AIRCRAFT SYSTEMS - PROPELLER

The propeller is a rotating airfoil, subject to induced drag, stalls, and other aerodynamic principles that apply to any airfoil. It provides the necessary thrust to pull, or in some cases push, the airplane through the air.

The engine power is used to rotate the propeller, which in turn generates thrust very similar to the manner in which a wing produces lift. The amount of thrust produced depends on the shape of the airfoil, the angle of attack of the propeller blade, and the r.p.m. of the engine. The propeller itself is twisted so the blade angle changes from hub to tip. The greatest angle of incidence, or the highest pitch, is at the hub while the smallest pitch is at the tip.

The reason for the twist is to produce uniform lift from the hub to the tip. As the blade rotates, there is a difference in the actual speed of the various portions of the blade. The tip of the blade travels faster than that part near the hub, because the tip travels a greater distance than the hub in the same length of time. Changing the angle of incidence (pitch) from the hub to the tip to correspond with the speed produces uniform lift throughout the length of the blade. If the propeller blade was designed with the same angle of incidence throughout its entire length, it would be inefficient, because as airspeed increases in flight, the portion near the hub would have a negative angle of attack while the blade tip would be stalled.

Small airplanes are equipped with either one of two types of propellers. One is the fixed-pitch, and the other is the controllable-pitch.

AIRCRAFT SYSTEMS - RECIPROCATING ENGINES

Most small airplanes are designed with reciprocating engines. The name is derived from the back-and-forth, or reciprocating, movement of the pistons. It is this motion that produces the mechanical energy needed to accomplish work. Two common means of classifying reciprocating engines are:
1. by cylinder arrangement with respect to the crankshaft—radial, in-line, v-type or opposed, or
2. by the method of cooling—liquid or air-cooled.

Radial engines were widely used during World War II, and many are still in service today. With these engines, a row or rows of cylinders are arranged in a circular pattern around the crankcase. The main advantage of a radial engine is the favorable power-to-weight ratio.

In-line engines have a comparatively small frontal area, but their power-to-weight ratios are relatively low. In addition, the rearmost cylinders of an air-cooled, in-line engine receive very little cooling air, so these engines are normally limited to four or six cylinders.

V-type engines provide more horsepower than in-line engines and still retain a small frontal area. Further improvements in engine design led to the development of the horizontally-opposed engine. Opposed-type engines are the most popular reciprocating engines used on small airplanes. These engines always have an even number of cylinders, since a cylinder on one side of the crankcase "opposes" a cylinder on the other side. The majority of these engines are air cooled and usually are mounted in a horizontal position when installed on fixed-wing airplanes. Opposed-type engines have high power-toweight ratios because they have a comparatively small, lightweight crankcase. In addition, the compact cylinder arrangement reduces the engine's frontal area and allows a streamlined installation that minimizes aerodynamic drag.

Powerplant—A complete engine and propeller combination with accessories.


The main parts of a reciprocating engine include the cylinders, crankcase, and accessory housing. The intake/exhaust valves, spark plugs, and pistons are located in the cylinders. The crankshaft and connecting rods are located in the crankcase. The magnetos are normally located on the engine accessory housing.

The basic principle for reciprocating engines involves the conversion of chemical energy, in the form of fuel, into mechanical energy. This occurs within the cylinders of the engine through a process known as the four-stroke operating cycle. These strokes are called intake, compression, power, and exhaust.

1. The intake stroke begins as the piston starts its downward travel. When this happens, the intake valve opens and the fuel/air mixture is drawn into the cylinder.
2. The compression stroke begins when the intake valve closes and the piston starts moving back to the top of the cylinder. This phase of the cycle is used to obtain a much greater power output from the fuel/air mixture once it is ignited.
3. The power stroke begins when the fuel/air mixture is ignited. This causes a tremendous pressure increase in the cylinder, and forces the piston downward away from the cylinder head, creating the power that turns the crankshaft.
4. The exhaust stroke is used to purge the cylinder of burned gases. It begins when the exhaust valve opens and the piston starts to move toward the cylinder head once again.

Even when the engine is operated at a fairly low speed, the four-stroke cycle takes place several hundred times each minute. In a four-cylinder engine, each cylinder operates on a different stroke. Continuous rotation of a crankshaft is maintained by the precise timing of the power strokes in each cylinder. Continuous operation of the engine depends on the simultaneous function of auxiliary systems, including the induction, ignition, fuel, oil, cooling, and exhaust systems.

AIRCRAFT SYSTEMS - POWERPLANT

This section covers the main systems found on small airplanes. These include the engine, propeller, and induction systems, as well as the ignition, fuel, lubrication, cooling, electrical, landing gear, autopilot, and environmental control systems. A comprehensive introduction to gas turbine engines is included at the end of this section.

POWERPLANT
The airplane engine and propeller, often referred to as a powerplant, work in combination to produce thrust. The powerplant propels the airplane and drives the various systems that support the operation of an airplane.

SECONDARY FLIGHT CONTROLS - ADJUSTABLE STABILIZER

Rather than using a movable tab on the trailing edge of the elevator, some airplanes have an adjustable stabilizer. With this arrangement, linkages pivot the horizontal stabilizer about its rear spar. This is accomplished by use of a jackscrew mounted on the leading edge of the stabilator.

On small airplanes, the jackscrew is cable-operated with a trim wheel or crank, and on larger airplanes, it is motor driven. The trimming effect and cockpit indications for an adjustable stabilizer are similar to those of a trim tab.
Since the primary and secondary flight control systems vary extensively between aircraft, you need to be familiar with the systems in your aircraft. A good source of information is the Airplane Flight Manual (AFM) or the Pilot's Operating Handbook (POH).

