, due to the sharp increase of parasite drag. As seen 

in Figure 4-8, at some given airspeed, total drag is at its 

minimum amount. In figuring the maximum endurance and 

range of aircraft, the power required to overcome drag is at 

a minimum if drag is at a minimum.
4-6 Total Drag

Airspeed

Minimum

Drag

Total Drag

Induced Drag

Form

Drag

Figure 4-8. Drag versus speed.

Stall

CL

CLMAX

CD

L/D

L/DMAX

18

16

14

12

10

8

6

4

2

0

.2000

.1800

.1600

.1400

.1200

.1000

.0800

.0600

.0400

.0200

0

Coefficient of drag (CD)

Lift/drag

Critical Angle of attack

0° 2° 4° 6° 8° 10° 12° 14° 16° 18° 20° 22°

CL

1.8

1.6

1.4

1.2

1.0

0.8

0.6

0.4

0.2

0

Figure 4-9. Lift coefficients at various angles of attack.

Lift/Drag Ratio

Drag is the price paid to obtain lift. The lift to drag ratio (L/D) 

is the amount of lift generated by a wing or airfoil compared 

to its drag. A ratio of L/D indicates airfoil efficiency. Aircraft 

with higher L/D ratios are more efficient than those with 

lower L/D ratios. In unaccelerated flight with the lift and drag 

data steady, the proportions of the CL and coefficient of drag 

(CD) can be calculated for specific AOA. [Figure 4-9]

The L/D ratio is determined by dividing the CL by the CD, 

which is the same as dividing the lift equation by the drag 

equation. All terms except coefficients cancel out. 

L = Lift in pounds

D = Drag

Where L is the lift force in pounds, CL is the lift coefficient, 

ρ is density expressed in slugs per cubic feet, V is velocity 

in feet per second, q is dynamic pressure per square feet, and 

S is the wing area in square feet.

CD= Ratio of drag pressure to dynamic pressure. Typically 

at low angles of attack, the drag coefficient is low and small 

changes in angle of attack create only slight changes in the 

drag coefficient. At high angles of attack, small changes in 

the angle of attack cause significant changes in drag.

L = CL . ρ . V2 . S

 2

D = CD . ρ . V2 . S

 2

The above formulas represent the coefficient of lift (CL) 

and the coefficient of drag (CD) respectively. The shape of 

an airfoil and other life producing devices (i.e., flaps) effect 

the production of lift and alter with changes in the AOA. The 

lift/drag ratio is used to express the relation between lift and 

drag and is determined by dividing the lift coefficient by the 

drag coefficient, CL/CD.

Notice in Figure 4-9 that the lift curve (red) reaches its 

maximum for this particular wing section at 20° AOA, and 

then rapidly decreases. 15° AOA is therefore the stalling 

angle. The drag curve (yellow) increases very rapidly from 

14° AOA and completely overcomes the lift curve at 21° 

AOA. The lift/drag ratio (green) reaches its maximum at 6° 

AOA, meaning that at this angle, the most lift is obtained for 

the least amount of drag.

Note that the maximum lift/drag ratio (L/DMAX) occurs at 

one specific CL and AOA. If the aircraft is operated in steady 

flight at L/DMAX, the total drag is at a minimum. Any AOA 

lower or higher than that for L/DMAX reduces the L/D and 
4-7

consequently increases the total drag for a given aircraft’s 

lift. Figure 4-8 depicts the L/DMAX by the lowest portion of 

the orange line labeled “total drag.” The configuration of an 

aircraft has a great effect on the L/D. 

Weight

Gravity is the pulling force that tends to draw all bodies to 

the center of the earth. The CG may be considered as a point 

at which all the weight of the aircraft is concentrated. If the 

aircraft were supported at its exact CG, it would balance in any 

attitude. It will be noted that CG is of major importance in an 

aircraft, for its position has a great bearing upon stability.

The location of the CG is determined by the general design 

of each particular aircraft. The designers determine how far 

the center of pressure (CP) will travel. They then fix the CG 

forward of the center of pressure for the corresponding flight 

speed in order to provide an adequate restoring moment to 

retain flight equilibrium.

Weight has a definite relationship to lift. This relationship 

is simple, but important in understanding the aerodynamics 

of flying. Lift is the upward force on the wing acting 

perpendicular to the relative wind. Lift is required to 

counteract the aircraft’s weight (which is caused by the force 

of gravity acting on the mass of the aircraft). This weight 

(gravity) force acts downward through the airplane’s CG. 

In stabilized level flight, when the lift force is equal to the 

weight force, the aircraft is in a state of equilibrium and 

neither gains nor loses altitude. If lift becomes less than 

weight, the aircraft loses altitude. When lift is greater than 

weight, the aircraft gains altitude.

Lift

The pilot can control the lift. Any time the control yoke 

or stick is moved fore or aft, the AOA is changed. As the 

AOA increases, lift increases (all other factors being equal). 

When the aircraft reaches the maximum AOA, lift begins 

to diminish rapidly. This is the stalling AOA, known as 

CL-MAX critical AOA. Examine Figure 4-9, noting how the 

CL increases until the critical AOA is reached, then decreases 

rapidly with any further increase in the AOA.

Before proceeding further with the topic of lift and how it 

can be controlled, velocity must be interjected. The shape 

of the wing or rotor cannot be effective unless it continually 

keeps “attacking” new air. If an aircraft is to keep flying, the 

lift-producing airfoil must keep moving. In a helicopter or 

gyro-plane this is accomplished by the rotation of the rotor 

blades. For other types of aircraft such as airplanes, weightshift control, or gliders, air must be moving across the lifting 

surface. This is accomplished by the forward speed of the 

aircraft. Lift is proportional to the square of the aircraft’s 

velocity. For example, an airplane traveling at 200 knots has 

four times the lift as the same airplane traveling at 100 knots, 

if the AOA and other factors remain constant. 

Actually, an aircraft could not continue to travel in level 

flight at a constant altitude and maintain the same AOA if 

the velocity is increased. The lift would increase and the 

aircraft would climb as a result of the increased lift force. 

Therefore, to maintain the lift and weight forces in balance, 

and to keep the aircraft straight and level (not accelerating 

upward) in a state of equilibrium, as velocity is increased, 

lift must be decreased. This is normally accomplished by 

reducing the AOA by lowering the nose. Conversely, as the 

aircraft is slowed, the decreasing velocity requires increasing 

the AOA to maintain lift sufficient to maintain flight. There 

is, of course, a limit to how far the AOA can be increased, if 

a stall is to be avoided.

All other factors being constant, for every AOA there is a 

corresponding airspeed required to maintain altitude in steady, 

unaccelerated flight (true only if maintaining “level flight”). 

Since an airfoil always stalls at the same AOA, if increasing 

weight, lift must also be increased. The only method of 

increasing lift is by increasing velocity if the AOA is held 

constant just short of the “critical,” or stalling, AOA.

Lift and drag also vary directly with the density of the air. 

Density is affected by several factors: pressure, temperature, 

and humidity. At an altitude of 18,000 feet, the density of 

the air has one-half the density of air at sea level. In order to 

maintain its lift at a higher altitude, an aircraft must fly at a 

greater true airspeed for any given AOA.

Warm air is less dense than cool air, and moist air is less dense 

than dry air. Thus, on a hot humid day, an aircraft must be 

flown at a greater true airspeed for any given AOA than on 

a cool, dry day.

If the density factor is decreased and the total lift must equal 

the total weight to remain in flight, it follows that one of the 

other factors must be increased. The factor usually increased 

is the airspeed or the AOA, because these are controlled 

directly by the pilot.

Lift varies directly with the wing area, provided there is no 

change in the wing’s planform. If the wings have the same 

proportion and airfoil sections, a wing with a planform area 

of 200 square feet lifts twice as much at the same AOA as a 

wing with an area of 100 square feet.

Two major aerodynamic factors from the pilot’s viewpoint 

are lift and velocity because they can be controlled readily 

and accurately. Of course, the pilot can also control density 
4-8

Vortex

Figure 4-10. Wingtip vortices.

by adjusting the altitude and can control wing area if the 

aircraft happens to have flaps of the type that enlarge wing 

area. However, for most situations, the pilot controls lift and 

velocity to maneuver an aircraft. For instance, in straight-andlevel flight, cruising along at a constant altitude, altitude is 

maintained by adjusting lift to match the aircraft’s velocity 

or cruise airspeed, while maintaining a state of equilibrium 

in which lift equals weight. In an approach to landing, when 

the pilot wishes to land as slowly as practical, it is necessary 

to increase lift to near maximum to maintain lift equal to the 

weight of the aircraft.

Wingtip Vortices

Formation of Vortices

The action of the airfoil that gives an aircraft lift also causes 

induced drag. When an airfoil is flown at a positive AOA, 

a pressure differential exists between the upper and lower 

surfaces of the airfoil. The pressure above the wing is less 

than atmospheric pressure and the pressure below the wing 

is equal to or greater than atmospheric pressure. Since air 

always moves from high pressure toward low pressure, 

and the path of least resistance is toward the airfoil’s tips, 

there is a spanwise movement of air from the bottom of the 

airfoil outward from the fuselage around the tips. This flow 

of air results in “spillage” over the tips, thereby setting up a 

whirlpool of air called a “vortex.” [Figure 4-10]

At the same time, the air on the upper surface has a tendency 

to flow in toward the fuselage and off the trailing edge. This 

air current forms a similar vortex at the inboard portion of the 

trailing edge of the airfoil, but because the fuselage limits the 

inward flow, the vortex is insignificant. Consequently, the 

deviation in flow direction is greatest at the outer tips where 

the unrestricted lateral flow is the strongest. 

As the air curls upward around the tip, it combines with the 

wash to form a fast-spinning trailing vortex. These vortices 

increase drag because of energy spent in producing the 

turbulence. Whenever an airfoil is producing lift, induced 

drag occurs, and wingtip vortices are created.

Just as lift increases with an increase in AOA, induced 

drag also increases. This occurs because as the AOA is 

increased, there is a greater pressure difference between the 

top and bottom of the airfoil, and a greater lateral flow of air; 

consequently, this causes more violent vortices to be set up, 

resulting in more turbulence and more induced drag.

In Figure 4-10, it is easy to see the formation of wingtip 

vortices. The intensity or strength of the vortices is directly 

proportional to the weight of the aircraft and inversely 

proportional to the wingspan and speed of the aircraft. The 

heavier and slower the aircraft, the greater the AOA and the 

stronger the wingtip vortices. Thus, an aircraft will create 

wingtip vortices with maximum strength occurring during 

the takeoff, climb, and landing phases of flight. These 

vortices lead to a particularly dangerous hazard to flight, 

wake turbulence.

Avoiding Wake Turbulence

Wingtip vortices are greatest when the generating aircraft is 

“heavy, clean, and slow.” This condition is most commonly 

encountered during approaches or departures because an 

aircraft’s AOA is at the highest to produce the lift necessary 

to land or take off. To minimize the chances of flying through 

an aircraft’s wake turbulence: 

• Avoid flying through another aircraft’s flightpath.

• Rotate prior to the point at which the preceding aircraft 

rotated, when taking off behind another aircraft.

• Avoid following another aircraft on a similar flightpath 

at an altitude within 1,000 feet. [Figure 4-11] 

• Approach the runway above a preceding aircraft’s 

path when landing behind another aircraft, and touch 

down after the point at which the other aircraft wheels 

contacted the runway. [Figure 4-12]

A hovering helicopter generates a down wash from its main 

rotor(s) similar to the vortices of an airplane. Pilots of small 

aircraft should avoid a hovering helicopter by at least three 

rotor disc diameters to avoid the effects of this down wash. 

In forward flight this energy is transformed into a pair of 

strong, high-speed trailing vortices similar to wing-tip 

vortices of larger fixed-wing aircraft. Helicopter vortices 

should be avoided because helicopter forward flight airspeeds 

are often very slow and can generate exceptionally strong 

wake turbulence.
4-9

500–1,000 ft

Nominally

several hundred ft/min

Sink rate

AVOID

Figure 4-11. Avoid following another aircraft at an altitude within 1,000 feet.

25

Wake ends

Wake begins

Rotation Touchdown

Figure 4-12. Avoid turbulence from another aircraft.

14

Figure 4-13. Ground effect changes airflow.

due to the interference of the ground (or water) surface with 

the airflow patterns about the aircraft in flight.

While the aerodynamic characteristics of the tail surfaces 

and the fuselage are altered by ground effect, the principal 

effects due to proximity of the ground are the changes in 

the aerodynamic characteristics of the wing. As the wing 

encounters ground effect and is maintained at a constant 

lift coefficient, there is consequent reduction in the upwash, 

downwash, and wingtip vortices.

Induced drag is a result of the airfoil’s work of sustaining 

the aircraft, and a wing or rotor lifts the aircraft simply by 

accelerating a mass of air downward. It is true that reduced 

pressure on top of an airfoil is essential to lift, but that is 

Wind is an important factor in avoiding wake turbulence 

because wingtip vortices drift with the wind at the speed of 

the wind. For example, a wind speed of 10 knots causes the 

vortices to drift at about 1,000 feet in a minute in the wind 

direction. When following another aircraft, a pilot should 

consider wind speed and direction when selecting an intended 

takeoff or landing point. If a pilot is unsure of the other 

aircraft’s takeoff or landing point, approximately 3 minutes 

provides a margin of safety that allows wake turbulence 

dissipation. For more information on wake turbulence, see 

Advisory Circular 90-23.

Ground Effect

It is possible to fly an aircraft just clear of the ground (or 

water) at a slightly slower airspeed than that required to 

sustain level flight at higher altitudes. This is the result of a 

phenomenon better known of than understood even by some 

experienced pilots.

When an aircraft in flight comes within several feet of the 

surface, ground or water, a change occurs in the threedimensional flow pattern around the aircraft because the 

vertical component of the airflow around the wing is restricted 

by the surface. This alters the wing’s upwash, downwash, 

and wingtip vortices. [Figure 4-13] Ground effect, then, is 
4-10

Thrust required

Velocity

Lift coefficient 

CL

Angle of attack

Out of ground effect

In ground effect

Out of ground effect

In ground effect

Figure 4-14. Ground effect changes drag and lift.

only one of the things contributing to the overall effect of 

pushing an air mass downward. The more downwash there 

is, the harder the wing pushes the mass of air down. At high 

angles of attack, the amount of induced drag is high; since 

this corresponds to lower airspeeds in actual flight, it can be 

said that induced drag predominates at low speed.

However, the reduction of the wingtip vortices due to ground 

effect alters the spanwise lift distribution and reduces the 

induced AOA and induced drag. Therefore, the wing will 

require a lower AOA in ground effect to produce the same 

CL. If a constant AOA is maintained, an increase in CL results. 

[Figure 4-14]

Ground effect also alters the thrust required versus velocity. 

Since induced drag predominates at low speeds, the reduction 

of induced drag due to ground effect will cause the most 

significant reduction of thrust required (parasite plus induced 

drag) at low speeds.

The reduction in induced flow due to ground effect causes 

a significant reduction in induced drag but causes no direct 

effect on parasite drag. As a result of the reduction in induced 

drag, the thrust required at low speeds will be reduced. Due 

to the change in upwash, downwash, and wingtip vortices, 

there may be a change in position (installation) error of 

the airspeed system, associated with ground effect. In the 

majority of cases, ground effect will cause an increase in 

the local pressure at the static source and produce a lower 

indication of airspeed and altitude. Thus, an aircraft may 

be airborne at an indicated airspeed less than that normally 

required.

In order for ground effect to be of significant magnitude, the 

wing must be quite close to the ground. One of the direct 

results of ground effect is the variation of induced drag with 

wing height above the ground at a constant CL. When the 

wing is at a height equal to its span, the reduction in induced 

drag is only 1.4 percent. However, when the wing is at a 

height equal to one-fourth its span, the reduction in induced 

drag is 23.5 percent and, when the wing is at a height equal 

to one-tenth its span, the reduction in induced drag is 47.6 

percent. Thus, a large reduction in induced drag will take 

place only when the wing is very close to the ground. Because 

of this variation, ground effect is most usually recognized 

during the liftoff for takeoff or just prior to touchdown when 

landing.

During the takeoff phase of flight, ground effect produces 

some important relationships. An aircraft leaving ground 

effect after takeoff encounters just the reverse of an aircraft 

entering ground effect during landing; i.e., the aircraft leaving 

ground effect will:

• Require an increase in AOA to maintain the same 

CL.

• Experience an increase in induced drag and thrust 

required.

• Experience a decrease in stability and a nose-up 

change in moment.

• Experience a reduction in static source pressure and 

increase in indicated airspeed.

Ground effect must be considered during takeoffs and 

landings. For example, if a pilot fails to understand the 

relationship between the aircraft and ground effect during 

takeoff, a hazardous situation is possible because the 

recommended takeoff speed may not be achieved. Due to 

the reduced drag in ground effect, the aircraft may seem 

capable of takeoff well below the recommended speed. As 

the aircraft rises out of ground effect with a deficiency of 

speed, the greater induced drag may result in marginal initial 

climb performance. In extreme conditions, such as high 

gross weight, high density altitude, and high temperature, a 

deficiency of airspeed during takeoff may permit the aircraft 

to become airborne but be incapable of sustaining flight 

out of ground effect. In this case, the aircraft may become 

airborne initially with a deficiency of speed, and then settle 

back to the runway. 
4-11

Pitching Rolling Yawing

Lateral axis Longitudinal axis Vertical axis

Figure 4-15. Axes of an airplane.

Figure 4-16. A weight-shift control aircraft.

A pilot should not attempt to force an aircraft to become 

airborne with a deficiency of speed. The manufacturer’s 

recommended takeoff speed is necessary to provide adequate 

initial climb performance. It is also important that a definite 

climb be established before a pilot retracts the landing gear 

or flaps. Never retract the landing gear or flaps prior to 

establishing a positive rate of climb, and only after achieving 

a safe altitude.

If, during the landing phase of flight, the aircraft is brought into 

ground effect with a constant AOA, the aircraft experiences 

an increase in CL and a reduction in the thrust required, and a 

“floating” effect may occur. Because of the reduced drag and 

power-off deceleration in ground effect, any excess speed at 

the point of flare may incur a considerable “float” distance. 

