Chapter 4 - Aerodynamics of Flight
Aerodynamics of Flight
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.
- 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).
Figure 4-1. Relationship of forces acting on an airplane.
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]
Figure 4-2. Force vectors during a stabilized climb.
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.
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 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.
Figure 4-3. Angle of attack at various speeds.
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]
Figure 4-4. Some aircraft have the ability to change the direction of thrust.
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 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 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]
Figure 4-5. Form drag.
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 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]
Figure 4-6. A wing root can cause interference drag.
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).
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.
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.
Figure 4-7. Wingtip vortex from a crop duster.
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.
Figure 4-8. Drag versus speed.
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]
Figure 4-9. Lift coefficients at various angles of attack.
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
D = CD . ρ . V2 . S
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 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.
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.
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, weight-shift 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 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-and-level 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.
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]
Figure 4-10. Wingtip vortices.
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]
Figure 4-11. Avoid following another aircraft at an altitude within 1,000 feet.
Figure 4-12. Avoid turbulence from another aircraft.
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.
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.
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 three-dimensional 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 due to the interference of the ground (or water) surface with the airﬂow patterns about the aircraft in ﬂight.
Figure 4-13. Ground effect changes airflow.
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 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]
Figure 4-14. Ground effect changes drag and lift.
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.
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]
Figure 4-15. Axes of an airplane.
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 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.
Figure 4-16. A weight-shift control aircraft.
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]
Figure 4-17. A powered parachute.
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 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 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 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
Figure 4-18. Types of static 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.
Figure 4-19. Damped versus undamped stability.
- 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.
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:
- Location of the wing with respect to the CG
- Location of the horizontal tail surfaces with respect to the CG
- 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.
Figure 4-20. Longitudinal stability.
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 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.
Figure 4-21. Effect of speed on downwash.
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.
Figure 4-22. Reduced power allows pitch down.
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 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.
Figure 4-23. Thrust line affects longitudinal stability.
Figure 4-24. Power changes affect longitudinal stability.
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.
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.
Figure 4-25. Dihedral for lateral stability.
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.
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 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.
Figure 4-26. Keel area for lateral stability.
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.
Figure 4-27. Fuselage and fin for vertical 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 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 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 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.
Figure 4-28. Forces during normal coordinated turn.
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. 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 constant-rate 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.
Figure 4-29. Normal, slipping, and skidding turns.
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 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.
Figure 4-30. Changes in lift during climb entry.
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.
Figure 4-31. Changes in speed during climb entry.
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 straight-and-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 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-and-level 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.
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 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.
Figure 4-32. Forces exerted when pulling out of a dive.
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.
Figure 4-33. Increase in stall speed and load factor.
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.
Figure 4-34. Inflight ice formation.
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.
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.
Figure 4-36. Propeller blade angle.
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.
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.
Figure 4-37. Propeller slippage.
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.
Figure 4-38. Propeller tips travel faster than the hub.
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 constant-speed 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.
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:
- Torque reaction from engine and propeller,
- Corkscrewing effect of the slipstream,
- Gyroscopic action of the propeller, and
- Asymmetric loading of the propeller (P-factor).
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]
Figure 4-39. Torque reaction.
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:
- Size and horsepower of engine,
- Size of propeller and the rpm,
- Size of the aircraft, and
- 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.
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]
Figure 4-40. Corkscrewing slipstream.
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.
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.
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.
Figure 4-41. Gyroscopic precession.
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.
Figure 4-42. Raising tail produces gyroscopic precession.
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 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.
Figure 4-43. Asymmetrical loading of propeller (P-factor).
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.
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:
- It is possible for a pilot to impose a dangerous overload on the aircraft structures.
- An increased load factor increases the stalling speed and makes stalls possible at seemingly safe ﬂight speeds.
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.
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 present-day 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 4-44. Two forces cause load factor during turns.
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.
Figure 4-45. Angle of bank changes load factor.
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 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]
Figure 4-46. Load factor changes stall speed.
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.
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.)
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:
- Stalls from steep turns or excessive maneuvering near the ground
- 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.
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°.
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.
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:
- Airspeed in a spin is very low, usually within 2 knots of the unaccelerated stalling speeds.
- 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.
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.
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.
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.
Figure 4-47. Typical Vg diagram.
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 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).
Figure 4-48. Rate of turn for a given airspeed (knots, TAS) and bank angle.
Figure 4-49. Rate of turn when increasing speed.
Figure 4-50. To achieve the same rate of turn of an aircraft traveling at 120 knots, an increase of bank angle is required.
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]
Figure 4-51. Radius at 120 knots with bank angle of 30°.
Figure 4-52. Radius at 240 knots.
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.
Figure 4-53. Another formula that can be used for radius.
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.
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.
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:
- The effect of the weight on the aircraft’s primary structure and its performance characteristics
- 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 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 load-carrying 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 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.
Figure 4-55. Effect of load distribution on balance.
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.
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, 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]
Figure 4-56. Wing airflow.
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 high-speed 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]
Figure 4-57. Critical Mach.
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, 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.”
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]
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.
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]
Figure 4-59. Shock waves.
Associated with “drag rise” are buffet (known as Mach buffet), trim and stability changes, and a decrease in control 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.
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]
Figure 4-60. Sweepback effect.
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.
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.
Figure 4-61. Wingtip stall.
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.
Figure 4-62. T-tail stall.
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 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]
Figure 4-63. Control surfaces.
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 non-differential 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 well-rounded 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 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.
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.