SECONDARY FLIGHT CONTROLS - GROUND ADJUSTABLE TABS

Many small airplanes have a non-moveable metal trim tab on the rudder. This tab is bent in one direction or the other while on the ground to apply a trim force to the rudder. The correct displacement is determined by trial-and-error process. Usually, small adjustments are necessary until you are satisfied that the airplane is no longer skidding left or right during normal cruising flight.

SECONDARY FLIGHT CONTROLS - ANTISERVO TABS

In addition to decreasing the sensitivity of the stabilator, an antiservo tab also functions as a trim device to relieve control pressure and maintain the stabilator in the desired position. The fixed end of the linkage is on the opposite side of the surface from the horn on the tab, and when the trailing edge of the stabilator moves up, the linkage forces the trailing edge of the tab up. When the stabilator moves down, the tab also moves down. This is different than trim tabs on elevators, which move opposite of the control surface.

This tab works in the same manner as the balance tab except that, instead of moving in the opposite direction, it moves in the same direction as the trailing edge of the stabilator. For example, when the trailing edge of the stabilator moves up, the linkage forces the trailing edge of the tab up. When the stabilator moves down, the tab also moves down.

SECONDARY FLIGHT CONTROLS - BALANCE TABS

The control forces may be excessively high in some airplanes, and in order to decrease them, the manufacturer may use balance tabs. They look like trim tabs and are hinged in approximately the same places as trim tabs. The essential difference between the two is that the balancing tab is coupled to the control surface rod so that when the primary control surface is moved in any direction, the tab automatically moves in the opposite direction. In this manner, the airflow striking the tab counter-balances some of the air pressure against the primary control surface, and enables the pilot to more easily move and hold the control surface in position.

If the linkage between the tab and the fixed surface is adjustable from the cockpit, the tab acts as a combination trim and balance tab, which can be adjusted to any desired deflection. Any time the control surface is deflected, the tab moves in the opposite direction and eases the load on the pilot.

SECONDARY FLIGHT CONTROLS - TRIM TABS

The most common installation on small airplanes is a single trim tab attached to the trailing edge of the elevator. A small, vertically mounted control wheel manually operates most trim tabs. However,
a trim crank may be found in some airplanes. The cockpit control includes a tab position indicator. Placing the trim control in the full nose-down position moves the tab to its full up position. With the tab up and into the airstream, the airflow over the horizontal tail surface tends to force the trailing edge of the elevator down. This causes the tail of the airplane to move up, and results in a nose-down pitch change.

If you set the trim tab to the full nose-up position, the tab moves to its full-down position. In this case, the air flowing under the horizontal tail surface hits the tab and tends to force the trailing edge of the elevator up, reducing the elevator's angle of attack. This causes a tail-down movement of the airplane and a nose-up pitch change.

In spite of the opposite direction movement of the trim tab and the elevator, control of trim is natural to a pilot. If you have to exert constant back pressure on the control column, the need for nose-up trim is indicated. The normal trim procedure is to continue trimming until the airplane is balanced and the nose-heavy condition is no longer apparent. Pilots normally establish the desired power, pitch attitude, and configuration first, and then trim the airplane to relieve control pressures that may exist for that flight condition. Any time power, pitch attitude, or configuration is changed, expect that retrimming will be necessary to relieve the control pressures for the new flight condition.

SECONDARY FLIGHT CONTROLS - TRIM SYSTEMS

Although the airplane can be operated throughout a wide range of attitudes, airspeeds, and power settings, it can only be designed to fly hands off within a very limited combination of these variables.

Therefore, trim systems are used to relieve the pilot of the need to maintain constant pressure on the flight controls. Trim systems usually consist of cockpit controls and small hinged devices attached to the trailing edge of one or more of the primary flight control surfaces. They are designed to help minimize a pilot's workload by aerodynamically assisting movement and position of the flight control surface to which they are attached. Common types of trim systems include trim tabs, balance tabs, antiservo tabs, ground adjustable tabs, and an adjustable stabilizer.

SECONDARY FLIGHT CONTROLS - SPOILERS

On some airplanes, high-drag devices called spoilers are deployed from the wings to spoil the smooth airflow, reducing lift and increasing drag. Spoilers are used for roll control on some aircraft, one of the
advantages being the elimination of adverse yaw. To turn right, for example, the spoiler on the right wing is raised, destroying some of the lift and creating more drag on the right. The right wing drops, and the airplane banks and yaws to the right. Deploying spoilers on both wings at the same time allows the aircraft to descend without gaining speed. Spoilers are also deployed to help shorten ground roll after landing. By destroying lift, they transfer weight to the wheels, improving braking effectiveness.

SECONDARY FLIGHT CONTROLS - LEADING EDGE DEVICES

High-lift devices also can be applied to the leading edge of the airfoil. The most common types are fixed slots, movable slats, and leading edge flaps.

Fixed slots direct airflow to the upper wing surface and delay airflow separation at higher angles of attack. The slot does not increase the wing camber, but allows a higher maximum coefficient of lift because the stall is delayed until the wing reaches greater angle of attacks.

Movable slats consist of leading edge segments, which move on tracks. At low angles of attack, each slat is held flush against the wing's leading edge by the high pressure that forms at the wing's leading edge. As the angle of attack increases, the high-pressure area moves aft below the lower surface of the wing, allowing the slats to move forward. Some slats, however, are pilot operated and can be deployed at any angle of attack. Opening a slat allows the air below the wing to flow over the wing's upper surface, delaying airflow separation.

Leading edge flaps, like trailing edge flaps, are used to increase both CLmax and the camber of the wings. This type of leading edge device is frequently used in conjunction with trailing edge flaps and can reduce the nose-down pitching movement produced by the latter.