As the aircraft nears the point of touchdown, ground effect is 

most realized at altitudes less than the wingspan. During the 

final phases of the approach as the aircraft nears the ground, 

a reduced power setting is necessary or the reduced thrust 

required would allow the aircraft to climb above the desired 

glidepath (GP).

Axes of an Aircraft

The axes of an aircraft are three imaginary lines that pass 

through an aircraft’s CG. The axes can be considered as 

imaginary axles around which the aircraft turns. The three axes 

pass through the CG at 90° angles to each other. The axis from 

nose to tail is the longitudinal axis, the axis that passes from 

wingtip to wingtip is the lateral axis, and the axis that passes 

vertically through the CG is the vertical axis. Whenever an 

aircraft changes its flight attitude or position in flight, it rotates 

about one or more of the three axes. [Figure 4-15]

The aircraft’s motion about its longitudinal axis resembles 

the roll of a ship from side to side. In fact, the names 

used to describe the motion about an aircraft’s three axes 

were originally nautical terms. They have been adapted to 

aeronautical terminology due to the similarity of motion of 

aircraft and seagoing ships. The motion about the aircraft’s 

longitudinal axis is “roll,” the motion about its lateral axis is 

“pitch,” and the motion about its vertical axis is “yaw.” Yaw 

is the horizontal (left and right) movement of the aircraft’s 

nose.

The three motions of the conventional airplane (roll, pitch, 

and yaw) are controlled by three control surfaces. Roll is 

controlled by the ailerons; pitch is controlled by the elevators; 

yaw is controlled by the rudder. The use of these controls 

is explained in Chapter 5, Flight Controls. Other types of 

aircraft may utilize different methods of controlling the 

movements about the various axes.

For example, weight-shift control aircraft control two axes, 

roll and pitch, using an “A” frame suspended from the 

flexible wing attached to a three-wheeled carriage. These 

aircraft are controlled by moving a horizontal bar (called a 

control bar) in roughly the same way hang glider pilots fly. 

[Figure 4-16] They are termed weight-shift control aircraft 
4-12

Figure 4-17. A powered parachute.

because the pilot controls the aircraft by shifting the CG. 

For more information on weight-shift control aircraft, see 

the Federal Aviation Administration (FAA) Weight-Shift 

Control Flying Handbook, FAA-H-8083-5. In the case of 

powered parachutes, aircraft control is accomplished by 

altering the airfoil via steering lines. 

A powered parachute wing is a parachute that has a cambered 

upper surface and a flatter under surface. The two surfaces are 

separated by ribs that act as cells, which open to the airflow 

at the leading edge and have internal ports to allow lateral 

airflow. The principle at work holds that the cell pressure is 

greater than the outside pressure, thereby forming a wing that 

maintains its airfoil shape in flight. The pilot and passenger 

sit in tandem in front of the engine which is located at the 

rear of a vehicle. The airframe is attached to the parachute 

via two attachment points and lines. Control is accomplished 

by both power and the changing of the airfoil via the control 

lines. [Figure 4-17]

Moment and Moment Arm

A study of physics shows that a body that is free to rotate 

will always turn about its CG. In aerodynamic terms, the 

mathematical measure of an aircraft’s tendency to rotate 

about its CG is called a “moment.” A moment is said to be 

equal to the product of the force applied and the distance at 

which the force is applied. (A moment arm is the distance 

from a datum [reference point or line] to the applied force.) 

For aircraft weight and balance computations, “moments” 

are expressed in terms of the distance of the arm times the 

aircraft’s weight, or simply, inch-pounds.

Aircraft designers locate the fore and aft position of the 

aircraft’s CG as nearly as possible to the 20 percent point 

of the mean aerodynamic chord (MAC). If the thrust line 

is designed to pass horizontally through the CG, it will not 

cause the aircraft to pitch when power is changed, and there 

will be no difference in moment due to thrust for a power-on 

or power-off condition of flight. Although designers have 

some control over the location of the drag forces, they are not 

always able to make the resultant drag forces pass through the 

CG of the aircraft. However, the one item over which they 

have the greatest control is the size and location of the tail. 

The objective is to make the moments (due to thrust, drag, and 

lift) as small as possible and, by proper location of the tail, 

to provide the means of balancing an aircraft longitudinally 

for any condition of flight.

The pilot has no direct control over the location of forces 

acting on the aircraft in flight, except for controlling the 

center of lift by changing the AOA. Such a change, however, 

immediately involves changes in other forces. Therefore, 

the pilot cannot independently change the location of one 

force without changing the effect of others. For example, 

a change in airspeed involves a change in lift, as well as a 

change in drag and a change in the up or down force on the 

tail. As forces such as turbulence and gusts act to displace 

the aircraft, the pilot reacts by providing opposing control 

forces to counteract this displacement.

Some aircraft are subject to changes in the location of the 

CG with variations of load. Trimming devices are used to 

counteract the forces set up by fuel burnoff, and loading or 

off-loading of passengers or cargo. Elevator trim tabs and 

adjustable horizontal stabilizers comprise the most common 

devices provided to the pilot for trimming for load variations. 

Over the wide ranges of balance during flight in large aircraft, 

the force which the pilot has to exert on the controls would 

become excessive and fatiguing if means of trimming were 

not provided.

Aircraft Design Characteristics 

Each aircraft handles somewhat differently because each 

resists or responds to control pressures in its own way. For 

example, a training aircraft is quick to respond to control 

applications, while a transport aircraft feels heavy on the 

controls and responds to control pressures more slowly. 

These features can be designed into an aircraft to facilitate 

the particular purpose of the aircraft by considering certain 

stability and maneuvering requirements. The following 

discussion summarizes the more important aspects of an 

aircraft’s stability, maneuverability and controllability 

qualities; how they are analyzed; and their relationship to 

various flight conditions.

Stability

Stability is the inherent quality of an aircraft to correct for 

conditions that may disturb its equilibrium, and to return to 

or to continue on the original flightpath. It is primarily an 

aircraft design characteristic. The flightpaths and attitudes an 

aircraft flies are limited by the aerodynamic characteristics of 
4-13

Positive static stability Neutral static stability Negative static stability CGCG CGCG

Applied

Force

Applied

Force

Applied

Force

Figure 4-18. Types of static stability.

the aircraft, its propulsion system, and its structural strength. 

These limitations indicate the maximum performance and 

maneuverability of the aircraft. If the aircraft is to provide 

maximum utility, it must be safely controllable to the full 

extent of these limits without exceeding the pilot’s strength 

or requiring exceptional flying ability. If an aircraft is to fly 

straight and steady along any arbitrary flightpath, the forces 

acting on it must be in static equilibrium. The reaction of 

any body when its equilibrium is disturbed is referred to as 

stability. The two types of stability are static and dynamic.

Static Stability

Static stability refers to the initial tendency, or direction of 

movement, back to equilibrium. In aviation, it refers to the 

aircraft’s initial response when disturbed from a given AOA, 

slip, or bank. 

• Positive static stability—the initial tendency of the 

aircraft to return to the original state of equilibrium 

after being disturbed [Figure 4-18] 

• Neutral static stability—the initial tendency of 

the aircraft to remain in a new condition after its 

equilibrium has been disturbed [Figure 4-18]

• Negative static stability—the initial tendency of the 

aircraft to continue away from the original state of 

equilibrium after being disturbed [Figure 4-18]

Dynamic Stability

Static stability has been defined as the initial tendency to 

return to equilibrium that the aircraft displays after being 

disturbed from its trimmed condition. Occasionally, the 

initial tendency is different or opposite from the overall 

tendency, so a distinction must be made between the two. 

Dynamic stability refers to the aircraft response over time 

when disturbed from a given AOA, slip, or bank. This type 

of stability also has three subtypes: [Figure 4-19]

• Positive dynamic stability—over time, the motion 

of the displaced object decreases in amplitude and, 

because it is positive, the object displaced returns 

toward the equilibrium state.

• Neutral dynamic stability—once displaced, the 

displaced object neither decreases nor increases in 

amplitude. A worn automobile shock absorber exhibits 

this tendency.

• Negative dynamic stability—over time, the motion 

of the displaced object increases and becomes more 

divergent.

Stability in an aircraft affects two areas significantly:

• Maneuverability—the quality of an aircraft that 

permits it to be maneuvered easily and to withstand 

the stresses imposed by maneuvers. It is governed by 

the aircraft’s weight, inertia, size and location of flight 

controls, structural strength, and powerplant. It too is 

an aircraft design characteristic.

• Controllability—the capability of an aircraft to 

respond to the pilot’s control, especially with regard to 

flightpath and attitude. It is the quality of the aircraft’s 

response to the pilot’s control application when 

maneuvering the aircraft, regardless of its stability 

characteristics.
4-14

Reduced power allows pitch down CG CL CG CL

Figure 4-20. Longitudinal stability.

Displacement

Positive static

(neutral dynamic)

Positive static

(positive dynamic)

Positive Static

(negative dynamic)

Time

Damped oscillation

Undamped oscillation

Divergent oscillation

Figure 4-19. Damped versus undamped stability.

Longitudinal Stability (Pitching)

In designing an aircraft, a great deal of effort is spent in 

developing the desired degree of stability around all three 

axes. But longitudinal stability about the lateral axis is 

considered to be the most affected by certain variables in 

various flight conditions.

Longitudinal stability is the quality that makes an aircraft 

stable about its lateral axis. It involves the pitching motion 

as the aircraft’s nose moves up and down in flight. A 

longitudinally unstable aircraft has a tendency to dive or 

climb progressively into a very steep dive or climb, or even 

a stall. Thus, an aircraft with longitudinal instability becomes 

difficult and sometimes dangerous to fly.

Static longitudinal stability or instability in an aircraft, is 

dependent upon three factors:

1. Location of the wing with respect to the CG

2. Location of the horizontal tail surfaces with respect 

to the CG

3. Area or size of the tail surfaces

In analyzing stability, it should be recalled that a body free 

to rotate always turns about its CG.

To obtain static longitudinal stability, the relation of the 

wing and tail moments must be such that, if the moments 

are initially balanced and the aircraft is suddenly nose up, 

the wing moments and tail moments change so that the sum 

of their forces provides an unbalanced but restoring moment 

which, in turn, brings the nose down again. Similarly, if the 

aircraft is nose down, the resulting change in moments brings 

the nose back up.

The CL in most asymmetrical airfoils has a tendency to 

change its fore and aft positions with a change in the AOA. 

The CL tends to move forward with an increase in AOA and 

to move aft with a decrease in AOA. This means that when the 

AOA of an airfoil is increased, the CL, by moving forward, 

tends to lift the leading edge of the wing still more. This 

tendency gives the wing an inherent quality of instability. 

(NOTE: CL is also known as the center of pressure (CP).)

Figure 4-20 shows an aircraft in straight-and-level flight. The 

line CG-CL-T represents the aircraft’s longitudinal axis from 

the CG to a point T on the horizontal stabilizer. 

Most aircraft are designed so that the wing’s CL is to the rear 

of the CG. This makes the aircraft “nose heavy” and requires 

that there be a slight downward force on the horizontal 

stabilizer in order to balance the aircraft and keep the nose 

from continually pitching downward. Compensation for this 

nose heaviness is provided by setting the horizontal stabilizer 

at a slight negative AOA. The downward force thus produced 

holds the tail down, counterbalancing the “heavy” nose. It 
4-15

Reduced power allows pitch down

Cruise speed

Low speed

High speed

CG

CG

CG

Balanced tail load

Lesser downward tail load

Greater downward tail load

Figure 4-21. Effect of speed on downwash.

Reduced power allows pitch down

Normal downwash

Reduced downwash

Thrust

Weight

CG

Lift

Thrust

Weight

CG

Lift

Figure 4-22. Reduced power allows pitch down.

is as if the line CG-CL-T were a lever with an upward force 

at CL and two downward forces balancing each other, one 

a strong force at the CG point and the other, a much lesser 

force, at point T (downward air pressure on the stabilizer). 

To better visualize this physics principle: If an iron bar were 

suspended at point CL, with a heavy weight hanging on it at 

the CG, it would take downward pressure at point T to keep 

the “lever” in balance.

Even though the horizontal stabilizer may be level when the 

aircraft is in level flight, there is a downwash of air from the 

wings. This downwash strikes the top of the stabilizer and 

produces a downward pressure, which at a certain speed is 

just enough to balance the “lever.” The faster the aircraft 

is flying, the greater this downwash and the greater the 

downward force on the horizontal stabilizer (except T-tails). 

[Figure 4-21] In aircraft with fixed-position horizontal 

stabilizers, the aircraft manufacturer sets the stabilizer at an 

angle that provides the best stability (or balance) during flight 

at the design cruising speed and power setting. 

If the aircraft’s speed decreases, the speed of the airflow 

over the wing is decreased. As a result of this decreased 

flow of air over the wing, the downwash is reduced, causing 

a lesser downward force on the horizontal stabilizer. In turn, 

the characteristic nose heaviness is accentuated, causing the 

aircraft’s nose to pitch down more. [Figure 4-22] This places 

the aircraft in a nose-low attitude, lessening the wing’s AOA 

and drag and allowing the airspeed to increase. As the aircraft 

continues in the nose-low attitude and its speed increases, 

the downward force on the horizontal stabilizer is once again 

increased. Consequently, the tail is again pushed downward 

and the nose rises into a climbing attitude.

As this climb continues, the airspeed again decreases, causing 

the downward force on the tail to decrease until the nose 

lowers once more. Because the aircraft is dynamically stable, 

the nose does not lower as far this time as it did before. The 

aircraft acquires enough speed in this more gradual dive to 

start it into another climb, but the climb is not as steep as 

the preceding one.

After several of these diminishing oscillations, in which 

the nose alternately rises and lowers, the aircraft finally 

settles down to a speed at which the downward force on the 

tail exactly counteracts the tendency of the aircraft to dive. 

When this condition is attained, the aircraft is once again in 

balanced flight and continues in stabilized flight as long as 

this attitude and airspeed are not changed.

A similar effect is noted upon closing the throttle. The 

downwash of the wings is reduced and the force at T in 

Figure 4-20 is not enough to hold the horizontal stabilizer 

down. It seems as if the force at T on the lever were allowing 

the force of gravity to pull the nose down. This is a desirable 

characteristic because the aircraft is inherently trying to 

regain airspeed and reestablish the proper balance.

Power or thrust can also have a destabilizing effect in that 

an increase of power may tend to make the nose rise. The 

aircraft designer can offset this by establishing a “high 

thrust line” wherein the line of thrust passes above the CG. 

[Figures 4-23 and 4-24] In this case, as power or thrust is 

increased a moment is produced to counteract the down 
4-16

Thrust line affect longitudinal stability

Below center of gravity

Through center of gravity

Above center of gravity

Thrust CG

Thrust CG

Thrust

CG

Figure 4-23. Thrust line affects longitudinal stability.

Power changes affect longitudinal stability

Cruise power

Idle power

Full power Lift

Thrust CGLift

Thrust CGLift

Thrust CG

Figure 4-24. Power changes affect longitudinal stability.

load on the tail. On the other hand, a very “low thrust line” 

would tend to add to the nose-up effect of the horizontal tail 

surface.

Conclusion: with CG forward of the CL and with an 

aerodynamic tail-down force, the aircraft usually tries to 

return to a safe flying attitude.

The following is a simple demonstration of longitudinal 

stability. Trim the aircraft for “hands off” control in level 

flight. Then, momentarily give the controls a slight push to 

nose the aircraft down. If, within a brief period, the nose rises 

to the original position and then stops, the aircraft is statically 

stable. Ordinarily, the nose passes the original position (that 

of level flight) and a series of slow pitching oscillations 

follows. If the oscillations gradually cease, the aircraft has 

positive stability; if they continue unevenly, the aircraft has 

neutral stability; if they increase, the aircraft is unstable.

Lateral Stability (Rolling)

Stability about the aircraft’s longitudinal axis, which extends 

from the nose of the aircraft to its tail, is called lateral 

stability. This helps to stabilize the lateral or “rolling effect” 

when one wing gets lower than the wing on the opposite side 

of the aircraft. There are four main design factors that make 

an aircraft laterally stable: dihedral, sweepback, keel effect, 

and weight distribution.

Dihedral

The most common procedure for producing lateral stability 

is to build the wings with an angle of one to three degrees 

above perpendicular to the longitudinal axis. The wings on 

either side of the aircraft join the fuselage to form a slight V or 

angle called “dihedral.” The amount of dihedral is measured 

by the angle made by each wing above a line parallel to the 

lateral axis.

Dihedral involves a balance of lift created by the wings’ AOA 

on each side of the aircraft’s longitudinal axis. If a momentary 

gust of wind forces one wing to rise and the other to lower, the 

aircraft banks. When the aircraft is banked without turning, 

the tendency to sideslip or slide downward toward the lowered 

wing occurs. [Figure 4-25] Since the wings have dihedral, 

the air strikes the lower wing at a much greater AOA than the 

higher wing. The increased AOA on the lower wing creates 

more lift than the higher wing. Increased lift causes the lower 

wing to begin to rise upward. As the wings approach the 

level position, the AOA on both wings once again are equal, 

causing the rolling tendency to subside. The effect of dihedral 

is to produce a rolling tendency to return the aircraft to a 

laterally balanced flight condition when a sideslip occurs.

The restoring force may move the low wing up too far, so 

that the opposite wing now goes down. If so, the process 

is repeated, decreasing with each lateral oscillation until a 

balance for wings-level flight is finally reached.
4-17

Normal angle of attack

Dihedral lateral stability

Lesser angle

of attack

Greater angle

of attack

Figure 4-25. Dihedral for lateral stability.

Keel area for lateral stability

CG

CG

CG centerline

Figure 4-26. Keel area for lateral stability.

Conversely, excessive dihedral has an adverse effect on 

lateral maneuvering qualities. The aircraft may be so stable 

laterally that it resists an intentional rolling motion. For this 

reason, aircraft that require fast roll or banking characteristics 

usually have less dihedral than those designed for less 

maneuverability.