As is true with trailing edge flaps, a small increment of leading edge flaps increases lift to a much greater extent than drag. As greater amounts of flaps are extended, drag increases at a greater rate than lift.

FLIGHT CONTROLS - CANARD

The term canard refers to a control surface with functions as a horizontal stabilizer but is located in front of the main wings. The term also is used to describe an airplane equipped with a canard. In effect, it is an airfoil similar to the horizontal surface on a conventional aft-tail design. The difference is that the canard actually creates lift and holds the nose up, as opposed to the aft-tail design that exerts downward force on the tail to prevent the nose from rotating downward.

Canard --A horizontal surface mounted ahead of the main wing to provide longitudinal stability and control. It may be a fixed, movable, or variable geometry surface, with or without control surfaces.

Canard Configuration --A configuration in which the span of the forward wings is substantially less than that of the main wing.

Although the Wright Flyer was configured as a canard with the horizontal surfaces in front of the lifting surface, it was not until recently that the canard configuration began appearing on newer airplanes. Canard designs include two types—one with a horizontal surface of about the same size as a normal aft-tail design, and the other with a surface of the same approximate size and airfoil of the aft-mounted wing known as a tandem wing configuration. Theoretically, the canard is considered more efficient because using the horizontal surface to help lift the weight of the aircraft should result in less drag for a given amount of lift. The canard's main advantage is in the area of stall characteristics. A properly designed canard or tandem wing will run out of authority to raise the nose of the aircraft at a point before the main wing will stall. This makes the aircraft stall-proof and results only in a descent rate that can be halted by adding power. Ailerons on the main wing remain effective throughout the recovery. Other canard configurations are designed so the canard stalls before the main wing, automatically lowering the nose and recovering the aircraft to a safe flying speed. Again, the ailerons remain effective throughout the stall. The canard design has several limitations. First, it is important that the forward lifting surface of a canard design stalls before the main wing. If the main wing stalls first, the lift remaining from the forward wing or canard would be well ahead of the CG, and the airplane would pitch up uncontrollably. Second, when the forward surface stalls first, or is limited in its ability to increase the angle of attack, the main wing never reaches a point where its maximum lift is created, sacrificing some performance. Third, use of flaps on the main wing causes design problems for the forward wing or canard. As lift on the main wing is increased by extension of flaps, the lift requirement of the canard is also increased. The forward wing or canard must be large enough to accommodate flap use, but not so large that it creates more lift than the main wing. Finally, the relationship of the main wing to the forward surface also makes a difference. When positioned closely in the vertical plane, downwash from the forward wing can have a negative effect on the lift of the main wing. Increasing vertical separation increases efficiency of the design. Efficiency is also increased, as the size of the two surfaces grows closer to being equal.

SECONDARY FLIGHT CONTROLS - FLAPS

Secondary flight control systems may consist of the flaps, leading edge devices, spoilers, and trim devices.

Flaps are the most common high-lift devices used on practically all airplanes. These surfaces, which are attached to the trailing edge of the wing, increase both lift and induced drag for any given angle of attack. Flaps allow a compromise between high cruising speed and low landing speed, because they may be extended when needed, and retracted into the wing's structure when not needed. There are four common types of flaps: plain, split, slotted, and Fowler flaps.

The plain flap is the simplest of the four types. It increases the airfoil camber, resulting in a significant increase in the coefficient of lift at a given angle of attack. At the same time, it greatly increases drag and moves the center of pressure aft on the airfoil, resulting in a nose-down pitching moment.

The split flap is deflected from the lower surface of the airfoil and produces a slightly greater increase in lift than does the plain flap. However, more drag is created because of the turbulent air pattern produced behind the airfoil. When fully extended, both plain and split flaps produce high drag with little additional lift.

The most popular flap on airplanes today is the slotted flap. Variations of this design are used for small airplanes as well as for large ones. Slotted flaps increase the lift coefficient significantly more than plain or spilt flaps. On small airplanes, the hinge is located below the lower surface of the flap, and when the flap is lowered, it forms a duct between the flap well in the wing and the leading edge of the flap.

When the slotted flap is lowered, high-energy air from the lower surface is ducted to the flap's upper surface. The high-energy air from the slot accelerates the upper surface boundary layer and delays airflow separation, providing a higher coefficient of lift. Thus, the slotted flap produces much greater increases in CLmax than the plain or split flap. While there are many types of slotted flaps, large airplanes often have double- and even triple-slotted flaps. These allow the maximum increase in drag without the airflow over the flaps separating and destroying the lift they produce.

Fowler flaps are a type of slotted flap. This flap design not only changes the camber of the wing, it also increases the wing area. Instead of rotating down on a hinge, it slides backwards on tracks. In the first portion of its extension, it increases the drag very little, but increases the lift a great deal as it increases both the area and camber. As the extension continues, the flap deflects downward, and during the last portion of its travel, it increases the drag with little additional increase in lift.

FLIGHT CONTROLS - STABILIZER

As mentioned earlier, a stabilizer is essentially a one-piece horizontal stabilizer with the same type of control system. Because stabilizers pivot around a central hinge point, they are extremely sensitive to control inputs and aerodynamic loads.

Antiservo tabs are incorporated on the trailing edge to decrease sensitivity. In addition, a balance weight is usually incorporated ahead of the main spar. The balance weight may project into the empennage or may be incorporated on the forward portion of the stabilizer tips.
 
When the control column is pulled back, it raises the stabilizer's trailing edge, rotating the airplane's nose up. Pushing the control column forward lowers the trailing edge of the stabilizer and pitches the nose of the airplane down. Without an anti-servo tab, the airplane would be prone to over controlling from pilot-induced control inputs.

FLIGHT CONTROLS - V-TAIL

The V-tail design utilizes two slanted tail surfaces to perform the same functions as the surfaces of a conventional elevator and rudder configuration. The fixed surfaces act as both horizontal and vertical stabilizers.

The movable surfaces, which are usually called ruddervators, are connected through a special linkage that allows the control wheel to move both surfaces simultaneously. On the other hand, displacement of the rudder pedals moves the surfaces differentially, thereby providing directional control. When both rudder and elevator controls are moved by the pilot, a control mixing mechanism moves each surface the appropriate amount. The control system for the V-tail is more complex than that required for a conventional tail. In addition, the V-tail design is more susceptible to Dutch roll tendencies than a conventional tail and total reduction in drag is only minimal.

Ruddervator—A pair of control surfaces on the tail of an aircraft arranged in the form of a V. These surfaces, when moved together by the control wheel, serve as elevators, and when moved differentially by the rudder pedals, serve as a rudder.

FLIGHT CONTROLS - T-TAIL

In a T-tail configuration, the elevator is above most of the effects of downwash from the propeller as well as airflow around the fuselage and/or wings during normal flight conditions. Operation of the elevators in this undisturbed air makes for control movements that are consistent throughout most flight regimes. T-tail designs have become popular on many light airplanes and on large aircraft, especially those with aft-fuselage mounted engines since the T-tail configuration removes the tail from the exhaust blast of the engines. Seaplanes and amphibians often have T-tails in order to keep the horizontal surfaces as far from the water as possible. An additional benefit is reduced vibration and noise inside the aircraft.

At slow speeds, the elevator on a T-tail aircraft must be moved through a larger number of degrees of travel to raise the nose a given amount as compared to a conventional-tail aircraft. This is because the conventional-tail aircraft has the downwash from the propeller pushing down on the tail to assist in raising the nose. Since controls on aircraft are rigged in such a manner as to require increasing control forces for increased control travel, the forces required to raise the nose of a T-tail aircraft are greater than for a conventional-tail aircraft. Longitudinal stability of a trimmed aircraft is the same for both types of configuration, but the pilot must be aware that at slow speeds during takeoffs and landings or stalls, the control forces will be greater than for similar size airplanes equipped with conventional tails.

T-tail airplanes also require additional design considerations to counter the problem of flutter. Since the weight of the horizontal surfaces is at the top of the vertical stabilizer, the moment arm created causes high loads on the vertical stabilizer which can result in flutter. Engineers must compensate for this by increasing the design stiffness of the vertical stabilizer, usually resulting in a weight penalty over conventional tail designs.

When flying at a very high angle of attack with a low airspeed and an aft CG, the T-tail airplane may be susceptible to a deep stall. In a deep stall, the airflow over the horizontal tail is blanketed by the disturbed airflow from the wings and fuselage. In these circumstances, elevator or stabilator control could be diminished, making it difficult to recover from the stall. It should be noted that an aft CG could be a contributing factor in these incidents since similar recovery problems are also found with conventional-tail aircraft with an aft CG.

Since flight at a high angle of attack with a low airspeed and an aft CG position can be dangerous, many airplanes have systems to compensate for this situation. The systems range from control stops to elevator down springs. An elevator down spring assists in lowering the nose to prevent a stall caused by the aft CG position. The stall occurs because the properly trimmed airplane is flying with the elevator in a trailing edge down position, forcing the tail up and the nose down. In this unstable condition, if the airplane encounters turbulence and slows down further, the trim tab no longer positions the elevator in the nose-down position. The elevator then streamlines, and the nose of the aircraft pitches upward. This aggravates the situation and can possibly result in a stall.

The elevator down spring produces a mechanical load on the elevator, causing it to move toward the nosedown position if not otherwise balanced. The elevator trim tab balances the elevator down spring to position the elevator in a trimmed position. When the trim tab becomes ineffective, the down spring drives the elevator to a nose down position. The nose of the aircraft lowers, speed builds up, and a stall is prevented.

The elevator must also have sufficient authority to hold the nose of the airplane up during the roundout for a landing. In this case, a forward CG may cause a problem. During the landing flare, power normally is reduced, which decreases the airflow over the empennage. This, coupled with the reduced landing speed, makes the elevator less effective.

From this discussion, it should be apparent that pilots must understand and follow proper loading procedures, particularly with regard to the CG position.

FLIGHT CONTROLS - RUDDER

The rudder controls movement of the airplane about its vertical axis. This motion is called yaw. Like the other primary control surfaces, the rudder is a movable surface hinged to a fixed surface, in this case, to the vertical stabilizer, or fin. Moving the left or right rudder pedal controls the rudder. When the rudder is deflected into the airflow, a horizontal force is exerted in the opposite direction.

By pushing the left pedal, the rudder moves left. This alters the airflow around the vertical stabilizer/rudder, and creates a sideward lift that moves the tail to the right and yaws the nose of the airplane to the left. Rudder effectiveness increases with speed, so large deflections at low speeds and small deflections at high speeds may be required to provide the desired reaction. In propeller-driven aircraft, any slipstream flowing over the rudder increases its effectiveness.

FLIGHT CONTROLS - ELEVATOR

The elevator controls pitch about the lateral axis. Like the ailerons on small airplanes, the elevator is connected to the control column in the cockpit by a series of mechanical linkages. Aft movement of the control column deflects the trailing edge of the elevator surface up. This is usually referred to as up elevator.

The up-elevator position decreases the camber of the elevator and creates a downward aerodynamic force, which is greater than the normal tail-down force that exists in straight-and-level flight. The overall effect causes the tail of the airplane to move down and the nose to pitch up. The pitching moment occurs about the center of gravity (CG). The strength of the pitching moment is determined by the distance between the CG and the horizontal tail surface, as well as by the aerodynamic effectiveness of the horizontal tail surface.