Sweepback

Sweepback is an addition to the dihedral that increases the 

lift created when a wing drops from the level position. A 

sweptback wing is one in which the leading edge slopes 

backward. When a disturbance causes an aircraft with 

sweepback to slip or drop a wing, the low wing presents its 

leading edge at an angle that is perpendicular to the relative 

airflow. As a result, the low wing acquires more lift, rises, 

and the aircraft is restored to its original flight attitude.

Sweepback also contributes to directional stability. When 

turbulence or rudder application causes the aircraft to yaw 

to one side, the right wing presents a longer leading edge 

perpendicular to the relative airflow. The airspeed of the right 

wing increases and it acquires more drag than the left wing. 

The additional drag on the right wing pulls it back, turning 

the aircraft back to its original path.

Keel Effect and Weight Distribution

An aircraft always has the tendency to turn the longitudinal 

axis of the aircraft into the relative wind. This “weather vane” 

tendency is similar to the keel of a ship and exerts a steadying 

influence on the aircraft laterally about the longitudinal 

axis. When the aircraft is disturbed and one wing dips, the 

fuselage weight acts like a pendulum returning the airplane 

to its original attitude. 

Laterally stable aircraft are constructed so that the greater 

portion of the keel area is above and behind the CG. 

[Figure 4-26] Thus, when the aircraft slips to one side, the 

combination of the aircraft’s weight and the pressure of the 

airflow against the upper portion of the keel area (both acting 

about the CG) tends to roll the aircraft back to wings-level 

flight.

Vertical Stability (Yawing)

Stability about the aircraft’s vertical axis (the sideways 

moment) is called yawing or directional stability. Yawing or 

directional stability is the most easily achieved stability in 

aircraft design. The area of the vertical fin and the sides of 

the fuselage aft of the CG are the prime contributors which 

make the aircraft act like the well known weather vane or 

arrow, pointing its nose into the relative wind.

In examining a weather vane, it can be seen that if exactly the 

same amount of surface were exposed to the wind in front 

of the pivot point as behind it, the forces fore and aft would 

be in balance and little or no directional movement would 

result. Consequently, it is necessary to have a greater surface 

aft of the pivot point than forward of it.

Similarly, the aircraft designer must ensure positive 

directional stability by making the side surface greater aft 

than ahead of the CG. [Figure 4-27] To provide additional 

positive stability to that provided by the fuselage, a vertical 

fin is added. The fin acts similar to the feather on an arrow 

in maintaining straight flight. Like the weather vane and the 

arrow, the farther aft this fin is placed and the larger its size, 

the greater the aircraft’s directional stability.

If an aircraft is flying in a straight line, and a sideward gust 

of air gives the aircraft a slight rotation about its vertical 

axis (i.e., the right), the motion is retarded and stopped by 

the fin because while the aircraft is rotating to the right, the 

air is striking the left side of the fin at an angle. This causes 

pressure on the left side of the fin, which resists the turning 

motion and slows down the aircraft’s yaw. In doing so, it 
4-18

Fuselage and fin for vertical stability

Area 

forward

of CG Area after center of gravity (CG)

CG

CG

Relative wind

yaw

yaw

Figure 4-27. Fuselage and fin for vertical stability.

acts somewhat like the weather vane by turning the aircraft 

into the relative wind. The initial change in direction of the 

aircraft’s flightpath is generally slightly behind its change 

of heading. Therefore, after a slight yawing of the aircraft 

to the right, there is a brief moment when the aircraft is still 

moving along its original path, but its longitudinal axis is 

pointed slightly to the right.

The aircraft is then momentarily skidding sideways, and 

during that moment (since it is assumed that although the 

yawing motion has stopped, the excess pressure on the left 

side of the fin still persists) there is necessarily a tendency 

for the aircraft to be turned partially back to the left. That is, 

there is a momentary restoring tendency caused by the fin.

This restoring tendency is relatively slow in developing and 

ceases when the aircraft stops skidding. When it ceases, the 

aircraft is flying in a direction slightly different from the 

original direction. In other words, it will not return of its own 

accord to the original heading; the pilot must reestablish the 

initial heading.

A minor improvement of directional stability may be obtained 

through sweepback. Sweepback is incorporated in the design 

of the wing primarily to delay the onset of compressibility 

during high-speed flight. In lighter and slower aircraft, 

sweepback aids in locating the center of pressure in the 

correct relationship with the CG. A longitudinally stable 

aircraft is built with the center of pressure aft of the CG.

Because of structural reasons, aircraft designers sometimes 

cannot attach the wings to the fuselage at the exact desired 

point. If they had to mount the wings too far forward, and at 

right angles to the fuselage, the center of pressure would not 

be far enough to the rear to result in the desired amount of 

longitudinal stability. By building sweepback into the wings, 

however, the designers can move the center of pressure 

toward the rear. The amount of sweepback and the position 

of the wings then place the center of pressure in the correct 

location.

The contribution of the wing to static directional stability is 

usually small. The swept wing provides a stable contribution 

depending on the amount of sweepback, but the contribution 

is relatively small when compared with other components.

Free Directional Oscillations (Dutch Roll)

Dutch roll is a coupled lateral/directional oscillation that is 

usually dynamically stable but is unsafe in an aircraft because 

of the oscillatory nature. The damping of the oscillatory mode 

may be weak or strong depending on the properties of the 

particular aircraft.

If the aircraft has a right wing pushed down, the positive 

sideslip angle corrects the wing laterally before the nose is 

realigned with the relative wind. As the wing corrects the 

position, a lateral directional oscillation can occur resulting 

in the nose of the aircraft making a figure eight on the 

horizon as a result of two oscillations (roll and yaw), which, 

although of about the same magnitude, are out of phase with 

each other.

In most modern aircraft, except high-speed swept wing 

designs, these free directional oscillations usually die out 

automatically in very few cycles unless the air continues to 

be gusty or turbulent. Those aircraft with continuing Dutch 

roll tendencies are usually equipped with gyro-stabilized yaw 

dampers. Manufacturers try to reach a midpoint between too 

much and too little directional stability. Because it is more 

desirable for the aircraft to have “spiral instability” than 

Dutch roll tendencies, most aircraft are designed with that 

characteristic.

Spiral Instability

Spiral instability exists when the static directional stability 

of the aircraft is very strong as compared to the effect of its 

dihedral in maintaining lateral equilibrium. When the lateral 

equilibrium of the aircraft is disturbed by a gust of air and a 

sideslip is introduced, the strong directional stability tends 

to yaw the nose into the resultant relative wind while the 

comparatively weak dihedral lags in restoring the lateral 

balance. Due to this yaw, the wing on the outside of the 
4-19

Level flight Medium banked turn Steeply banked turn

Centrifugal

 force

Horizontal

component

Verticalcomponent

Resultant load

Centrifugal

force

Horizontal

component

Verticalcomponent

Resultant load

Lift

Total lift

Total lift

WeightWeightWeight

Figure 4-28. Forces during normal coordinated turn.

turning moment travels forward faster than the inside wing 

and, as a consequence, its lift becomes greater. This produces 

an overbanking tendency which, if not corrected by the pilot, 

results in the bank angle becoming steeper and steeper. At 

the same time, the strong directional stability that yaws the 

aircraft into the relative wind is actually forcing the nose 

to a lower pitch attitude. A slow downward spiral begins 

which, if not counteracted by the pilot, gradually increases 

into a steep spiral dive. Usually the rate of divergence in the 

spiral motion is so gradual the pilot can control the tendency 

without any difficulty.

All aircraft are affected to some degree by this characteristic, 

although they may be inherently stable in all other normal 

parameters. This tendency explains why an aircraft cannot 

be flown “hands off” indefinitely.

Much research has gone into the development of control 

devices (wing leveler) to correct or eliminate this instability. 

The pilot must be careful in application of recovery controls 

during advanced stages of this spiral condition or excessive 

loads may be imposed on the structure. Improper recovery 

from spiral instability leading to inflight structural failures 

has probably contributed to more fatalities in general aviation 

aircraft than any other factor. Since the airspeed in the spiral 

condition builds up rapidly, the application of back elevator 

force to reduce this speed and to pull the nose up only 

“tightens the turn,” increasing the load factor. The results 

of the prolonged uncontrolled spiral are inflight structural 

failure or crashing into the ground, or both. The most common 

recorded causes for pilots who get into this situation are: 

loss of horizon reference, inability to control the aircraft by 

reference to instruments, or a combination of both. 

Aerodynamic Forces in Flight Maneuvers

Forces in Turns

If an aircraft were viewed in straight-and-level flight from 

the front [Figure 4-28], and if the forces acting on the aircraft 

could be seen, lift and weight would be apparent: two forces. 

If the aircraft were in a bank it would be apparent that lift 

did not act directly opposite to the weight, rather it now acts 

in the direction of the bank. A basic truth about turns: when 

the aircraft banks, lift acts inward toward the center of the 

turn, as well as upward.

Newton’s First Law of Motion, the Law of Inertia, states 

that an object at rest or moving in a straight line remains 

at rest or continues to move in a straight line until acted on 

by some other force. An aircraft, like any moving object, 

requires a sideward force to make it turn. In a normal turn, 

this force is supplied by banking the aircraft so that lift is 

exerted inward, as well as upward. The force of lift during a 

turn is separated into two components at right angles to each 

other. One component, which acts vertically and opposite 

to the weight (gravity), is called the “vertical component of 

lift.” The other, which acts horizontally toward the center 

of the turn, is called the “horizontal component of lift,” or 

centripetal force. The horizontal component of lift is the force 

that pulls the aircraft from a straight flightpath to make it 

turn. Centrifugal force is the “equal and opposite reaction” 

of the aircraft to the change in direction and acts equal and 

opposite to the horizontal component of lift. This explains 

why, in a correctly executed turn, the force that turns the 

aircraft is not supplied by the rudder. The rudder is used to 

correct any deviation between the straight track of the nose 

and tail of the aircraft. A good turn is one in which the nose 

and tail of the aircraft track along the same path. If no rudder 

is used in a turn, the nose of the aircraft yaws to the outside 

of the turn. The rudder is used to bring the nose back in line 

with the relative wind.

An aircraft is not steered like a boat or an automobile. In 

order for an aircraft to turn, it must be banked. If it is not 

banked, there is no force available to cause it to deviate from 

a straight flightpath. Conversely, when an aircraft is banked, 

it turns, provided it is not slipping to the inside of the turn. 
4-20

Normal turn Slipping turn Skidding turn

Centrifugal 

force equals 

horizontal lift

Centrifugal 

force less than 

horizontal lift

Centrifugal force 

greater than 

horizontal lift

Vertical lift

Vertical lift

Vertical lift

Load

Load

Load

Centrifugal

force

Horizontal

lift 

Centrifugal

force

Horizontal

lift 

Centrifugal force

Horizontal

lift 

Lift

Lift

Lift

WeightWeightWeight

Figure 4-29. Normal, slipping, and skidding turns.

Good directional control is based on the fact that the aircraft 

attempts to turn whenever it is banked. Pilots should keep 

this fact in mind when attempting to hold the aircraft in 

straight-and-level flight.

Merely banking the aircraft into a turn produces no change 

in the total amount of lift developed. Since the lift during the 

bank is divided into vertical and horizontal components, the 

amount of lift opposing gravity and supporting the aircraft’s 

weight is reduced. Consequently, the aircraft loses altitude 

unless additional lift is created. This is done by increasing 

the AOA until the vertical component of lift is again equal 

to the weight. Since the vertical component of lift decreases 

as the bank angle increases, the AOA must be progressively 

increased to produce sufficient vertical lift to support the 

aircraft’s weight. An important fact for pilots to remember 

when making constant altitude turns is that the vertical 

component of lift must be equal to the weight to maintain 

altitude.

At a given airspeed, the rate at which an aircraft turns depends 

upon the magnitude of the horizontal component of lift. It is 

found that the horizontal component of lift is proportional 

to the angle of bank—that is, it increases or decreases 

respectively as the angle of bank increases or decreases. As 

the angle of bank is increased, the horizontal component of 

lift increases, thereby increasing the ROT. Consequently, at 

any given airspeed, the ROT can be controlled by adjusting 

the angle of bank.

To provide a vertical component of lift sufficient to hold 

altitude in a level turn, an increase in the AOA is required. 

Since the drag of the airfoil is directly proportional to its 

AOA, induced drag increases as the lift is increased. This, 

in turn, causes a loss of airspeed in proportion to the angle 

of bank. A small angle of bank results in a small reduction 

in airspeed while a large angle of bank results in a large 

reduction in airspeed. Additional thrust (power) must be 

applied to prevent a reduction in airspeed in level turns. The 

required amount of additional thrust is proportional to the 

angle of bank.

To compensate for added lift, which would result if the 

airspeed were increased during a turn, the AOA must be 

decreased, or the angle of bank increased, if a constant altitude 

is to be maintained. If the angle of bank is held constant and 

the AOA decreased, the ROT decreases. In order to maintain 

a constant-ROT as the airspeed is increased, the AOA must 

remain constant and the angle of bank increased.

An increase in airspeed results in an increase of the turn 

radius, and centrifugal force is directly proportional to the 

radius of the turn. In a correctly executed turn, the horizontal 

component of lift must be exactly equal and opposite to the 

centrifugal force. As the airspeed is increased in a constantrate level turn, the radius of the turn increases. This increase 

in the radius of turn causes an increase in the centrifugal 

force, which must be balanced by an increase in the horizontal 

component of lift, which can only be increased by increasing 

the angle of bank.

In a slipping turn, the aircraft is not turning at the rate 

appropriate to the bank being used, since the aircraft is 

yawed toward the outside of the turning flightpath. The 

aircraft is banked too much for the ROT, so the horizontal lift 

component is greater than the centrifugal force. [Figure 4-29]

Equilibrium between the horizontal lift component and 

centrifugal force is reestablished by either decreasing the 

bank, increasing the ROT, or a combination of the two 

changes.

A skidding turn results from an excess of centrifugal force 

over the horizontal lift component, pulling the aircraft 

toward the outside of the turn. The ROT is too great for the 

angle of bank. Correction of a skidding turn thus involves a 
4-21

Steady climb

normal lift

Climb entry

increased lift

Level flight 

normal lift

L

L

L

L

Figure 4-30. Changes in lift during climb entry.

T

D

T

D T D

Steady climb

forces balanced

constant speed

Climb entry drag

greater than thrust

speed slowing

Level flight

forces balanced

constant speed

W

L W

W L

L

Figure 4-31. Changes in speed during climb entry.

reduction in the ROT, an increase in bank, or a combination 

of the two changes.

To maintain a given ROT, the angle of bank must be varied 

with the airspeed. This becomes particularly important in 

high-speed aircraft. For instance, at 400 miles per hour (mph), 

an aircraft must be banked approximately 44° to execute a 

standard-rate turn (3° per second). At this angle of bank, 

only about 79 percent of the lift of the aircraft comprises the 

vertical component of the lift. This causes a loss of altitude 

unless the AOA is increased sufficiently to compensate for 

the loss of vertical lift.

Forces in Climbs

For all practical purposes, the wing’s lift in a steady state 

normal climb is the same as it is in a steady level flight at 

the same airspeed. Although the aircraft’s flightpath changed 

when the climb was established, the AOA of the wing with 

respect to the inclined flightpath reverts to practically the 

same values, as does the lift. There is an initial momentary 

change as shown in Figure 4-30. During the transition from 

straight-and-level flight to a climb, a change in lift occurs 

when back elevator pressure is first applied. Raising the 

aircraft’s nose increases the AOA and momentarily increases 

the lift. Lift at this moment is now greater than weight and 

starts the aircraft climbing. After the flightpath is stabilized 

on the upward incline, the AOA and lift again revert to about 

the level flight values.

If the climb is entered with no change in power setting, the 

airspeed gradually diminishes because the thrust required 

to maintain a given airspeed in level flight is insufficient to 

maintain the same airspeed in a climb. When the flightpath 

is inclined upward, a component of the aircraft’s weight 

acts in the same direction as, and parallel to, the total drag 

of the aircraft, thereby increasing the total effective drag. 

Consequently, the total drag is greater than the power, and 

the airspeed decreases. The reduction in airspeed gradually 

results in a corresponding decrease in drag until the total 

drag (including the component of weight acting in the same 

direction) equals the thrust. [Figure 4-31] Due to momentum, 

the change in airspeed is gradual, varying considerably with 

differences in aircraft size, weight, total drag, and other 

factors. Consequently, the total drag is greater than the thrust, 

and the airspeed decreases.

Generally, the forces of thrust and drag, and lift and weight, 

again become balanced when the airspeed stabilizes but at 

a value lower than in straight-and-level flight at the same 

power setting. Since the aircraft’s weight is acting not only 

downward but rearward with drag while in a climb, additional 

power is required to maintain the same airspeed as in level 

flight. The amount of power depends on the angle of climb. 

When the climb is established steep enough that there is 

insufficient power available, a slower speed results. 

The thrust required for a stabilized climb equals drag plus a 

percentage of weight dependent on the angle of climb. For 

example, a 10° climb would require thrust to equal drag plus 

17 percent of weight. To climb straight up would require 

thrust to equal all of weight and drag. Therefore, the angle 

of climb for climb performance is dependent on the amount 

of excess power available to overcome a portion of weight. 

Note that aircraft are able to sustain a climb due to excess 

thrust. When the excess thrust is gone, the aircraft is no 

longer able to climb. At this point, the aircraft has reached 

its “absolute ceiling.”

Forces in Descents

As in climbs, the forces which act on the aircraft go through 

definite changes when a descent is entered from straightand-level flight. For the following example, the aircraft is 

descending at the same power as used in straight-and-level 

flight.

As forward pressure is applied to the control yoke to initiate 

the descent, the AOA is decreased momentarily. Initially, 

the momentum of the aircraft causes the aircraft to briefly 
4-22

CF

W

W

CF

L

L

CF

CF

W

L

Figure 4-32. Forces exerted when pulling out of a dive.

continue along the same flightpath. For this instant, the AOA 

decreases causing the total lift to decrease. With weight now 

being greater than lift, the aircraft begins to descend. At the 

same time, the flightpath goes from level to a descending 

flightpath. Do not confuse a reduction in lift with the 

inability to generate sufficient lift to maintain level flight. 