Moving the control column forward has the opposite effect. In this case, elevator camber increases, creating more lift (less tail-down force) on the horizontal stabilizer/elevator. This moves the tail upward and pitches the nose down. Again, the pitching moment occurs about the CG.

As mentioned earlier in the coverage on stability, power, thrust line, and the position of the horizontal tail surfaces on the empennage are factors in how effective the elevator is in controlling pitch. For example, the horizontal tail surfaces may be attached near the lower part of the vertical stabilizer, at the midpoint, or at the high point, as in the T-tail design.

FLIGHT CONTROLS - COUPLED AILERONS AND RUDDER

Coupled ailerons and rudder means these controls are linked. This is accomplished with rudder-aileron interconnect springs, which help correct for aileron drag by automatically deflecting the rudder at the same time the ailerons are deflected. For example, when the control yoke is moved to produce a left roll, the interconnect cable and spring pulls forward on the left rudder pedal just enough to prevent the nose of the airplane from yawing to the right. The force applied to the rudder by the springs can be overridden if it becomes necessary to slip the airplane.

FLIGHT CONTROLS - FRISE-TYPE AILERONS

With a Frise-type aileron, when pressure is applied to the control wheel, the aileron that is being raised pivots on an offset hinge. This projects the leading edge of the aileron into the airflow and creates drag. This helps equalize the drag created by the lowered aileron on the opposite wing and reduces adverse yaw.

The Frise-type aileron also forms a slot so that air flows smoothly over the lowered aileron, making it more effective at high angles of attack. Frise-type ailerons also may be designed to function differentially. Like the differential aileron, the Frise-type aileron does not eliminate adverse yaw entirely. Coordinated rudder application is still needed wherever ailerons are applied.

FLIGHT CONTROLS - DIFFERENTIAL AILERONS

With differential ailerons, one aileron is raised a greater distance than the other aileron is lowered for a given movement of the control wheel. This produces an increase in drag on the descending wing. The greater drag results from deflecting the up aileron on the descending wing to a greater angle than the down aileron on the rising wing. While adverse yaw is reduced, it is not eliminated completely.

FLIGHT CONTROLS - ADVERSE YAW

Since the downward deflected aileron produces more lift, it also produces more drag. This added drag attempts to yaws the airplane's nose in the direction of the raised wing. This is called adverse yaw.

The rudder is used to counteract adverse yaw, and the amount of rudder control required is greatest at low airspeeds, high angles of attack, and with large aileron deflections. However, with lower airspeeds, the vertical stabilizer/rudder combination becomes less effective, and magnifies the control problems associated with adverse yaw.

All turns are coordinated by use of ailerons, rudder, and elevator. Applying aileron pressure is necessary to place the airplane in the desired angle of bank, while simultaneously applying rudder pressure to counteract the resultant adverse yaw. During a turn, applying elevator pressure because more lift is required than when in straight-and level flight must increase the angle of attack. The steeper the turn, the more back elevator pressure is needed.

As the desired angle of bank is established, aileron and rudder pressures should be relaxed. This will stop the bank from increasing because the aileron and rudder control surfaces will be neutral in their streamlined position. Elevator pressure should be held constant to maintain a constant altitude.

The rollout from a turn is similar to the roll-in the flight controls are applied in the opposite direction. Aileron and rudder are applied in the direction rollout or toward the high wing. As the angle decreases, the elevator pressure should be relaxed necessary to maintain altitude.

FLIGHT CONTROLS - AILERONS

Ailerons control roll about the longitudinal axis. The ailerons are attached to the outboard trailing edge of each wing and move in the opposite direction from each other. Ailerons are connected by cables, bell cranks, pulleys or push-pull tubes to each other and to the control wheel.

Moving the control wheel to the right causes the right aileron to deflect upward and the left aileron to deflect downward. The upward deflection of the right aileron decreases the camber resulting in decreased lift on the right wing. The corresponding downward deflection of the left aileron increases the camber resulting in increased lift on the left wing. Thus, the increased lift on the left wing and the decreased lift on the right wing cause the airplane to roll to the right.

FLIGHT CONTROLS - PRIMARY FLIGHT CONTROLS

Aircraft flight control systems are classified as primary and secondary. The primary control systems consist of those that are required to safely control an airplane during flight. These include the ailerons, elevator (or stabilizer), and rudder. Secondary control systems improve the performance characteristics of the airplane, or relieve the pilot of excessive control forces. Examples of secondary control systems are wing flaps and trim systems.

Airplane control systems are carefully designed to provide a natural feel, and at the same time, allow adequate responsiveness to control inputs. At low airspeeds, the controls usually feel soft and sluggish, and the airplane responds slowly to control applications. At high speeds, the controls feel firm and the response is more rapid.

Movement of any of the three primary flight control surfaces changes the airflow and pressure distribution over and around the airfoil. These changes affect the lift and drags produced by the airfoil/control surface combination, and allow a pilot to control the airplane about its three axes of rotation.

Design features limit the amount of deflection of flight control surfaces. For example, control-stop mechanisms may be incorporated into the flight controls or movement of the control column and/or rudder pedals may be limited. The purpose of these design limits is to prevent the pilot from inadvertently over controlling and over stressing the aircraft during normal maneuvers.

A properly designed airplane should be stable and easily controlled during maneuvering. Control surface inputs cause movement about the three axes of rotation. The types of stability an airplane exhibits also relate to the three axes of rotation.

HIGH-SPEED FLIGHT- FLIGHT CONTROLS

On high-speed airplanes, flight controls are divided into primary flight controls and secondary or auxiliary flight controls. The primary flight controls maneuver the airplane 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 airplanes, 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 airplane from the wings onto the wheels for better braking performance.

Jet transport airplanes have small ailerons. The space for ailerons is limited because as much of the wing trailing edge as possible is needed for flaps. Another reason is that a conventional size aileron would cause wing twist at high speed. Because the ailerons are necessarily small, 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 types, they will extend further on one side and retract on the other side.

If they are the Non-Differential type, they will extend further on one side but will 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 angle of attack without airflow separation, an airplane's wing leading edge should have a well-rounded almost blunt shape that the airflow can adhere to at the higher angle of attack. With this shape, the airflow separation will start at the trailing edge and progress forward gradually as angle of attack 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 angle of attack. To utilize trailing edge flaps, and thus increase the maximum lift coefficient, the wing must go to a higher angle of attack 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 maximum lift coefficient (CLmax) is increased considerably. In fact, a 50-knot reduction in stall speed is not uncommon.

The operational requirements of a large jet transport airplane necessitate large pitch trim changes. Some of these requirements are:
- The requirement for a large CG range.
- The need to cover a large speed range.
- The need to cope with possibly large trim changes due to wing leading edge and trailing edge high- lift devices without limiting the amount of elevator remaining.
- The need to reduce 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 airplane 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 airplanes equipped with a variable incidence horizontal stabilizer, the elevator is smaller and less effective in isolation than it is on a fixed-tail airplane. In comparison to other flight controls, the variable incidence horizontal stabilizer is enormously powerful in its effect. Its use and effect must be fully understood and appreciated by flight crewmembers.

Because of the size and high speeds of jet transport airplanes, the forces required moving the control surfaces can be beyond the strength of the pilot. Consequently, hydraulic or electrical power units actuate the control surfaces. Moving the controls in the cockpit 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 that causes the control surface to move.

HIGH-SPEED FLIGHT- MACH BUFFET BOUNDARIES

Thus far, only the Mach buffet that results from excessive speed has been addressed. It must be remembered that Mach buffet is a function of the speed of the airflow over the wing—not necessarily the speed of the airplane. Any time that too great a lift demand is made on the wing, whether from too fast an airspeed or from too high an angle of attack near the MMO, the "high-speed" buffet will occur.

However, there are also occasions when the buffet can be experienced at much lower speeds known as the "low-speed Mach buffet." The most likely situation that could cause the low speed buffet would be when the airplane is flown at too slow a speed for its weight and altitude necessitating a high angle of attack. This very high angle of attack would have the effect of increasing airflow velocity over the upper surface of the wing to the point that all of the same effects of the shock waves and buffet would occur as in the high-speed buffet situation. The angle of attack of the wing has the greatest effect on inducing the Mach buffet at either the high-speed or low-speed boundaries for the airplane. The conditions that increase the angle of attack, hence the speed of the airflow over the wing and chances of Mach buffet are as follows:

- High Altitudes—The higher an airplane flies, the thinner the air and the greater the angle of attack required to produce the lift needed to maintain level flight.
- Heavy Weights—The heavier the airplane, the greater the lift required of the wing, and all other things being equal, the greater the angle of attack.
- "G" Loading—An increase in the "G" loading on the airplane has the same effect as increasing the weight of the airplane. Whether turns, rough control usage, or turbulence causes the increase in "G" forces, the effect of increasing the wing's angle of attack is the same.

HIGH-SPEED FLIGHT- SWEEPBACK

Most of the difficulties of transonic flight are associated with shock wave induced flow separation. Therefore, any means of delaying or alleviating the shock-induced separation will improve aerodynamic performance. One method is wing sweep back. Sweep back 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.

On a straight wing airplane, the airflow strikes the wing leading edge at 90°, and its full impact produces pressure and lift. A wing with sweep back 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 rise will peak. In other words, sweep will delay the onset of compressibility effects. The Mach number, which produces a sharp change in drag coefficient, is termed the "force divergence" Mach numbers 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, sweep back reduces the magnitude in the changes of drag, lift or moment coefficients. In other words, the use of sweep back will "soften" the force divergence. A disadvantage of swept wings is that they tend to stall at the wingtips rather than at the wing roots.

This is because the boundary layer tends flows span wise 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 center of lift), a wingtip stall will cause the center of lift to move forward on the wing, forcing the nose to rise further. The tendency for tip stall is greatest when wing sweep and taper are combined.

The stall situation can be aggravated by a T-tail configuration, which affords little or no pre-stall warning in the form of tail control surface buffet. 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 angle of attack.

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 airplane (specifically the enormous increase in drag at low speeds) can cause an increasingly descending flight path with no change in pitch attitude, further increasing the angle of attack. In this situation, without reliable angle of attack 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 airplane stalled.

It is a characteristic of T-tail airplanes 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, preprogrammed stick forces automatically move the stick forward, preventing the stall from developing. A "g" limited may also be incorporated into the system to prevent the pitch down generated by the stick pusher from imposing excessive loads on the airplane. A "stick shaker," on the other hand provides stall warning when the airspeed is 5 to 7 percent above stall speed.

HIGH-SPEED FLIGHT- SHOCK WAVES

When an airplane flies at subsonic speeds, the air ahead is "warned" of the airplane's coming by a pressure change transmitted ahead of the airplane at the speed of sound. Because of this warning, the air begins to move aside before the airplane arrives and is prepared to let it pass easily. When the airplane's speed reaches the speed of sound, the pressure change can no longer warn the air ahead because the airplane is keeping up with its own pressure waves. Rather, the air particles pile up in front of the airplane causing a sharp decrease in the flow velocity directly in front of the airplane with a corresponding increase in air pressure and density.