The flightpath is being manipulated with available thrust in 

reserve and with the elevator.

To descend at the same airspeed as used in straight-andlevel flight, the power must be reduced as the descent is 

entered. The component of weight acting forward along the 

flightpath increases as the angle of rate of descent increases 

and, conversely, decreases as the angle of rate of descent 

decreases. The component of weight acting forward along the 

flightpath increases as the angle of rate of descent increases 

and, conversely, decreases as the angle of rate of descent 

decreases.

Stalls 

An aircraft stall results from a rapid decrease in lift caused by 

the separation of airflow from the wing’s surface brought on 

by exceeding the critical AOA. A stall can occur at any pitch 

attitude or airspeed. Stalls are one of the most misunderstood 

areas of aerodynamics because pilots often believe an airfoil 

stops producing lift when it stalls. In a stall, the wing does 

not totally stop producing lift. Rather, it can not generate 

adequate lift to sustain level flight.

Since the CL increases with an increase in AOA, at some 

point the CL peaks and then begins to drop off. This peak is 

called the CL-MAX. The amount of lift the wing produces drops 

dramatically after exceeding the CL-MAX or critical AOA, but 

as stated above, it does not completely stop producing lift. 

In most straight-wing aircraft, the wing is designed to stall 

the wing root first. The wing root reaches its critical AOA 

first making the stall progress outward toward the wingtip. 

By having the wing root stall first, aileron effectiveness is 

maintained at the wingtips, maintaining controllability of 

the aircraft. Various design methods are used to achieve 

the stalling of the wing root first. In one design, the wing is 

“twisted” to a higher AOA at the wing root. Installing stall 

strips on the first 20–25 percent of the wing’s leading edge 

is another method to introduce a stall prematurely. 

The wing never completely stops producing lift in a stalled 

condition. If it did, the aircraft would fall to the Earth. Most 

training aircraft are designed for the nose of the aircraft to 

drop during a stall, reducing the AOA and “unstalling” the 

wing. The “nose-down” tendency is due to the CL being aft 

of the CG. The CG range is very important when it comes 

to stall recovery characteristics. If an aircraft is allowed to 

be operated outside of the CG, the pilot may have difficulty 

recovering from a stall. The most critical CG violation would 

occur when operating with a CG which exceeds the rear limit. 

In this situation, a pilot may not be able to generate sufficient 

force with the elevator to counteract the excess weight aft of 

the CG. Without the ability to decrease the AOA, the aircraft 

continues in a stalled condition until it contacts the ground.

The stalling speed of a particular aircraft is not a fixed value 

for all flight situations, but a given aircraft always stalls at 

the same AOA regardless of airspeed, weight, load factor, 

or density altitude. Each aircraft has a particular AOA where 

the airflow separates from the upper surface of the wing and 

the stall occurs. This critical AOA varies from 16° to 20° 

depending on the aircraft’s design. But each aircraft has only 

one specific AOA where the stall occurs.

There are three flight situations in which the critical AOA 

can be exceeded: low speed, high speed, and turning.

The aircraft can be stalled in straight-and-level flight by flying 

too slowly. As the airspeed decreases, the AOA must be 

increased to retain the lift required for maintaining altitude. 

The lower the airspeed becomes, the more the AOA must 

be increased. Eventually, an AOA is reached which results 

in the wing not producing enough lift to support the aircraft 

which starts settling. If the airspeed is reduced further, the 

aircraft stalls, since the AOA has exceeded the critical angle 

and the airflow over the wing is disrupted.

Low speed is not necessary to produce a stall. The wing can 

be brought into an excessive AOA at any speed. For example, 

an aircraft is in a dive with an airspeed of 100 knots when the 

pilot pulls back sharply on the elevator control. [Figure 4-32]

Gravity and centrifugal force prevent an immediate alteration 
4-23

100

90

80

60

40

20

0

0° 10° 20° 30° 40° 50° 60° 70° 80° 90°

13

12

11

10

9

8

7

6

5

4

3

2

1

Percent increase in stall speed

 Load factor or “G”

 Bank Angle

Load factor

Stall speed increase

Figure 4-33. Increase in stall speed and load factor.

Figure 4-34. Inflight ice formation.

of the flightpath, but the aircraft’s AOA changes abruptly 

from quite low to very high. Since the flightpath of the aircraft 

in relation to the oncoming air determines the direction of the 

relative wind, the AOA is suddenly increased, and the aircraft 

would reach the stalling angle at a speed much greater than 

the normal stall speed.

The stalling speed of an aircraft is also higher in a level turn 

than in straight-and-level flight. [Figure 4-33] Centrifugal 

force is added to the aircraft’s weight and the wing must 

produce sufficient additional lift to counterbalance the load 

imposed by the combination of centrifugal force and weight. 

In a turn, the necessary additional lift is acquired by applying 

back pressure to the elevator control. This increases the wing’s 

AOA, and results in increased lift. The AOA must increase 

as the bank angle increases to counteract the increasing load 

caused by centrifugal force. If at any time during a turn the 

AOA becomes excessive, the aircraft stalls.

At this point, the action of the aircraft during a stall should 

be examined. To balance the aircraft aerodynamically, the CL 

is normally located aft of the CG. Although this makes the 

aircraft inherently nose-heavy, downwash on the horizontal 

stabilizer counteracts this condition. At the point of stall, when 

the upward force of the wing’s lift and the downward tail 

force cease, an unbalanced condition exists. This allows the 

aircraft to pitch down abruptly, rotating about its CG. During 

this nose-down attitude, the AOA decreases and the airspeed 

again increases. The smooth flow of air over the wing begins 

again, lift returns, and the aircraft is again flying. Considerable 

altitude may be lost before this cycle is complete.

Airfoil shape and degradation of that shape must also be 

considered in a discussion of stalls. For example, if ice, 

snow, and frost are allowed to accumulate on the surface of 

an aircraft, the smooth airflow over the wing is disrupted. 

This causes the boundary layer to separate at an AOA lower 

than that of the critical angle. Lift is greatly reduced, altering 

expected aircraft performance. If ice is allowed to accumulate 

on the aircraft during flight [Figure 4-34], the weight of 

the aircraft is increased while the ability to generate lift is 

decreased. As little as 0.8 millimeter of ice on the upper 

wing surface increases drag and reduces aircraft lift by 25 

percent.

Pilots can encounter icing in any season, anywhere in the 

country, at altitudes of up to 18,000 feet and sometimes 

higher. Small aircraft, including commuter planes, are most 

vulnerable because they fly at lower altitudes where ice is 

more prevalent. They also lack mechanisms common on jet 

aircraft that prevent ice buildup by heating the front edges 

of wings.

Icing can occur in clouds any time the temperature drops 

below freezing and super-cooled droplets build up on an 

aircraft and freeze. (Super-cooled droplets are still liquid 

even though the temperature is below 32 °Fahrenheit (F), 

or 0 °Celsius (C).

Basic Propeller Principles

The aircraft propeller consists of two or more blades and a 

central hub to which the blades are attached. Each blade of 

an aircraft propeller is essentially a rotating wing. As a result 

of their construction, the propeller blades are like airfoils 

and produce forces that create the thrust to pull, or push, 

the aircraft through the air. The engine furnishes the power 

needed to rotate the propeller blades through the air at high 

speeds, and the propeller transforms the rotary power of the 

engine into forward thrust.

A cross-section of a typical propeller blade is shown in 

Figure 4-35. This section or blade element is an airfoil 

comparable to a cross-section of an aircraft wing. One 

surface of the blade is cambered or curved, similar to the 

upper surface of an aircraft wing, while the other surface is 

flat like the bottom surface of a wing. The chord line is an 

imaginary line drawn through the blade from its leading edge 

to its trailing edge. As in a wing, the leading edge is the thick 

edge of the blade that meets the air as the propeller rotates.
4-24

Propeller blade angle

Relative wind

Chord line

Thrust

Forward velocity

Rotational velocity

blade angle

Pitch or

of attack

Angle

Figure 4-36. Propeller blade angle.

Airfoil sections of propeller blade

Figure 4-35. Airfoil sections of propeller blade.

Blade angle, usually measured in degrees, is the angle 

between the chord of the blade and the plane of rotation 

and is measured at a specific point along the length of the 

blade. [Figure 4-36] Because most propellers have a flat 

blade “face,” the chord line is often drawn along the face 

of the propeller blade. Pitch is not blade angle, but because 

pitch is largely determined by blade angle, the two terms are 

often used interchangeably. An increase or decrease in one is 

usually associated with an increase or decrease in the other.

The pitch of a propeller may be designated in inches. A 

propeller designated as a “74-48” would be 74 inches in 

length and have an effective pitch of 48 inches. The pitch 

is the distance in inches, which the propeller would screw 

through the air in one revolution if there were no slippage.

When specifying a fixed-pitch propeller for a new type of 

aircraft, the manufacturer usually selects one with a pitch 

that operates efficiently at the expected cruising speed of the 

aircraft. Every fixed-pitch propeller must be a compromise 

because it can be efficient at only a given combination of 

airspeed and revolutions per minute (rpm). Pilots cannot 

change this combination in flight.

When the aircraft is at rest on the ground with the engine 

operating, or moving slowly at the beginning of takeoff, 

the propeller efficiency is very low because the propeller is 

restrained from advancing with sufficient speed to permit 

its fixed-pitch blades to reach their full efficiency. In this 

situation, each propeller blade is turning through the air at 

an AOA that produces relatively little thrust for the amount 

of power required to turn it.

To understand the action of a propeller, consider first its 

motion, which is both rotational and forward. As shown by 

the vectors of propeller forces in Figure 4-36, each section of 

a propeller blade moves downward and forward. The angle 

at which this air (relative wind) strikes the propeller blade is 

its AOA. The air deflection produced by this angle causes the 

dynamic pressure at the engine side of the propeller blade to 

be greater than atmospheric pressure, thus creating thrust.

The shape of the blade also creates thrust because it is 

cambered like the airfoil shape of a wing. As the air flows 

past the propeller, the pressure on one side is less than that 

on the other. As in a wing, a reaction force is produced in the 

direction of the lesser pressure. The airflow over the wing 

has less pressure, and the force (lift) is upward. In the case 

of the propeller, which is mounted in a vertical instead of a 

horizontal plane, the area of decreased pressure is in front of 

the propeller, and the force (thrust) is in a forward direction. 

Aerodynamically, thrust is the result of the propeller shape 

and the AOA of the blade.

Thrust can be considered also in terms of the mass of air 

handled by the propeller. In these terms, thrust equals mass 

of air handled multiplied by slipstream velocity minus 

velocity of the aircraft. The power expended in producing 

thrust depends on the rate of air mass movement. On average, 

thrust constitutes approximately 80 percent of the torque 

(total horsepower absorbed by the propeller). The other 

20 percent is lost in friction and slippage. For any speed of 

rotation, the horsepower absorbed by the propeller balances 

the horsepower delivered by the engine. For any single 

revolution of the propeller, the amount of air handled depends 

on the blade angle, which determines how big a “bite” of 

air the propeller takes. Thus, the blade angle is an excellent 

means of adjusting the load on the propeller to control the 

engine rpm.

The blade angle is also an excellent method of adjusting the 

AOA of the propeller. On constant-speed propellers, the blade 

angle must be adjusted to provide the most efficient AOA at 

all engine and aircraft speeds. Lift versus drag curves, which 

are drawn for propellers, as well as wings, indicate that the 

most efficient AOA is small, varying from +2° to +4°. The 

actual blade angle necessary to maintain this small AOA 

varies with the forward speed of the aircraft.
4-25

Propeller tips travel faster than hubs

Effective Pitch

Geometric Pitch

Slip

Figure 4-37. Propeller slippage.

Propeller travels faster at the tips

60 in.

40 in.

20 in.

Short travel distance—slow speed

—129 kno st

Moderate travel distance—moderate speed—259 knots

Greater travel distance—very high speed—389 knots

2,500 rpm

2,500 rpm

2,500 rpm

Figure 4-38. Propeller tips travel faster than the hub.

Fixed-pitch and ground-adjustable propellers are designed 

for best efficiency at one rotation and forward speed. They 

are designed for a given aircraft and engine combination. A 

propeller may be used that provides the maximum efficiency 

for takeoff, climb, cruise, or high-speed flight. Any change in 

these conditions results in lowering the efficiency of both the 

propeller and the engine. Since the efficiency of any machine 

is the ratio of the useful power output to the actual power 

input, propeller efficiency is the ratio of thrust horsepower to 

brake horsepower. Propeller efficiency varies from 50 to 87 

percent, depending on how much the propeller “slips.”

Propeller slip is the difference between the geometric pitch of 

the propeller and its effective pitch. [Figure 4-37] Geometric 

pitch is the theoretical distance a propeller should advance 

in one revolution; effective pitch is the distance it actually 

advances. Thus, geometric or theoretical pitch is based on 

no slippage, but actual or effective pitch includes propeller 

slippage in the air.

The reason a propeller is “twisted” is that the outer parts of the 

propeller blades, like all things that turn about a central point, 

travel faster than the portions near the hub. [Figure 4-38] If 

the blades had the same geometric pitch throughout their 

lengths, portions near the hub could have negative AOAs 

while the propeller tips would be stalled at cruise speed. 

Twisting or variations in the geometric pitch of the blades 

permits the propeller to operate with a relatively constant 

AOA along its length when in cruising flight. Propeller 

blades are twisted to change the blade angle in proportion 

to the differences in speed of rotation along the length of 

the propeller, keeping thrust more nearly equalized along 

this length.

Usually 1° to 4° provides the most efficient lift/drag ratio, 

but in flight the propeller AOA of a fixed-pitch propeller 

varies—normally from 0° to 15°. This variation is caused 

by changes in the relative airstream, which in turn results 

from changes in aircraft speed. Thus, propeller AOA is the 

product of two motions: propeller rotation about its axis and 

its forward motion.

A constant-speed propeller automatically keeps the blade 

angle adjusted for maximum efficiency for most conditions 

encountered in flight. During takeoff, when maximum power 

and thrust are required, the constant-speed propeller is at 

a low propeller blade angle or pitch. The low blade angle 

keeps the AOA small and efficient with respect to the relative 

wind. At the same time, it allows the propeller to handle a 

smaller mass of air per revolution. This light load allows 

the engine to turn at high rpm and to convert the maximum 

amount of fuel into heat energy in a given time. The high rpm 

also creates maximum thrust because, although the mass of 

air handled per revolution is small, the rpm and slipstream 

velocity are high, and with the low aircraft speed, there is 

maximum thrust.

After liftoff, as the speed of the aircraft increases, the constantspeed propeller automatically changes to a higher angle (or 

pitch). Again, the higher blade angle keeps the AOA small 

and efficient with respect to the relative wind. The higher 

blade angle increases the mass of air handled per revolution. 

This decreases the engine rpm, reducing fuel consumption 

and engine wear, and keeps thrust at a maximum.

After the takeoff climb is established in an aircraft having 

a controllable-pitch propeller, the pilot reduces the power 

output of the engine to climb power by first decreasing the 

manifold pressure and then increasing the blade angle to 

lower the rpm.
4-26

Torque reaction

Reaction

Action

Figure 4-39. Torque reaction.

Yaw

force

slipstream

Figure 4-40. Corkscrewing slipstream.

At cruising altitude, when the aircraft is in level flight and 

less power is required than is used in takeoff or climb, the 

pilot again reduces engine power by reducing the manifold 

pressure and then increasing the blade angle to decrease the 

rpm. Again, this provides a torque requirement to match the 

reduced engine power. Although the mass of air handled per 

revolution is greater, it is more than offset by a decrease in 

slipstream velocity and an increase in airspeed. The AOA is 

still small because the blade angle has been increased with 

an increase in airspeed.

Torque and P-Factor

To the pilot, “torque” (the left turning tendency of the 

airplane) is made up of four elements which cause or produce 

a twisting or rotating motion around at least one of the 

airplane’s three axes. These four elements are:

1. Torque reaction from engine and propeller,

2. Corkscrewing effect of the slipstream,

3. Gyroscopic action of the propeller, and

4. Asymmetric loading of the propeller (P-factor).

Torque Reaction

Torque reaction involves Newton’s Third Law of Physics—

for every action, there is an equal and opposite reaction. As 

applied to the aircraft, this means that as the internal engine 

parts and propeller are revolving in one direction, an equal 

force is trying to rotate the aircraft in the opposite direction. 

[Figure 4-39] 

When the aircraft is airborne, this force is acting around 

the longitudinal axis, tending to make the aircraft roll. To 

compensate for roll tendency, some of the older aircraft are 

rigged in a manner to create more lift on the wing that is being 

forced downward. The more modern aircraft are designed 

with the engine offset to counteract this effect of torque.

NOTE: Most United States built aircraft engines rotate the 

propeller clockwise, as viewed from the pilot’s seat. The 

discussion here is with reference to those engines.

Generally, the compensating factors are permanently set 

so that they compensate for this force at cruising speed, 

since most of the aircraft’s operating lift is at that speed. 

However, aileron trim tabs permit further adjustment for 

other speeds.

When the aircraft’s wheels are on the ground during the 

takeoff roll, an additional turning moment around the vertical 

axis is induced by torque reaction. As the left side of the 

aircraft is being forced down by torque reaction, more weight 

is being placed on the left main landing gear. This results in 

more ground friction, or drag, on the left tire than on the right, 

causing a further turning moment to the left. The magnitude 

of this moment is dependent on many variables. Some of 

these variables are: 

1. Size and horsepower of engine,

2. Size of propeller and the rpm,

3. Size of the aircraft, and

4. Condition of the ground surface.

This yawing moment on the takeoff roll is corrected by the 

pilot’s proper use of the rudder or rudder trim.

Corkscrew Effect

The high-speed rotation of an aircraft propeller gives a 

corkscrew or spiraling rotation to the slipstream. At high 

propeller speeds and low forward speed (as in the takeoffs 

and approaches to power-on stalls), this spiraling rotation 

is very compact and exerts a strong sideward force on the 

aircraft’s vertical tail surface. [Figure 4-40] 

When this spiraling slipstream strikes the vertical fin it causes 

a turning moment about the aircraft’s vertical axis. The more 

compact the spiral, the more prominent this force is. As the 

forward speed increases, however, the spiral elongates and 

becomes less effective.The corkscrew flow of the slipstream 

also causes a rolling moment around the longitudinal axis. 
4-27

Resultant force 90° Effective

force

Applied force

Yaw

Figure 4-41. Gyroscopic precession.