As the airplane's speed increases beyond the speed of sound, the pressure and density of the compressed air ahead of it increase the area of compression extending some distance ahead of the airplane. At some point in the air stream, 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 air stream is slowed to subsonic without a change in direction. Such as when the air stream is accelerated to sonic speed over the chambered portion of a wing, and then decelerates to subsonic speed as the area of maximum camber is passed. A shock wave will form 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 air stream passing through a normal shock wave will experience these changes:

- The air stream is slowed to subsonic.
- The airflow immediately behind the shock wave does not change direction.
- The static pressure and density of the air stream behind the wave is greatly increased.
- The energy of the air stream (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 first. Further increases in Mach number, however, can enlarge the supersonic area on the 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.

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 down wash, 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 down wash behind the wing. Thus, a decrease in down wash decreases the horizontal tail's pitch control effectiveness. Movement of the wing center of pressure affects the wing pitching moment. If the center of pressure 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 airplanes, which places the horizontal stabilizer as far as practical from the turbulence of the wings.

HIGH-SPEED FLIGHT- BOUNDARY LAYER

Air has viscosity, and will encounter resistance to flow over a surface. The viscous nature of airflow reduces the local velocity's on a surface and is responsible for skin friction drag. As the air passes over the wing's surface, the air particles nearest the surface come to rest. The next layer of particles is slowed down but not stopped. Some small but measurable distance from the surface, the air particles are moving at free stream velocity. The layer of air over the wing's surface, which is slowed down or stopped by viscosity, is termed the "boundary layer." Typical boundary layer thickness on an airplane range from small fractions of an inch near the leading edge of a wing to the order of 12 inches at the aft end of a large airplane such as a Boeing 747.

There are two different types of boundary layer flow: laminar and turbulent. 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. 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. 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 angle of attack is increased.

"Vortex Generators" are used to delay or prevent shock wave induced boundary layer separation encountered in transonic flight. Vortex generators are small low aspect ratio airfoils placed at a 12° to 15° angle of attack to the air stream. They are usually spaced a few inches apart along the wing ahead of the ailerons or other control surfaces. Vortex generators create a vortex that mixes the boundary airflow with the high-energy airflow just above the surface. This produces higher surface velocity's and increases the energy of the boundary layer. Thus, a stronger shock wave will be necessary to produce airflow separation.

HIGH-SPEED FLIGHT- MACH NUMBER VS AIRSPEED

Speeds such as Mach Crit and MMO for a specific airplane occur at a given Mach number. The true airspeed (TAS), however, varies with outside air temperature. Therefore, true airspeeds corresponding to a specific Mach number can vary considerably (as much as 75 – 100 knots). When an airplane cruising at a constant Mach number enters an area of higher outside air temperatures, true airspeed and required fuel increases, and range decreases. Conversely, when entering an area of colder outside air temperatures, true airspeed and fuel flow decreases, and range increases. In a jet airplane operating at high altitude, the indicated airspeed (IAS) for any given Mach number decreases with an increase in altitude above a certain level. The reverse occurs during descent. Normally, climbs and descents are accomplished using indicated airspeed in the lower altitudes and Mach number in the higher altitudes.

Unlike operations in the lower altitudes, the indicated airspeed (IAS) at which a jet airplane stalls increases significantly with altitude. This is due to the fact that true airspeed (TAS) increase with altitude. At high true airspeeds, air compression causes airflow distortion over the wings and in the pitot system. At the same time, the indicated airspeed (IAS) representing MMO decreases with altitude. Eventually, the airplane can reach an altitude where there is little or no difference between the two.

HIGH-SPEED FLIGHT- SPEED RANGES

The speed of sound varies with temperature. Under standard temperature conditions of 15°C, the speed of sound at sea level is 661 knots. At 40,000 feet, where the temperature is –55°C, the speed of sound decreases to 574 knots. In 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 airplane to the speed of sound in the same atmospheric conditions. An airplane traveling at the speed of sound is traveling at Mach 1.0. Airplane speed regimes are defined as follows:

Subsonic—Mach numbers below 0.75
Transonic—Mach numbers from .075 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 airplanes, civilian jet airplanes normally operate in a cruise speed range of Mach 0.78 to Mach 0.90. The speed of an airplane in which airflow over any part of the wing first reaches (but does not exceed) Mach 1.0 is termed that airplane's critical Mach number or "Mach Crit". Thus, critical Mach number is the boundary between subsonic and transonic flight and is an important point of reference for all compressibility effects encountered in transonic flight. Shock waves, buffet, and airflow separation take place above critical Mach numbers. A jet airplane 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," and is typically the speed chosen for high-speed cruise operations. At some point beyond high-speed cruise are the turbine powered airplane's maximum operating limit speeds: VMO/MMO.

VMO is the maximum operating speed expressed in terms of knots. VMO limits ram air pressure acting against the structure and prevents flutter. MMO is the maximum operating speed expressed in terms of Mach number. An airplane should not be flown in excess of this speed. Doing so risks encountering the full effects of compressibility, including possible loss of control.

HIGH-SPEED FLIGHT-SUPERSONIC VS SUBSONIC 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 below about 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 airplane. 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 airplane itself may be flying subsonic. At some extreme angles of attack, in some airplanes, the speed of the air over the top surface of the wing may be double the airplane's speed. It is therefore entirely possible to have both supersonic and subsonic airflow on an airplane at the same time. When flow velocity's reach sonic speeds at some location on an airplane (such as the area of maximum camber on the wing), further acceleration will result 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.

EFFECT OF LOAD DISTRIBUTION

The effect of the position of the center of gravity on the load imposed on an airplane's wing in flight is not generally realized, although it may be very significant to climb and cruising performance.

Contrary to the beliefs of some pilots, an airplane with forward loading is "heavier" and consequently, slower than the same airplane with the center of gravity furthers aft. With forward loading, "nose-up" trim is required in most airplanes 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 angle of attack of the wing, which results in more drag and, in turn, produces a higher stalling speed.