Applied force

Resultant

force

Yaw

Effective

force

Figure 4-42. Raising tail produces gyroscopic precession.

Note that this rolling moment caused by the corkscrew flow 

of the slipstream is to the right, while the rolling moment 

caused by torque reaction is to the left—in effect one may 

be counteracting the other. However, these forces vary 

greatly and it is the pilot’s responsibility to apply proper 

corrective action by use of the flight controls at all times. 

These forces must be counteracted regardless of which is 

the most prominent at the time.

Gyroscopic Action

Before the gyroscopic effects of the propeller can be 

understood, it is necessary to understand the basic principle 

of a gyroscope. All practical applications of the gyroscope 

are based upon two fundamental properties of gyroscopic 

action: rigidity in space and precession. The one of interest 

for this discussion is precession.

Precession is the resultant action, or deflection, of a spinning 

rotor when a deflecting force is applied to its rim. As can be 

seen in Figure 4-41, when a force is applied, the resulting 

force takes effect 90° ahead of and in the direction of 

rotation.

The rotating propeller of an airplane makes a very good 

gyroscope and thus has similar properties. Any time a force 

is applied to deflect the propeller out of its plane of rotation, 

the resulting force is 90° ahead of and in the direction of 

rotation and in the direction of application, causing a pitching 

moment, a yawing moment, or a combination of the two 

depending upon the point at which the force was applied.

This element of torque effect has always been associated with 

and considered more prominent in tailwheel-type aircraft, 

and most often occurs when the tail is being raised during 

the takeoff roll. [Figure 4-42] This change in pitch attitude 

has the same effect as applying a force to the top of the 

propeller’s plane of rotation. The resultant force acting 90° 

ahead causes a yawing moment to the left around the vertical 

axis. The magnitude of this moment depends on several 

variables, one of which is the abruptness with which the tail 

is raised (amount of force applied). However, precession, 

or gyroscopic action, occurs when a force is applied to any 

point on the rim of the propeller’s plane of rotation; the 

resultant force will still be 90° from the point of application 

in the direction of rotation. Depending on where the force is 

applied, the airplane is caused to yaw left or right, to pitch 

up or down, or a combination of pitching and yawing.

It can be said that, as a result of gyroscopic action, any yawing 

around the vertical axis results in a pitching moment, and any 

pitching around the lateral axis results in a yawing moment. 

To correct for the effect of gyroscopic action, it is necessary 

for the pilot to properly use elevator and rudder to prevent 

undesired pitching and yawing.

Asymmetric Loading (P-Factor)

When an aircraft is flying with a high AOA, the “bite” of 

the downward moving blade is greater than the “bite” of the 

upward moving blade. This moves the center of thrust to the 

right of the prop disc area, causing a yawing moment toward 

the left around the vertical axis. To prove this explanation is 

complex because it would be necessary to work wind vector 

problems on each blade while considering both the AOA of 

the aircraft and the AOA of each blade.

This asymmetric loading is caused by the resultant velocity, 

which is generated by the combination of the velocity of the 

propeller blade in its plane of rotation and the velocity of the 

air passing horizontally through the propeller disc. With the 

aircraft being flown at positive AOAs, the right (viewed from 

the rear) or downswinging blade, is passing through an area of 

resultant velocity which is greater than that affecting the left 

or upswinging blade. Since the propeller blade is an airfoil, 

increased velocity means increased lift. The downswinging 

blade has more lift and tends to pull (yaw) the aircraft’s nose 

to the left.

When the aircraft is flying at a high AOA, the downward 

moving blade has a higher resultant velocity, creating more 

lift than the upward moving blade. [Figure 4-43] This might 

be easier to visualize if the propeller shaft was mounted 

perpendicular to the ground (like a helicopter). If there 

were no air movement at all, except that generated by the 
4-28

Low angle of attack High angle of attack

Load on

upward moving

propeller blade

Load on

downward moving

propeller blade

Load on

downward moving

propeller blade

Load on

upward moving

propeller blade

Figure 4-43. Asymmetrical loading of propeller (P-factor).

propeller itself, identical sections of each blade would have 

the same airspeed. With air moving horizontally across this 

vertically mounted propeller, the blade proceeding forward 

into the flow of air has a higher airspeed than the blade 

retreating with the airflow. Thus, the blade proceeding into 

the horizontal airflow is creating more lift, or thrust, moving 

the center of thrust toward that blade. Visualize rotating the 

vertically mounted propeller shaft to shallower angles relative 

to the moving air (as on an aircraft). This unbalanced thrust 

then becomes proportionately smaller and continues getting 

smaller until it reaches the value of zero when the propeller 

shaft is exactly horizontal in relation to the moving air.

The effects of each of these four elements of torque vary in 

value with changes in flight situations. In one phase of flight, 

one of these elements may be more prominent than another. 

In another phase of flight, another element may be more 

prominent. The relationship of these values to each other 

varies with different aircraft—depending on the airframe, 

engine, and propeller combinations, as well as other design 

features. To maintain positive control of the aircraft in all 

flight conditions, the pilot must apply the flight controls as 

necessary to compensate for these varying values.

Load Factors

In aerodynamics, load factor is the ratio of the maximum load 

an aircraft can sustain to the gross weight of the aircraft. The 

load factor is measured in Gs (acceleration of gravity), a unit 

of force equal to the force exerted by gravity on a body at rest 

and indicates the force to which a body is subjected when it 

is accelerated. Any force applied to an aircraft to deflect its 

flight from a straight line produces a stress on its structure, 

and the amount of this force is the load factor. While a course 

in aerodynamics is not a prerequisite for obtaining a pilot’s 

license, the competent pilot should have a solid understanding 

of the forces that act on the aircraft, the advantageous use 

of these forces, and the operating limitations of the aircraft 

being flown. 

For example, a load factor of 3 means the total load on an 

aircraft’s structure is three times its gross weight. Since load 

factors are expressed in terms of Gs, a load factor of 3 may 

be spoken of as 3 Gs, or a load factor of 4 as 4 Gs.

If an aircraft is pulled up from a dive, subjecting the pilot to 

3 Gs, he or she would be pressed down into the seat with a 

force equal to three times his or her weight. Since modern 

aircraft operate at significantly higher speeds than older 

aircraft, increasing the magnitude of the load factor, this 

effect has become a primary consideration in the design of 

the structure of all aircraft.

With the structural design of aircraft planned to withstand 

only a certain amount of overload, a knowledge of load 

factors has become essential for all pilots. Load factors are 

important for two reasons:

1. It is possible for a pilot to impose a dangerous overload 

on the aircraft structures.

2. An increased load factor increases the stalling speed 

and makes stalls possible at seemingly safe flight 

speeds.

Load Factors in Aircraft Design

The answer to the question “How strong should an aircraft 

be?” is determined largely by the use to which the aircraft is 

subjected. This is a difficult problem because the maximum 

possible loads are much too high for use in efficient design. 

It is true that any pilot can make a very hard landing or an 

extremely sharp pull up from a dive, which would result in 

abnormal loads. However, such extremely abnormal loads 

must be dismissed somewhat if aircraft are built that take off 

quickly, land slowly, and carry worthwhile payloads.

The problem of load factors in aircraft design becomes how 

to determine the highest load factors that can be expected in 

normal operation under various operational situations. These 

load factors are called “limit load factors.” For reasons of 

safety, it is required that the aircraft be designed to withstand 

these load factors without any structural damage. Although 

the Code of Federal Regulations (CFR) requires the aircraft 

structure be capable of supporting one and one-half times 

these limit load factors without failure, it is accepted that 

parts of the aircraft may bend or twist under these loads and 

that some structural damage may occur.

This 1.5 load limit factor is called the “factor of safety” and 

provides, to some extent, for loads higher than those expected 

under normal and reasonable operation. This strength reserve 

is not something which pilots should willfully abuse; rather, 

it is there for protection when encountering unexpected 

conditions.
4-29

Two forces cause load factor during turns

Load factor = 2 Gs

Centrifugal

force = 1.73 Gs

Gravity = 1G

60° 50° 40° 30° 20° 10°

Figure 4-44. Two forces cause load factor during turns.

7

6

5

4

3

2

1

0

0° 10° 20° 30° 40° 50° 60° 70° 80° 90°

Load factor (G units)

 Bank angle

Figure 4-45. Angle of bank changes load factor.

The above considerations apply to all loading conditions, 

whether they be due to gusts, maneuvers, or landings. The 

gust load factor requirements now in effect are substantially 

the same as those that have been in existence for years. 

Hundreds of thousands of operational hours have proven 

them adequate for safety. Since the pilot has little control over 

gust load factors (except to reduce the aircraft’s speed when 

rough air is encountered), the gust loading requirements are 

substantially the same for most general aviation type aircraft 

regardless of their operational use. Generally, the gust load 

factors control the design of aircraft which are intended for 

strictly nonacrobatic usage.

An entirely different situation exists in aircraft design with 

maneuvering load factors. It is necessary to discuss this matter 

separately with respect to: (1) aircraft designed in accordance 

with the category system (i.e., normal, utility, acrobatic); and 

(2) older designs built according to requirements which did 

not provide for operational categories.

Aircraft designed under the category system are readily 

identified by a placard in the flight deck, which states the 

operational category (or categories) in which the aircraft 

is certificated. The maximum safe load factors (limit load 

factors) specified for aircraft in the various categories are:

CATEGORY LIMIT LOAD FACTOR

Normal1 3.8 to –1.52

Utility (mild acrobatics, including spins) 4.4 to –1.76

Acrobatic 6.0 to –3.00

1 For aircraft with gross weight of more than 4,000 pounds, 

the limit load factor is reduced. To the limit loads given above, 

a safety factor of 50 percent is added.

There is an upward graduation in load factor with the 

increasing severity of maneuvers. The category system 

provides for maximum utility of an aircraft. If normal 

operation alone is intended, the required load factor (and 

consequently the weight of the aircraft) is less than if the 

aircraft is to be employed in training or acrobatic maneuvers 

as they result in higher maneuvering loads.

Aircraft that do not have the category placard are designs that 

were constructed under earlier engineering requirements in 

which no operational restrictions were specifically given to 

the pilots. For aircraft of this type (up to weights of about 

4,000 pounds), the required strength is comparable to presentday utility category aircraft, and the same types of operation 

are permissible. For aircraft of this type over 4,000 pounds, 

the load factors decrease with weight. These aircraft should 

be regarded as being comparable to the normal category 

aircraft designed under the category system, and they should 

be operated accordingly.

Load Factors in Steep Turns

In a constant altitude, coordinated turn in any aircraft, the 

load factor is the result of two forces: centrifugal force and 

gravity. [Figure 4-44] For any given bank angle, the ROT 

varies with the airspeed—the higher the speed, the slower the 

ROT. This compensates for added centrifugal force, allowing 

the load factor to remain the same.

Figure 4-45 reveals an important fact about turns—the load 

factor increases at a terrific rate after a bank has reached 

45° or 50°. The load factor for any aircraft in a 60° bank is 

2 Gs. The load factor in an 80° bank is 5.76 Gs. The wing 

must produce lift equal to these load factors if altitude is to 

be maintained.

It should be noted how rapidly the line denoting load factor 

rises as it approaches the 90° bank line, which it never quite 

reaches because a 90° banked, constant altitude turn is not 
4-30

3

2

5

1

4

0

0 1 2 3 4 5 6 7 8 20 40 60 80 100 120 140 160 180 200 220 240 260

Ratio of acceleration 

Vs to unaccelerated 

Vs

Unaccelerated stall speed

 “G” Load Accelerated stall speed

40

50

60

70

80

100

120

150

40

50

60

70

80

100

120

150

Figure 4-46. Load factor changes stall speed.

mathematically possible. An aircraft may be banked to 90°, 

but not in a coordinated turn. An aircraft which can be held in 

a 90° banked slipping turn is capable of straight knife-edged 

flight. At slightly more than 80°, the load factor exceeds the 

limit of 6 Gs, the limit load factor of an acrobatic aircraft.

For a coordinated, constant altitude turn, the approximate 

maximum bank for the average general aviation aircraft is 60°. 

This bank and its resultant necessary power setting reach the 

limit of this type of aircraft. An additional 10° bank increases 

the load factor by approximately 1 G, bringing it close to the 

yield point established for these aircraft. [Figure 4-46]

Load Factors and Stalling Speeds

Any aircraft, within the limits of its structure, may be stalled 

at any airspeed. When a sufficiently high AOA is imposed, 

the smooth flow of air over an airfoil breaks up and separates, 

producing an abrupt change of flight characteristics and a 

sudden loss of lift, which results in a stall.

A study of this effect has revealed that the aircraft’s 

stalling speed increases in proportion to the square root of 

the load factor. This means that an aircraft with a normal 

unaccelerated stalling speed of 50 knots can be stalled at 100 

knots by inducing a load factor of 4 Gs. If it were possible 

for this aircraft to withstand a load factor of nine, it could be 

stalled at a speed of 150 knots. A pilot should be aware:

• Of the danger of inadvertently stalling the aircraft by 

increasing the load factor, as in a steep turn or spiral; 

• When intentionally stalling an aircraft above its 

design maneuvering speed, a tremendous load factor 

is imposed.

Figures 4-45 and 4-46 show that banking an aircraft greater 

than 72° in a steep turn produces a load factor of 3, and the 

stalling speed is increased significantly. If this turn is made 

in an aircraft with a normal unaccelerated stalling speed of 

45 knots, the airspeed must be kept greater than 75 knots to 

prevent inducing a stall. A similar effect is experienced in a 

quick pull up, or any maneuver producing load factors above 

1 G. This sudden, unexpected loss of control, particularly in 

a steep turn or abrupt application of the back elevator control 

near the ground, has caused many accidents.

Since the load factor is squared as the stalling speed doubles, 

tremendous loads may be imposed on structures by stalling 

an aircraft at relatively high airspeeds.

The maximum speed at which an aircraft may be stalled safely 

is now determined for all new designs. This speed is called 

the “design maneuvering speed” (VA) and must be entered in 

the FAA-approved Airplane Flight Manual/Pilot’s Operating 

Handbook (AFM/POH) of all recently designed aircraft. For 

older general aviation aircraft, this speed is approximately 1.7 

times the normal stalling speed. Thus, an older aircraft which 

normally stalls at 60 knots must never be stalled at above 102 

knots (60 knots x 1.7 = 102 knots). An aircraft with a normal 

stalling speed of 60 knots stalled at 102 knots undergoes a load 

factor equal to the square of the increase in speed, or 2.89 Gs 

(1.7 x 1.7 = 2.89 Gs). (The above figures are approximations 

to be considered as a guide, and are not the exact answers to 

any set of problems. The design maneuvering speed should be 

determined from the particular aircraft’s operating limitations 

provided by the manufacturer.)
4-31

Since the leverage in the control system varies with different 

aircraft (some types employ “balanced” control surfaces while 

others do not), the pressure exerted by the pilot on the controls 

cannot be accepted as an index of the load factors produced 

in different aircraft. In most cases, load factors can be judged 

by the experienced pilot from the feel of seat pressure. Load 

factors can also be measured by an instrument called an 

“accelerometer,” but this instrument is not common in general 

aviation training aircraft. The development of the ability to 

judge load factors from the feel of their effect on the body is 

important. A knowledge of these principles is essential to the 

development of the ability to estimate load factors.

A thorough knowledge of load factors induced by varying 

degrees of bank and the VA aids in the prevention of two of 

the most serious types of accidents:

1. Stalls from steep turns or excessive maneuvering near 

the ground

2. Structural failures during acrobatics or other violent 

maneuvers resulting from loss of control

Load Factors and Flight Maneuvers

Critical load factors apply to all flight maneuvers except 

unaccelerated straight flight where a load factor of 1 G is 

always present. Certain maneuvers considered in this section 

are known to involve relatively high load factors.

Turns

Increased load factors are a characteristic of all banked 

turns. As noted in the section on load factors in steep turns, 

load factors become significant to both flight performance 

and load on wing structure as the bank increases beyond 

approximately 45°.

The yield factor of the average light plane is reached 

at a bank of approximately 70° to 75°, and the stalling 

speed is increased by approximately one-half at a bank of 

approximately 63°.

Stalls

The normal stall entered from straight-and-level flight, or an 

unaccelerated straight climb, does not produce added load 

factors beyond the 1 G of straight-and-level flight. As the stall 

occurs, however, this load factor may be reduced toward zero, 

the factor at which nothing seems to have weight. The pilot 

experiences a sensation of “floating free in space.” If recovery 

is effected by snapping the elevator control forward, negative 

load factors (or those that impose a down load on the wings 

and raise the pilot from the seat) may be produced.

During the pull up following stall recovery, significant load 

factors are sometimes induced. These may be further increased 

inadvertently during excessive diving (and consequently high 

airspeed) and abrupt pull ups to level flight. One usually leads 

to the other, thus increasing the load factor. Abrupt pull ups 

at high diving speeds may impose critical loads on aircraft 

structures and may produce recurrent or secondary stalls by 

increasing the AOA to that of stalling.

As a generalization, a recovery from a stall made by diving 

only to cruising or design maneuvering airspeed, with a 

gradual pull up as soon as the airspeed is safely above stalling, 

can be effected with a load factor not to exceed 2 or 2.5 Gs. A 

higher load factor should never be necessary unless recovery 

has been effected with the aircraft’s nose near or beyond the 

vertical attitude, or at extremely low altitudes to avoid diving 

into the ground.

Spins

A stabilized spin is not different from a stall in any element 

other than rotation and the same load factor considerations 

apply to spin recovery as apply to stall recovery. Since spin 

recoveries are usually effected with the nose much lower 

than is common in stall recoveries, higher airspeeds and 

consequently higher load factors are to be expected. The 

load factor in a proper spin recovery usually is found to be 

about 2.5 Gs.

The load factor during a spin varies with the spin characteristics 

of each aircraft, but is usually found to be slightly above the 

1 G of level flight. There are two reasons for this:

1. Airspeed in a spin is very low, usually within 2 knots 

of the unaccelerated stalling speeds.

2. Aircraft pivots, rather than turns, while it is in a 

spin.

High Speed Stalls

The average light plane is not built to withstand the repeated 

application of load factors common to high speed stalls. 

The load factor necessary for these maneuvers produces a 

stress on the wings and tail structure, which does not leave 

a reasonable margin of safety in most light aircraft.

The only way this stall can be induced at an airspeed above 

normal stalling involves the imposition of an added load 

factor, which may be accomplished by a severe pull on the 

elevator control. A speed of 1.7 times stalling speed (about 

102 knots in a light aircraft with a stalling speed of 60 knots) 

produces a load factor of 3 Gs. Only a very narrow margin 

for error can be allowed for acrobatics in light aircraft. To 

illustrate how rapidly the load factor increases with airspeed, 

a high-speed stall at 112 knots in the same aircraft would 

produce a load factor of 4 Gs.
4-32

Chandelles and Lazy Eights

A chandelle is a maximum performance climbing turn 

beginning from approximately straight-and-level flight, 

and ending at the completion of a precise 180° of turn in a 

wings-level, nose-high attitude at the minimum controllable 

airspeed. In this flight maneuver, the aircraft is in a steep 

climbing turn and almost stalls to gain altitude while changing 

direction. A lazy eight derives its name from the manner in 

which the extended longitudinal axis of the aircraft is made 

to trace a flight pattern in the form of a figure “8” lying on 

its side. It would be difficult to make a definite statement 

concerning load factors in these maneuvers as both involve 

smooth, shallow dives and pull ups. The load factors incurred 

depend directly on the speed of the dives and the abruptness 

of the pull ups during these maneuvers.

Generally, the better the maneuver is performed, the less 

extreme the load factor induced. A chandelle or lazy eight in 

which the pull-up produces a load factor greater than 2 Gs will 

not result in as great a gain in altitude, and in low-powered 

aircraft it may result in a net loss of altitude.

The smoothest pull up possible, with a moderate load factor, 

delivers the greatest gain in altitude in a chandelle and results 

in a better overall performance in both chandelles and lazy 

eights. The recommended entry speed for these maneuvers 

is generally near the manufacturer’s design maneuvering 

speed which allows maximum development of load factors 

without exceeding the load limits.

Rough Air

All standard certificated aircraft are designed to withstand 

loads imposed by gusts of considerable intensity. Gust load 

factors increase with increasing airspeed, and the strength 

used for design purposes usually corresponds to the highest 

level flight speed. In extremely rough air, as in thunderstorms 

or frontal conditions, it is wise to reduce the speed to the 

design maneuvering speed. Regardless of the speed held, 

there may be gusts that can produce loads which exceed the 

load limits.

Each specific aircraft is designed with a specific G loading 

that can be imposed on the aircraft without causing structural 

damage. There are two types of load factors factored into 

aircraft design, limit load and ultimate load. The limit load 

is a force applied to an aircraft that causes a bending of the 

aircraft structure that does not return to the original shape. 

The ultimate load is the load factor applied to the aircraft 

beyond the limit load and at which point the aircraft material 

experiences structural failure (breakage). Load factors lower 

than the limit load can be sustained without compromising 

the integrity of the aircraft structure.

Speeds up to but not exceeding the maneuvering speed allows 

an aircraft to stall prior to experiencing an increase in load 

factor that would exceed the limit load of the aircraft.

Most AFM/POH now include turbulent air penetration 

information, which help today’s pilots safely fly aircraft 

capable of a wide range of speeds and altitudes. It is important 

for the pilot to remember that the maximum “never-exceed” 

placard dive speeds are determined for smooth air only. High 

speed dives or acrobatics involving speed above the known 

maneuvering speed should never be practiced in rough or 

turbulent air.

Vg Diagram

The flight operating strength of an aircraft is presented 

on a graph whose vertical scale is based on load factor. 

[Figure 4-47] The diagram is called a Vg diagram—velocity 

versus G loads or load factor. Each aircraft has its own Vg 

diagram which is valid at a certain weight and altitude.

The lines of maximum lift capability (curved lines) are the 

first items of importance on the Vg diagram. The aircraft in 

the Figure 4-47 is capable of developing no more than +1 G 

at 62 mph, the wing level stall speed of the aircraft. Since the 

maximum load factor varies with the square of the airspeed, 

the maximum positive lift capability of this aircraft is 2 G at 

92 mph, 3 G at 112 mph, 4.4 G at 137 mph, and so forth. Any 

load factor above this line is unavailable aerodynamically 

(i.e., the aircraft cannot fly above the line of maximum lift 

capability because it stalls). The same situation exists for 

negative lift flight with the exception that the speed necessary 

to produce a given negative load factor is higher than that to 

produce the same positive load factor.

If the aircraft is flown at a positive load factor greater than 

the positive limit load factor of 4.4, structural damage 

is possible. When the aircraft is operated in this region, 

objectionable permanent deformation of the primary structure 

may take place and a high rate of fatigue damage is incurred. 

Operation above the limit load factor must be avoided in 

normal operation.

There are two other points of importance on the Vg diagram. 

One point is the intersection of the positive limit load factor 

and the line of maximum positive lift capability. The airspeed 

at this point is the minimum airspeed at which the limit load 

can be developed aerodynamically. Any airspeed greater than 

this provides a positive lift capability sufficient to damage 

the aircraft. Conversely, any airspeed less than this does not 

provide positive lift capability sufficient to cause damage 

from excessive flight loads. The usual term given to this speed 

is “maneuvering speed,” since consideration of subsonic 
4-33

0

–1

2

–2

1

–3

6

5

4

7

3

20 40 60 80 100 120 140 160 180 200 220 240

–

–

–

Load factor

Never exceed speed

Indicated airspeed (mph)

Structural damage

Caution range

Structural failure

Normal operating range

Accelerated stall

Structural Damage

Normal stall speed

Maneuvering speed

Level flight at 1 G

Figure 4-47. Typical Vg diagram.

aerodynamics would predict minimum usable turn radius or 

maneuverability to occur at this condition. The maneuver 

speed is a valuable reference point, since an aircraft operating 

below this point cannot produce a damaging positive flight 

load. Any combination of maneuver and gust cannot create 

damage due to excess airload when the aircraft is below the 

maneuver speed.

The other point of importance on the Vg diagram is the 

intersection of the negative limit load factor and line of 

maximum negative lift capability. Any airspeed greater than 

this provides a negative lift capability sufficient to damage 

the aircraft; any airspeed less than this does not provide 

negative lift capability sufficient to damage the aircraft from 

excessive flight loads.

The limit airspeed (or redline speed) is a design reference 

point for the aircraft—this aircraft is limited to 225 mph. 

If flight is attempted beyond the limit airspeed, structural 

damage or structural failure may result from a variety of 

phenomena.

The aircraft in flight is limited to a regime of airspeeds 

and Gs which do not exceed the limit (or redline) speed, 

do not exceed the limit load factor, and cannot exceed the 

maximum lift capability. The aircraft must be operated 

within this “envelope” to prevent structural damage and 

ensure the anticipated service lift of the aircraft is obtained. 

The pilot must appreciate the Vg diagram as describing the 

allowable combination of airspeeds and load factors for 

safe operation. Any maneuver, gust, or gust plus maneuver 

outside the structural envelope can cause structural damage 

and effectively shorten the service life of the aircraft.

Rate of Turn

The rate of turn (ROT) is the number of degrees (expressed in 

degrees per second) of heading change that an aircraft makes. 

The ROT can be determined by taking the constant of 1,091, 

multiplying it by the tangent of any bank angle and dividing 

that product by a given airspeed in knots as illustrated in 

Figure 4-48. If the airspeed is increased and the ROT desired 

is to be constant, the angle of bank must be increased, 

otherwise, the ROT decreases. Likewise, if the airspeed is 

held constant, an aircraft’s ROT increases if the bank angle 

is increased. The formula in Figures 4-48 through 4-50

depicts the relationship between bank angle and airspeed as 

they affect the ROT. 

NOTE: All airspeed discussed in this section is true airspeed 

(TAS).
4-34

Rate of turn for a given airspeed (knots, TAS)

and bank angle 

 1,091 x tangent of the bank angle

airspeed (in knots) ROT =

 1,091 x tangent of 30°

120 knots

ROT =

 1,091 x 0.5773 (tangent of 30°)

120 knots

ROT =

ROT = 5.25 degrees per second

Example The rate of turn for an aircraft in a 

coordinated turn of 30° and traveling at 

120 knots would have a ROT as follows.

Figure 4-48. Rate of turn for a given airspeed (knots, TAS) and 

bank angle. Increase in the speed 

 1,091 x tangent of 30°

240 knots

ROT =

ROT = 2.62 degrees per second

An increase in speed causes a decrease in 

the rate of turn when using the same bank angle.

Example Suppose we were to increase the speed to 240 

knots, what is the rate of turn? Using the same

formula from above we see that:

Figure 4-49. Rate of turn when increasing speed.

To maintain the same Rate of Turn of an aircraft 

traveling at 125 knots (approximately 5.25° per 

second using a 30° bank) but using an airspeed of

240 knots requires an increased bank angle.

 1,091 x tangent of X

240 knots ROT (5.25) =

240 x 5.25 = 1,091 x tangent of X

240 x 5.25 = tangent of X 1,091

1.1549 = tangent of X

49° = X

Example Suppose we wanted to know what bank angle 

would give us a rate of turn of 5.25° per second 

at 240 knots. A slight rearrangement of the formula 

would indicate it will take a 49° angle of bank to 

achieve the same ROT used at the lower airspeed 

of 120 knots.

Figure 4-50. To achieve the same rate of turn of an aircraft traveling 

at 120 knots, an increase of bank angle is required.

120 knots

11.26 x tangent of bank angle R = 

 1202

11.26 x tangent of 30°

R =

V2

R = 11.26 x 0.5773

14,400

R = 2,215 feet

The radius of a turn required by an aircraft traveling at 120 knots 

and using a bank angle of 30° is 2,215 feet.

Figure 4-51. Radius at 120 knots with bank angle of 30°.

Airspeed significantly effects an aircraft’s ROT. If airspeed is 

increased, the ROT is reduced if using the same angle of bank 

used at the lower speed. Therefore, if airspeed is increased 

as illustrated in Figure 4-49, it can be inferred that the angle 

of bank must be increased in order to achieve the same ROT 

achieved in Figure 4-50. 

What does this mean on a practicable side? If a given 

airspeed and bank angle produces a specific ROT, additional 

conclusions can be made. Knowing the ROT is a given 

number of degrees of change per second, the number of 

seconds it takes to travel 360° (a circle) can be determined 

by simple division. For example, if moving at 120 knots with 

a 30° bank angle, the ROT is 5.25° per second and it takes 

68.6 seconds (360° divided by 5.25 = 68.6 seconds) to make 

a complete circle. Likewise, if flying at 240 knots TAS and 

using a 30° angle of bank, the ROT is only about 2.63° per 

second and it takes about 137 seconds to complete a 360° 

circle. Looking at the formula, any increase in airspeed is 

directly proportional to the time the aircraft takes to travel 

an arc. 

So why is this important to understand? Once the ROT is 

understood, a pilot can determine the distance required to make 

that particular turn which is explained in radius of turn.

Radius of Turn

The radius of turn is directly linked to the ROT, which 

explained earlier is a function of both bank angle and 

airspeed. If the bank angle is held constant and the airspeed 

is increased, the radius of the turn changes (increases). A 

higher airspeed causes the aircraft to travel through a longer 

arc due to a greater speed. An aircraft traveling at 120 knots 

is able to turn a 360° circle in a tighter radius than an aircraft 

traveling at 240 knots. In order to compensate for the increase 

in airspeed, the bank angle would need to be increased. 

The radius of turn (R) can be computed using a simple 

formula. The radius of turn is equal to the velocity squared 

(V2) divided by 11.26 times the tangent of the bank angle.

 R = V2

11.26 x tangent of bank angle

Using the examples provided in Figures 4-48 through 4-50, the 

turn radius for each of the two speeds can be computed. 

Note that if the speed is doubled, the radius is squared. 

[Figures 4-51 and 4-52] 
4-35

240 knots

11.26 x tangent of bank angle R = 

 2402

11.26 x tangent of 30°

R =

V2

R = 11.26 x 0.57735

 57,600

R = 8,861 feet

 (four times the radius at 120 knots)

The radius of a turn required by an aircraft traveling at 240 knots 

using the same bank angle in Figure 4-51 is 8,861 feet. Speed is a 

major factor in a turn.

Figure 4-52. Radius at 240 knots.

Figure 4-53. Another formula that can be used for radius.

r = speed (fps) x 360

ROT

Pi (π)

2

202.8 x 68.6

π

2 r =

13,912

π

2 r =

4,428

2 r = = 2,214 feet

Another way to determine the radius of turn is speed in 

using feet per second (fps), π (3.1415) and the ROT. Using 

the example on page 4-34 in the upper right column, it 

was determined that an aircraft with a ROT of 5.25 degrees 

per second required 68.6 seconds to make a complete 

circle. An aircraft’s speed (in knots) can be converted to fps 

by multiplying it by a constant of 1.69. Therefore, an aircraft 

traveling at 120 knots (TAS) travels at 202.8 fps. Knowing the 

speed in fps (202.8) multiplied by the time an aircraft takes 

to complete a circle (68.6 seconds) can determine the size of 

the circle; 202.8 times 68.6 equals 13,912 feet. Dividing by 

π yields a diameter of 4,428 feet, which when divided by 2 

equals a radius of 2,214 feet [Figure 4-53], a foot within that 

determined through use of the formula in Figure 4-51.

In Figure 4-54, the pilot enters a canyon and decides to turn 

180° to exit. The pilot uses a 30° bank angle in his turn.

Weight and Balance

The aircraft’s weight and balance data is important 

information for a pilot that must be frequently reevaluated. 

Although the aircraft was weighed during the certification 

process, this data is not valid indefinitely. Equipment changes 

or modifications affect the weight and balance data. Too often 

pilots reduce the aircraft weight and balance into a “rule of 

thumb” such as: “If I have three passengers, I can load only 

100 gallons of fuel; four passengers, 70 gallons.”

Weight and balance computations should be part of every 

preflight briefing. Never assume three passengers are always 

of equal weight. Instead, do a full computation of all items 

to be loaded on the aircraft, including baggage, as well as 

the pilot and passenger. It is recommended that all bags be 

weighed to make a precise computation of how the aircraft 

CG is positioned.

The importance of the CG was stressed in the discussion 

of stability, controllability, and performance. Unequal load 

distribution causes accidents. A competent pilot understands 

and respects the effects of CG on an aircraft. 

Weight and balance are critical components in the utilization 

of an aircraft to its fullest potential. The pilot must know 

how much fuel can be loaded onto the aircraft without 

violating CG limits, as well as weight limits to conduct 

long or short flights with or without a full complement of 

allowable passengers. For example, an aircraft has four seats 

and can carry 60 gallons of fuel. How many passengers can 

the aircraft safely carry? Can all those seats be occupied at 

all times with the varying fuel loads? Four people who each 

weigh 150 pounds leads to a different weight and balance 

computation than four people who each weigh 200 pounds. 

The second scenario loads an additional 200 pounds onto the 

aircraft and is equal to about 30 gallons of fuel.

The additional weight may or may not place the CG outside 

of the CG envelope, but the maximum gross weight could 

be exceeded. The excess weight can overstress the aircraft 

and degrade the performance. 

Aircraft are certificated for weight and balance for two 

principal reasons:

1. The effect of the weight on the aircraft’s primary 

structure and its performance characteristics

2. The effect of the location of this weight on flight 

characteristics, particularly in stall and spin recovery 

and stability

Aircraft, such as balloons and weight-shift control, do not 

require weight and balance computations because the load 

is suspended below the lifting mechanism. The CG range 

in these types of aircraft is such that it is difficult to exceed 

loading limits. For example, the rear seat position and fuel 

of a weight-shift control aircraft are as close as possible to 

the hang point with the aircraft in a suspended attitude. Thus, 

load variations have little effect on the CG. This also holds 

true for the balloon basket or gondola. While it is difficult 
4-36

5,000 feet

20 20

I0 I0

I0 I0

20 20

I0 I0

2

STBY PWR TEST

2,215 feet

4,430 feet

5 , 0 0 0 f e e t

20 20

I0 I0

I0 I0

20 20

I0 

2

STBY PWR TEST

3,014 feet

6,028 feet

 V2 

11.26 x tangent of the bank angle 30°

R = 1202

11.26 x 0.5773

14,400

6.50096

R = = = 2,215 feet

 V2 

11.26 x tangent of the bank angle 30°

R = 1402

11.26 x 0.5773

19,600

6.50096

R = = = 3,014 feet

120 knots

140 knots

Figure 4-54. Two aircraft have flown into a canyon by error. The canyon is 5,000 feet across and has sheer cliffs on both sides. The pilot in 

the top image is flying at 120 knots. After realizing the error, the pilot banks hard and uses a 30° bank angle to reverse course. This aircraft 

requires about 4,000 feet to turn 180°, and makes it out of the canyon safely. The pilot in the bottom image is flying at 140 knots and also 

uses a 30° angle of bank in an attempt to reverse course. The aircraft, although flying just 20 knots faster than the aircraft in the top image, 

requires over 6,000 feet to reverse course to safety. Unfortunately, the canyon is only 5,000 feet across and the aircraft will hit the canyon 

wall. The point is that airspeed is the most influential factor in determining how much distance is required to turn. Many pilots have made 

the error of increasing the steepness of their bank angle when a simple reduction of speed would have been more appropriate.
4-37

to exceed CG limits in these aircraft, pilots should never 

overload an aircraft because overloading causes structural 

damage and failures. Weight and balance computations are 

not required, but pilots should calculate weight and remain 

within the manufacturer’s established limit.

Effect of Weight on Flight Performance

The takeoff/climb and landing performance of an aircraft are 

determined on the basis of its maximum allowable takeoff and 

landing weights. A heavier gross weight results in a longer 

takeoff run and shallower climb, and a faster touchdown 

speed and longer landing roll. Even a minor overload may 

make it impossible for the aircraft to clear an obstacle that 

normally would not be a problem during takeoff under more 

favorable conditions.

The detrimental effects of overloading on performance are not 

limited to the immediate hazards involved with takeoffs and 

landings. Overloading has an adverse effect on all climb and 

cruise performance which leads to overheating during climbs, 

added wear on engine parts, increased fuel consumption, 

slower cruising speeds, and reduced range.

The manufacturers of modern aircraft furnish weight and 

balance data with each aircraft produced. Generally, this 

information may be found in the FAA-approved AFM/POH 

and easy-to-read charts for determining weight and balance 

data are now provided. Increased performance and loadcarrying capability of these aircraft require strict adherence 

to the operating limitations prescribed by the manufacturer. 

Deviations from the recommendations can result in structural 

damage or complete failure of the aircraft’s structure. Even 

if an aircraft is loaded well within the maximum weight 

limitations, it is imperative that weight distribution be 

within the limits of CG location. The preceding brief study 

of aerodynamics and load factors points out the reasons for 

this precaution. The following discussion is background 

information into some of the reasons why weight and balance 

conditions are important to the safe flight of an aircraft.

In some aircraft, it is not possible to fill all seats, baggage 

compartments, and fuel tanks, and still remain within 

approved weight or balance limits. For example, in several 

popular four-place aircraft, the fuel tanks may not be filled to 

capacity when four occupants and their baggage are carried. 

In a certain two-place aircraft, no baggage may be carried 

in the compartment aft of the seats when spins are to be 

practiced. It is important for a pilot to be aware of the weight 

and balance limitations of the aircraft being flown and the 

reasons for these limitations.

Effect of Weight on Aircraft Structure

The effect of additional weight on the wing structure of an 

aircraft is not readily apparent. Airworthiness requirements 

prescribe that the structure of an aircraft certificated in the 

normal category (in which acrobatics are prohibited) must 

be strong enough to withstand a load factor of 3.8 Gs to take 

care of dynamic loads caused by maneuvering and gusts. This 

means that the primary structure of the aircraft can withstand 

a load of 3.8 times the approved gross weight of the aircraft 

without structural failure occurring. If this is accepted as 

indicative of the load factors that may be imposed during 

operations for which the aircraft is intended, a 100-pound 

overload imposes a potential structural overload of 380 

pounds. The same consideration is even more impressive in 

the case of utility and acrobatic category aircraft, which have 

load factor requirements of 4.4 and 6.0, respectively.

Structural failures which result from overloading may 

be dramatic and catastrophic, but more often they affect 

structural components progressively in a manner that 

is difficult to detect and expensive to repair. Habitual 

overloading tends to cause cumulative stress and damage 

that may not be detected during preflight inspections and 

result in structural failure later during completely normal 

operations. The additional stress placed on structural parts 

by overloading is believed to accelerate the occurrence of 

metallic fatigue failures.

A knowledge of load factors imposed by flight maneuvers 

and gusts emphasizes the consequences of an increase in 

the gross weight of an aircraft. The structure of an aircraft 

about to undergo a load factor of 3 Gs, as in recovery from 

a steep dive, must be prepared to withstand an added load 

of 300 pounds for each 100-pound increase in weight. It 

should be noted that this would be imposed by the addition 

of about 16 gallons of unneeded fuel in a particular aircraft. 

FAA-certificated civil aircraft have been analyzed structurally 

and tested for flight at the maximum gross weight authorized 

and within the speeds posted for the type of flights to be 

performed. Flights at weights in excess of this amount are 

quite possible and often are well within the performance 

capabilities of an aircraft. This fact should not mislead the 

pilot, as the pilot may not realize that loads for which the 

aircraft was not designed are being imposed on all or some 

part of the structure.

In loading an aircraft with either passengers or cargo, the 

structure must be considered. Seats, baggage compartments, 

and cabin floors are designed for a certain load or 

concentration of load and no more. For example, a light 
4-38

Down load on tail

Stronger

CG

Down load on tail

Lighter

CG

Forward CG Center of lift

AFT CG

Load imposed by tail

Gross weight

Load imposed by tail

Gross weight

Figure 4-55. Effect of load distribution on balance.

plane baggage compartment may be placarded for 20 pounds 

because of the limited strength of its supporting structure 

even though the aircraft may not be overloaded or out of CG 

limits with more weight at that location.

Effect of Weight on Stability and Controllability

Overloading also effects stability. An aircraft that is stable 

and controllable when loaded normally may have very 

different flight characteristics when overloaded. Although 

the distribution of weight has the most direct effect on this, 

an increase in the aircraft’s gross weight may be expected 

to have an adverse effect on stability, regardless of location 

of the CG. The stability of many certificated aircraft is 

completely unsatisfactory if the gross weight is exceeded.

Effect of Load Distribution

The effect of the position of the CG on the load imposed 

on an aircraft’s wing in flight is significant to climb and 

cruising performance. An aircraft with forward loading is 

“heavier” and consequently, slower than the same aircraft 

with the CG further aft.

Figure 4-55 illustrates why this is true. With forward loading, 

“nose-up” trim is required in most aircraft to maintain level 

cruising flight. Nose-up trim involves setting the tail surfaces 

to produce a greater down load on the aft portion of the 

fuselage, which adds to the wing loading and the total lift 

required from the wing if altitude is to be maintained. This 

requires a higher AOA of the wing, which results in more 

drag and, in turn, produces a higher stalling speed.

With aft loading and “nose-down” trim, the tail surfaces exert 

less down load, relieving the wing of that much wing loading 

and lift required to maintain altitude. The required AOA 

of the wing is less, so the drag is less, allowing for a faster 

cruise speed. Theoretically, a neutral load on the tail surfaces 

in cruising flight would produce the most efficient overall 

performance and fastest cruising speed, but it would also 

result in instability. Modern aircraft are designed to require 

a down load on the tail for stability and controllability.

A zero indication on the trim tab control is not necessarily 

the same as “neutral trim” because of the force exerted 

by downwash from the wings and the fuselage on the tail 

surfaces.

The effects of the distribution of the aircraft’s useful load 

have a significant influence on its flight characteristics, even 

when the load is within the CG limits and the maximum 

permissible gross weight. Important among these effects 

are changes in controllability, stability, and the actual load 

imposed on the wing.

Generally, an aircraft becomes less controllable, especially 

at slow flight speeds, as the CG is moved further aft. An 

aircraft which cleanly recovers from a prolonged spin with 

the CG at one position may fail completely to respond to 

normal recovery attempts when the CG is moved aft by one 

or two inches.

It is common practice for aircraft designers to establish an aft 

CG limit that is within one inch of the maximum which allows 

normal recovery from a one-turn spin. When certificating an 

aircraft in the utility category to permit intentional spins, the 

aft CG limit is usually established at a point several inches 

forward of that permissible for certification in the normal 

category.

Another factor affecting controllability, which has become 

more important in current designs of large aircraft, is the effect 

of long moment arms to the positions of heavy equipment 

and cargo. The same aircraft may be loaded to maximum 

gross weight within its CG limits by concentrating fuel, 

passengers, and cargo near the design CG, or by dispersing 

fuel and cargo loads in wingtip tanks and cargo bins forward 

and aft of the cabin.

With the same total weight and CG, maneuvering the 

aircraft or maintaining level flight in turbulent air requires 

the application of greater control forces when the load is 

dispersed. The longer moment arms to the positions of the 

heavy fuel and cargo loads must be overcome by the action 

of the control surfaces. An aircraft with full outboard wing 

tanks or tip tanks tends to be sluggish in roll when control 

situations are marginal, while one with full nose and aft cargo 

bins tends to be less responsive to the elevator controls.
4-39

The rearward CG limit of an aircraft is determined largely 

by considerations of stability. The original airworthiness 

requirements for a type certificate specify that an aircraft in 

flight at a certain speed dampens out vertical displacement of 

the nose within a certain number of oscillations. An aircraft 

loaded too far rearward may not do this. Instead, when the 

nose is momentarily pulled up, it may alternately climb and 

dive becoming steeper with each oscillation. This instability 

is not only uncomfortable to occupants, but it could even 

become dangerous by making the aircraft unmanageable 

under certain conditions.

The recovery from a stall in any aircraft becomes progressively 

more difficult as its CG moves aft. This is particularly 

important in spin recovery, as there is a point in rearward 

loading of any aircraft at which a “flat” spin develops. A 

flat spin is one in which centrifugal force, acting through a 

CG located well to the rear, pulls the tail of the aircraft out 

away from the axis of the spin, making it impossible to get 

the nose down and recover.

An aircraft loaded to the rear limit of its permissible CG 

range handles differently in turns and stall maneuvers and 

has different landing characteristics than when it is loaded 

near the forward limit.

The forward CG limit is determined by a number of 

considerations. As a safety measure, it is required that the 

trimming device, whether tab or adjustable stabilizer, be 

capable of holding the aircraft in a normal glide with the 

power off. A conventional aircraft must be capable of a full 

stall, power-off landing in order to ensure minimum landing 

speed in emergencies. A tailwheel-type aircraft loaded 

excessively nose-heavy is difficult to taxi, particularly in 

high winds. It can be nosed over easily by use of the brakes, 

and it is difficult to land without bouncing since it tends to 

pitch down on the wheels as it is slowed down and flared 

for landing. Steering difficulties on the ground may occur in 

nosewheel-type aircraft, particularly during the landing roll 

and takeoff. To summarize the effects of load distribution:

• The CG position influences the lift and AOA of the 

wing, the amount and direction of force on the tail, 

and the degree of deflection of the stabilizer needed 

to supply the proper tail force for equilibrium. The 

latter is very important because of its relationship to 

elevator control force.

• The aircraft stalls at a higher speed with a forward CG 

location. This is because the stalling AOA is reached 

at a higher speed due to increased wing loading.

• Higher elevator control forces normally exist with a 

forward CG location due to the increased stabilizer 

deflection required to balance the aircraft.

• The aircraft cruises faster with an aft CG location 

because of reduced drag. The drag is reduced because 

a smaller AOA and less downward deflection of the 

stabilizer are required to support the aircraft and 

overcome the nose-down pitching tendency.

• The aircraft becomes less stable as the CG is moved 

rearward. This is because when the CG is moved 

rearward it causes an increase in the AOA. Therefore, 

the wing contribution to the aircraft’s stability is 

now decreased, while the tail contribution is still 

stabilizing. When the point is reached that the wing 

and tail contributions balance, then neutral stability 

exists. Any CG movement further aft results in an 

unstable aircraft.

• A forward CG location increases the need for greater 

back elevator pressure. The elevator may no longer 

be able to oppose any increase in nose-down pitching. 

Adequate elevator control is needed to control the 

aircraft throughout the airspeed range down to the 

stall.

A detailed discussion and additional information relating to 

weight and balance can be found in Chapter 9, Weight and 

Balance. 

High Speed Flight

Subsonic Versus Supersonic Flow

In subsonic aerodynamics, the theory of lift is based upon the 

forces generated on a body and a moving gas (air) in which 

it is immersed. At speeds of approximately 260 knots, air 

can be considered incompressible in that, at a fixed altitude, 

its density remains nearly constant while its pressure varies. 

Under this assumption, air acts the same as water and is 

classified as a fluid. Subsonic aerodynamic theory also 

assumes the effects of viscosity (the property of a fluid that 

tends to prevent motion of one part of the fluid with respect 

to another) are negligible, and classifies air as an ideal fluid, 

conforming to the principles of ideal-fluid aerodynamics such 

as continuity, Bernoulli’s principle, and circulation.

In reality, air is compressible and viscous. While the effects of 

these properties are negligible at low speeds, compressibility 

effects in particular become increasingly important as speed 

increases. Compressibility (and to a lesser extent viscosity) is 

of paramount importance at speeds approaching the speed of 

sound. In these speed ranges, compressibility causes a change 

in the density of the air around an aircraft.

During flight, a wing produces lift by accelerating the airflow 

over the upper surface. This accelerated air can, and does, 

reach sonic speeds even though the aircraft itself may be 

flying subsonic. At some extreme AOAs, in some aircraft, 
4-40

Wing airflow

Possible separation

M = 0.50

M = 0.72

M = 0.77

Maximum local velocity is less than sonic

Critical Mach number

Maximum local velocity equal to sonic

Normal shock wave

Supersonic

flow

Subsonic

Figure 4-56. Wing airflow.

CD (Drag coefficient)

 M (Mach number)

0.5 1.0

CL= 0.3

Force divergence Mach number

Critical Mach number

Figure 4-57. Critical Mach.

the speed of the air over the top surface of the wing may be 

double the aircraft’s speed. It is therefore entirely possible 

to have both supersonic and subsonic airflow on an aircraft 

at the same time. When flow velocities reach sonic speeds at 

some location on an aircraft (such as the area of maximum 

camber on the wing), further acceleration results in the onset 

of compressibility effects such as shock wave formation, 

drag increase, buffeting, stability, and control difficulties. 

Subsonic flow principles are invalid at all speeds above this 

point. [Figure 4-56]

Speed Ranges

The speed of sound varies with temperature. Under standard 

temperature conditions of 15 °C, the speed of sound at sea 

level is 661 knots. At 40,000 feet, where the temperature is 

–55 °C, the speed of sound decreases to 574 knots. In highspeed flight and/or high-altitude flight, the measurement of 

speed is expressed in terms of a “Mach number”—the ratio 

of the true airspeed of the aircraft to the speed of sound in 

the same atmospheric conditions. An aircraft traveling at 

the speed of sound is traveling at Mach 1.0. Aircraft speed 

regimes are defined approximately as follows:

Subsonic—Mach numbers below 0.75

Transonic—Mach numbers from 0.75 to 1.20

Supersonic—Mach numbers from 1.20 to 5.00

Hypersonic—Mach numbers above 5.00

While flights in the transonic and supersonic ranges are 

common occurrences for military aircraft, civilian jet aircraft 

normally operate in a cruise speed range of Mach 0.7 to 

Mach 0.90.

The speed of an aircraft in which airflow over any part of 

the aircraft or structure under consideration first reaches 

(but does not exceed) Mach 1.0 is termed “critical Mach 

number” or “Mach Crit.” Thus, critical Mach number is 

the boundary between subsonic and transonic flight and is 

largely dependent on the wing and airfoil design. Critical 

Mach number is an important point in transonic flight. When 

shock waves form on the aircraft, airflow separation followed 

by buffet and aircraft control difficulties can occur. Shock 

waves, buffet, and airflow separation take place above critical 

Mach number. A jet aircraft typically is most efficient when 

cruising at or near its critical Mach number. At speeds 5–10 

percent above the critical Mach number, compressibility 

effects begin. Drag begins to rise sharply. Associated with 

the “drag rise” are buffet, trim and stability changes, and a 

decrease in control surface effectiveness. This is the point 

of “drag divergence.” [Figure 4-57]

VMO/MMO is defined as the maximum operating limit speed. 

VMO is expressed in knots calibrated airspeed (KCAS), while 

MMO is expressed in Mach number. The VMO limit is usually 

associated with operations at lower altitudes and deals with 

structural loads and flutter. The MMO limit is associated with 

operations at higher altitudes and is usually more concerned 

with compressibility effects and flutter. At lower altitudes, 

structural loads and flutter are of concern; at higher altitudes, 

compressibility effects and flutter are of concern. 

Adherence to these speeds prevents structural problems due 

to dynamic pressure or flutter, degradation in aircraft control 

response due to compressibility effects (e.g., Mach Tuck, 

aileron reversal, or buzz), and separated airflow due to shock 

waves resulting in loss of lift or vibration and buffet. Any of 

these phenomena could prevent the pilot from being able to 

adequately control the aircraft.

For example, an early civilian jet aircraft had a VMO limit of 

306 KCAS up to approximately FL 310 (on a standard day). 

At this altitude (FL 310), an MMO of 0.82 was approximately 

equal to 306 KCAS. Above this altitude, an MMO of 0.82 

always equaled a KCAS less than 306 KCAS and, thus, 
4-41

became the operating limit as you could not reach the VMO

limit without first reaching the MMO limit. For example, at 

FL 380, an MMO of 0.82 is equal to 261 KCAS.

Mach Number Versus Airspeed

It is important to understand how airspeed varies with Mach 

number. As an example, consider how the stall speed of a 

jet transport aircraft varies with an increase in altitude. The 

increase in altitude results in a corresponding drop in air 

density and outside temperature. Suppose this jet transport 

is in the clean configuration (gear and flaps up) and weighs 

550,000 pounds. The aircraft might stall at approximately 

152 KCAS at sea level. This is equal to (on a standard day) 

a true velocity of 152 KTAS and a Mach number of 0.23. 

At FL 380, the aircraft will still stall at approximately 152 

KCAS but the true velocity is about 287 KTAS with a Mach 

number of 0.50.

Although the stalling speed has remained the same for our 

purposes, both the Mach number and TAS have increased. 

With increasing altitude, the air density has decreased; this 

requires a faster true airspeed in order to have the same 

pressure sensed by the pitot tube for the same KCAS or KIAS 

(for our purposes, KCAS and KIAS are relatively close to 

each other). The dynamic pressure the wing experiences at 

FL 380 at 287 KTAS is the same as at sea level at 152 KTAS. 

However, it is flying at higher Mach number.

Another factor to consider is the speed of sound. A decrease 

in temperature in a gas results in a decrease in the speed of 

sound. Thus, as the aircraft climbs in altitude with outside 

temperature dropping, the speed of sound is dropping. At 

sea level, the speed of sound is approximately 661 KCAS, 

while at FL 380 it is 574 KCAS. Thus, for our jet transport 

aircraft, the stall speed (in KTAS) has gone from 152 at sea 

level to 287 at FL 380. Simultaneously, the speed of sound 

(in KCAS) has decreased from 661 to 574 and the Mach 

number has increased from 0.23 (152 KTAS divided by 661 

KTAS) to 0.50 (287 KTAS divided by 574 KTAS). All the 

while the KCAS for stall has remained constant at 152. This 

describes what happens when the aircraft is at a constant 

KCAS with increasing altitude, but what happens when the 

pilot keeps Mach constant during the climb? In normal jet 

flight operations, the climb is at 250 KIAS (or higher (e.g. 

heavy)) to 10,000 feet and then at a specified en route climb 

airspeed (such as about 330 if a DC10) until reaching an 

altitude in the “mid-twenties” where the pilot then climbs at 

a constant Mach number to cruise altitude.

Assuming for illustration purposes that the pilot climbs at a 

MMO of 0.82 from sea level up to FL 380. KCAS goes from 

543 to 261. The KIAS at each altitude would follow the 

same behavior and just differ by a few knots. Recall from 

the earlier discussion that the speed of sound is decreasing 

with the drop in temperature as the aircraft climbs. The Mach 

number is simply the ratio of the true airspeed to the speed 

of sound at flight conditions. The significance of this is that 

at a constant Mach number climb, the KCAS (and KTAS or 

KIAS as well) is falling off.

If the aircraft climbed high enough at this constant MMO 

with decreasing KIAS, KCAS, and KTAS, it would begin to 

approach its stall speed. At some point the stall speed of the 

aircraft in Mach number could equal the MMO of the aircraft, 

and the pilot could neither slow up (without stalling) nor 

speed up (without exceeding the max operating speed of the 

aircraft). This has been dubbed the “coffin corner.”

Boundary Layer

The viscous nature of airflow reduces the local velocities on 

a surface and is responsible for skin friction. As discussed 

earlier in the chapter, the layer of air over the wing’s surface 

that is slowed down or stopped by viscosity, is the boundary 

layer. There are two different types of boundary layer flow: 

laminar and turbulent. 

Laminar Boundary Layer Flow

The laminar boundary layer is a very smooth flow, while 

the turbulent boundary layer contains swirls or “eddies.” 

The laminar flow creates less skin friction drag than the 

turbulent flow, but is less stable. Boundary layer flow over 

a wing surface begins as a smooth laminar flow. As the flow 

continues back from the leading edge, the laminar boundary 

layer increases in thickness. 

Turbulent Boundary Layer Flow

At some distance back from the leading edge, the smooth 

laminar flow breaks down and transitions to a turbulent flow. 

From a drag standpoint, it is advisable to have the transition 

from laminar to turbulent flow as far aft on the wing as 

possible, or have a large amount of the wing surface within 

the laminar portion of the boundary layer. The low energy 

laminar flow, however, tends to break down more suddenly 

than the turbulent layer.

Boundary Layer Separation

Another phenomenon associated with viscous flow is 

separation. Separation occurs when the airflow breaks away 

from an airfoil. The natural progression is from laminar 

boundary layer to turbulent boundary layer and then to 

airflow separation. Airflow separation produces high drag 

and ultimately destroys lift. The boundary layer separation 

point moves forward on the wing as the AOA is increased. 

[Figure 4-58]
4-42

Transition region

Laminar sublayer

Laminar boundary layer

Turbulent boundary layer

Figure 4-58. Boundary layer.

Vortex generators are used to delay or prevent shock wave 

induced boundary layer separation encountered in transonic 

flight. They are small low aspect ratio airfoils placed at a 12° 

to 15° AOA to the airstream. Usually spaced a few inches 

apart along the wing ahead of the ailerons or other control 

surfaces, vortex generators create a vortex which mixes the 

boundary airflow with the high energy airflow just above the 

surface. This produces higher surface velocities and increases 

the energy of the boundary layer. Thus, a stronger shock wave 

is necessary to produce airflow separation.

Shock Waves

When an airplane flies at subsonic speeds, the air ahead is 

“warned” of the airplane’s coming by a pressure change 

transmitted ahead of the airplane at the speed of sound. 

Because of this warning, the air begins to move aside before 

the airplane arrives and is prepared to let it pass easily. When 

the airplane’s speed reaches the speed of sound, the pressure 

change can no longer warn the air ahead because the airplane 

is keeping up with its own pressure waves. Rather, the air 

particles pile up in front of the airplane causing a sharp 

decrease in the flow velocity directly in front of the airplane 

with a corresponding increase in air pressure and density.

As the airplane’s speed increases beyond the speed of sound, 

the pressure and density of the compressed air ahead of it 

increase, the area of compression extending some distance 

ahead of the airplane. At some point in the airstream, the air 

particles are completely undisturbed, having had no advanced 

warning of the airplane’s approach, and in the next instant the 

same air particles are forced to undergo sudden and drastic 

changes in temperature, pressure, density, and velocity. 

The boundary between the undisturbed air and the region 

of compressed air is called a shock or “compression” wave. 

This same type of wave is formed whenever a supersonic 

airstream is slowed to subsonic without a change in direction, 

such as when the airstream is accelerated to sonic speed 

over the cambered portion of a wing, and then decelerated 

to subsonic speed as the area of maximum camber is passed. 

A shock wave forms as a boundary between the supersonic 

and subsonic ranges.

Whenever a shock wave forms perpendicular to the airflow, it 

is termed a “normal” shock wave, and the flow immediately 

behind the wave is subsonic. A supersonic airstream passing 

through a normal shock wave experiences these changes:

• The airstream is slowed to subsonic.

• The airflow immediately behind the shock wave does 

not change direction.

• The static pressure and density of the airstream behind 

the wave is greatly increased.

 • The energy of the airstream (indicated by total 

pressure—dynamic plus static) is greatly reduced.

Shock wave formation causes an increase in drag. One of 

the principal effects of a shock wave is the formation of a 

dense high pressure region immediately behind the wave. 

The instability of the high pressure region, and the fact that 

part of the velocity energy of the airstream is converted to 

heat as it flows through the wave is a contributing factor 

in the drag increase, but the drag resulting from airflow 

separation is much greater. If the shock wave is strong, 

the boundary layer may not have sufficient kinetic energy 

to withstand airflow separation. The drag incurred in the 

transonic region due to shock wave formation and airflow 

separation is known as “wave drag.” When speed exceeds 

the critical Mach number by about 10 percent, wave drag 

increases sharply. A considerable increase in thrust (power) 

is required to increase flight speed beyond this point into the 

supersonic range where, depending on the airfoil shape and 

the angle of attack, the boundary layer may reattach.

Normal shock waves form on the wing’s upper surface and 

form an additional area of supersonic flow and a normal shock 

wave on the lower surface. As flight speed approaches the 

speed of sound, the areas of supersonic flow enlarge and the 

shock waves move nearer the trailing edge. [Figure 4-59]

Associated with “drag rise” are buffet (known as Mach 

buffet), trim and stability changes, and a decrease in control 
4-43

Shock waves

M = 0.82

M = 0.95

M = 1.05

Normal shock wave

Supersonic flow

Normal shock wave

Supersonic flow

Subsonic airflow

Bow wave

Figure 4-59. Shock waves.

Spanwise flowTrue airspeed Mach 0.85

Airspeed sensed by wing Mach 0.70

Figure 4-60. Sweepback effect.

force effectiveness. The loss of lift due to airflow separation 

results in a loss of downwash, and a change in the position of 

the center pressure on the wing. Airflow separation produces 

a turbulent wake behind the wing, which causes the tail 

surfaces to buffet (vibrate). The nose-up and nose-down pitch 

control provided by the horizontal tail is dependent on the 

downwash behind the wing. Thus, an increase in downwash 

decreases the horizontal tail’s pitch control effectiveness 

since it effectively increases the angle of attack that the tail 

surface is seeing. Movement of the wing CP affects the wing 

pitching moment. If the CP moves aft, a diving moment 

referred to as “Mach tuck” or “tuck under” is produced, 

and if it moves forward, a nose-up moment is produced. 

This is the primary reason for the development of the T-tail 

configuration on many turbine-powered aircraft, which 

places the horizontal stabilizer as far as practical from the 

turbulence of the wings.

Sweepback

Most of the difficulties of transonic flight are associated with 

shock wave induced flow separation. Therefore, any means of 

delaying or alleviating the shock induced separation improves 

aerodynamic performance. One method is wing sweepback. 

Sweepback theory is based upon the concept that it is only the 

component of the airflow perpendicular to the leading edge 

of the wing that affects pressure distribution and formation 

of shock waves. [Figure 4-60]

On a straight wing aircraft, the airflow strikes the wing 

leading edge at 90°, and its full impact produces pressure and 

lift. A wing with sweepback is struck by the same airflow at 

an angle smaller than 90°. This airflow on the swept wing has 

the effect of persuading the wing into believing that it is flying 

slower than it really is; thus the formation of shock waves is 

delayed. Advantages of wing sweep include an increase in 

critical Mach number, force divergence Mach number, and 

the Mach number at which drag rises peaks. In other words, 

sweep delays the onset of compressibility effects.

The Mach number, which produces a sharp change in drag 

coefficient, is termed the “force divergence” Mach number 

and, for most airfoils, usually exceeds the critical Mach 

number by 5 to 10 percent. At this speed, the airflow separation 

induced by shock wave formation can create significant 

variations in the drag, lift, or pitching moment coefficients. 

In addition to the delay of the onset of compressibility effects, 

sweepback reduces the magnitude in the changes of drag, lift 

or moment coefficients. In other words, the use of sweepback 

“softens” the force divergence.
4-44

Intitial stall area

Figure 4-62. T-tail stall.

Stalled

Prestall

Figure 4-61. Wingtip stall.

A disadvantage of swept wings is that they tend to stall at the 

wingtips rather than at the wing roots. [Figure 4-61] This is 

because the boundary layer tends to flow spanwise toward the 

tips and to separate near the leading edges. Because the tips of 

a swept wing are on the aft part of the wing (behind the CL), 

a wingtip stall causes the CL to move forward on the wing, 

forcing the nose to rise further. The tendency for tip stall is 

greatest when wing sweep and taper are combined. 

The stall situation can be aggravated by a T-tail configuration, 

which affords little or no pre-stall warning in the form of 

tail control surface buffet. [Figure 4-62] The T-tail, being 

above the wing wake remains effective even after the wing 

has begun to stall, allowing the pilot to inadvertently drive 

the wing into a deeper stall at a much greater AOA. If the 

horizontal tail surfaces then become buried in the wing’s 

wake, the elevator may lose all effectiveness, making it 

impossible to reduce pitch attitude and break the stall. In 

the pre-stall and immediate post-stall regimes, the lift/drag 

qualities of a swept wing aircraft (specifically the enormous 

increase in drag at low speeds) can cause an increasingly 

descending flightpath with no change in pitch attitude, further 

increasing the AOA. In this situation, without reliable AOA 

information, a nose-down pitch attitude with an increasing 

airspeed is no guarantee that recovery has been effected, 

and up-elevator movement at this stage may merely keep 

the aircraft stalled.

It is a characteristic of T-tail aircraft to pitch up viciously 

when stalled in extreme nose-high attitudes, making 

recovery difficult or violent. The stick pusher inhibits this 

type of stall. At approximately one knot above stall speed, 

pre-programmed stick forces automatically move the stick 

forward, preventing the stall from developing. A G-limiter 

may also be incorporated into the system to prevent the pitch 

down generated by the stick pusher from imposing excessive 

loads on the aircraft. A “stick shaker,” on the other hand 

provides stall warning when the airspeed is five to seven 

percent above stall speed.

Mach Buffet Boundaries

Mach buffet is a function of the speed of the airflow over the 

wing—not necessarily the speed of the aircraft. Any time that 

too great a lift demand is made on the wing, whether from too 

fast an airspeed or from too high an AOA near the MMO, the 

“high-speed” buffet occurs. There are also occasions when 

the buffet can be experienced at much lower speeds known 

as the “low-speed Mach buffet.”

An aircraft flown at a speed too slow for its weight and 

altitude necessitating a high AOA is the most likely situation 
4-45

to cause a low-speed Mach buffet. This very high AOA has 

the effect of increasing airflow velocity over the upper surface 

of the wing until the same effects of the shock waves and 

buffet occur as in the high-speed buffet situation. The AOA of 

the wing has the greatest effect on inducing the Mach buffet at 

either the high-speed or low-speed boundaries for the aircraft. 

The conditions that increase the AOA, the speed of the airflow 

over the wing, and chances of Mach buffet are:

• High altitudes—the higher an aircraft flies, the thinner 

the air and the greater the AOA required to produce 

the lift needed to maintain level flight.

• Heavy weights—the heavier the aircraft, the greater 

the lift required of the wing, and all other things being 

equal, the greater the AOA.

• G loading—an increase in the G loading on the aircraft 

has the same effect as increasing the weight of the 

aircraft. Whether the increase in G forces is caused 

by turns, rough control usage, or turbulence, the effect 

of increasing the wing’s AOA is the same.

High Speed Flight Controls

On high-speed aircraft, flight controls are divided into primary 

flight controls and secondary or auxiliary flight controls. The 

primary flight controls maneuver the aircraft about the pitch, 

roll, and yaw axes. They include the ailerons, elevator, and 

rudder. Secondary or auxiliary flight controls include tabs, 

leading edge flaps, trailing edge flaps, spoilers, and slats.

Spoilers are used on the upper surface of the wing to spoil 

or reduce lift. High speed aircraft, due to their clean low 

drag design use spoilers as speed brakes to slow them 

down. Spoilers are extended immediately after touchdown 

to dump lift and thus transfer the weight of the aircraft from 

the wings onto the wheels for better braking performance. 

[Figure 4-63]

Jet transport aircraft have small ailerons. The space for 

ailerons is limited because as much of the wing trailing 

edge as possible is needed for flaps. Also, a conventional 

size aileron would cause wing twist at high speed. For that 

reason, spoilers are used in unison with ailerons to provide 

additional roll control.

Some jet transports have two sets of ailerons, a pair of 

outboard low-speed ailerons and a pair of high-speed inboard 

ailerons. When the flaps are fully retracted after takeoff, the 

outboard ailerons are automatically locked out in the faired 

position.

When used for roll control, the spoiler on the side of the 

up-going aileron extends and reduces the lift on that side, 

causing the wing to drop. If the spoilers are extended as 

speed brakes, they can still be used for roll control. If they 

are the differential type, they extend further on one side 

and retract on the other side. If they are the non-differential 

type, they extend further on one side but do not retract on the 

other side. When fully extended as speed brakes, the nondifferential spoilers remain extended and do not supplement 

the ailerons.

To obtain a smooth stall and a higher AOA without airflow 

separation, the wing’s leading edge should have a wellrounded almost blunt shape that the airflow can adhere to 

at the higher AOA. With this shape, the airflow separation 

starts at the trailing edge and progresses forward gradually 

as AOA is increased.

The pointed leading edge necessary for high-speed flight 

results in an abrupt stall and restricts the use of trailing edge 

flaps because the airflow cannot follow the sharp curve 

around the wing leading edge. The airflow tends to tear loose 

rather suddenly from the upper surface at a moderate AOA. 

To utilize trailing edge flaps, and thus increase the CL-MAX, 

the wing must go to a higher AOA without airflow separation. 

Therefore, leading edge slots, slats, and flaps are used to 

improve the low-speed characteristics during takeoff, climb, 

and landing. Although these devices are not as powerful as 

trailing edge flaps, they are effective when used full span in 

combination with high-lift trailing edge flaps. With the aid 

of these sophisticated high-lift devices, airflow separation is 

delayed and the CL-MAX is increased considerably. In fact, a 

50 knot reduction in stall speed is not uncommon.

The operational requirements of a large jet transport aircraft 

necessitate large pitch trim changes. Some requirements 

are:

• A large CG range

• A large speed range

• The ability to perform large trim changes due to 

wing leading edge and trailing edge high-lift devices 

without limiting the amount of elevator remaining

• Maintaining trim drag to a minimum

These requirements are met by the use of a variable incidence 

horizontal stabilizer. Large trim changes on a fixed-tail 

aircraft require large elevator deflections. At these large 

deflections, little further elevator movement remains in the 

same direction. A variable incidence horizontal stabilizer 

is designed to take out the trim changes. The stabilizer is 

larger than the elevator, and consequently does not need to 

be moved through as large an angle. This leaves the elevator 

streamlining the tail plane with a full range of movement up 

and down. The variable incidence horizontal stabilizer can 
4-46

737 Control Surfaces

727 Control Surfaces

Takeoff Flaps

Flaps Retracted

Landing Flaps

Aileron

Leading edge slat

Leading edge flap

Foreflap Midflap Aftflap

Inboard wing

Outboard wing

Inboard wing

Outboard wing

Foreflap

Midflap

Aftflap

Aileron

Inboard wing

Outboard wing

Aileron

Rudder

Elevator

Tab

Stabilizer

Inboard flap

Flight spoilers

Ground spoiler

Outboard flap

Tab

Aileron

Leading edge slats

Leading edge flaps

Stabilizer

Elevator

Control tab

Upper rudder

Anti-balance tabs

Control tab

Inboard aileron

Leading edge flaps

Lower rudder

Vortex generators

Pitot tubes

Ground spoilers

Inboard flaps

Flight spoilers

Outboard flap

Balance tab

Outboard aileron

Slats

Fence

Figure 4-63. Control surfaces.

be set to handle the bulk of the pitch control demand, with 

the elevator handling the rest. On aircraft equipped with a 

variable incidence horizontal stabilizer, the elevator is smaller 

and less effective in isolation than it is on a fixed-tail aircraft. 

In comparison to other flight controls, the variable incidence 

horizontal stabilizer is enormously powerful in its effect. 

Because of the size and high speeds of jet transport aircraft, 

the forces required to move the control surfaces can be 

beyond the strength of the pilot. Consequently, the control 

surfaces are actuated by hydraulic or electrical power units. 

Moving the controls in the flight deck signals the control 

angle required, and the power unit positions the actual 

control surface. In the event of complete power unit failure, 

movement of the control surface can be effected by manually 

controlling the control tabs. Moving the control tab upsets 

the aerodynamic balance which causes the control surface 

to move.

Chapter Summary

In order to sustain an aircraft in flight, a pilot must understand 

how thrust, drag, lift, and weight act on the aircraft. By 

understanding the aerodynamics of flight, how design, 

weight, load factors, and gravity affect an aircraft during 

flight maneuvers from stalls to high speed flight, the pilot 

learns how to control the balance between these forces. For 

information on stall speeds, load factors, and other important 

aircraft data, always consult the AFM/POH for specific 

information pertaining to the aircraft being flown.