With aft loading and "nose-down" trim, the tail surfaces will exert less down load, relieving the wing of that much wing loading and lift required maintaining altitude. The required angle of attack 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 would also result in instability. Consequently, modern airplanes are designed to require a down load on the tail for stability and controllability.

Remember that a zero indication on the trim tab control is not necessarily the same as "neutral trim" because of the force exerted by down wash from the wings and the fuselage on the tail surfaces.

The effects of the distribution of the airplane's useful load have a significant influence on its flight characteristics, even when the load is within the center- of-gravity 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 airplane becomes less controllable, especially at slow flight speeds, as the center of gravity is moved further aft. An airplane which cleanly recovers from a prolonged spin with the center of gravity at one position may fail completely to respond to normal recovery attempts when the center of gravity is moved aft by 1 or 2 inches. It is common practice for airplane designers to establish an aft center-of-gravity limit that is within 1 inch of the maximum which will allow normal recovery from a one-turn spin. When certifications an airplane in the utility category to permit intentional spins, the aft center-of-gravity limit is usually established at a point several inches forward that which is permissible for certification in the normal category.

Another factor affecting controllability, which is becoming more important in current designs of large airplanes, is the effect of long moment arms to the positions of heavy equipment and cargo. The same airplane may be loaded to maximum gross weight within its center-of-gravity limits by concentrating fuel, passengers, and cargo near the design center of gravity; 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 center of gravity, maneuvering the airplane or maintaining level flight in turbulent air will require the application of greater control forces when the load is dispersed. This is true because of the longer moment arms to the positions of the heavy fuel and cargo loads that must be overcome by the action of the control surfaces. An airplane 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 center-of-gravity limit of an airplane is determined largely by considerations of stability. The original airworthiness requirements for a type certificate specify that an airplane in flight at a certain speed will dampen out vertical displacement of the nose within a certain number of oscillations. An airplane loaded too far rearward may not do this; instead when the nose is momentarily pulled up, it may alternately climb and dive becomes steeper with each oscillation. This instability is not only uncomfortable to occupants, but it could even become dangerous by making the airplane unmanageable under certain conditions.

The recovery from a stall in any airplane becomes progressively more difficult as its center of gravity moves aft. This is particularly important in spin recovery, as there is a point in rearward loading of any airplane at which a "flat" spin will develop. A flat spin is one in which centrifugal force, acting through a center of gravity located well to the rear, will pull the tail of the airplane out away from the axis of the spin, making it impossible to get the nose down and recover.

An airplane loaded to the rear limit of its permissible center-of-gravity range will handle differently in turns and stall maneuvers and has different landing characteristics than when it is loaded near the forward limit. The forward center-of-gravity 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 airplane in a normal glide with the power off. A conventional airplane must be capable of a full stall, power-off landing in order to ensure minimum landing speed in emergencies. A tail wheel-type airplane loaded excessively nose heavy will be difficult to taxi, particularly in high winds. It can be nosed over easily by use of the brakes, and it will be 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 nose wheel-type airplanes, particularly during the landing roll and takeoff.

- The CG position influences the lift and angle of attack 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 airplane will stall at a higher speed with a forward CG location. This is because the stalling angle of attack 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 airplane.
- The airplane will cruise faster with an aft CG location because of reduced drag. The drag is reduced because a smaller angle of attack and less downward deflection of the stabilizer are required to support the airplane and overcome the nose-down pitching tendency.
- The airplane becomes less stable as the CG is moved rearward. This is because when the CG is moved rearward it causes an increase in the angle of attack. Therefore, the wing contribution to the airplane'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 will result in an unstable airplane.
- 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 airplane throughout the airspeed range down to the stall.

EFFECTS OF WEIGHT ON STABILITY AND CONTROL ABILITY

The effects that overloading has on stability also are not generally recognized. An airplane, which is observed to be quite stable and controllable when loaded normally, may be discovered to have very different flight characteristics when it is overloaded.

Although the distribution of weight has the most direct effect on this, an increase in the airplane's
gross weight may be expected to have an adverse effect on stability, regardless of location of the center of gravity. The stability of many certificate airplanes is completely unsatisfactory if the gross weight is exceeded.

EFFECT OF WEIGHT ON AIRPLANE STRUCTURE

The effect of additional weight on the wing structure of an airplane is not readily apparent. Airworthiness requirements prescribe that the structure of an airplane certificate in the normal category (in which acrobatics are prohibited) must be strong enough to withstand a load factor of 3.8 to take care of dynamic loads caused by maneuvering and gusts. This means that the primary structure of the airplane can withstand a load of 3.8 times the approved gross weight of the airplane without structural failure occurring. If this is accepted as indicative of the load factors that may be imposed during operations for which the airplane 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 airplanes, 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 which is difficult to detect and expensive to repair. One of the most serious results of habitual overloading is that its results tend to be cumulative, and may 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.

Knowledge of load factors imposed by flight maneuvers and gusts will emphasize the consequences of an increase in the gross weight of an airplane. The structure of an airplane about to undergo a load factor of 3 G's, as in the 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 airplane. The FAA certificate civil airplane has 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 airplane. Nonetheless, this fact should not be allowed to mislead the pilot, as the pilot may not realize that loads for which the airplane was not designed are being imposed on all or some part of the structure.

In loading an airplane 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. As an example, a light plane baggage compartment may be placarded for 20 pounds because of the limited strength of its supporting structure even though the airplane may not be overloaded or out of center-of-gravity limits with more weight at that location.

EFFECTS OF WEIGHT ON FLIGHT PERFORMANCE

The takeoff/climb and landing performance of an airplane are determined on the basis of its maximum allowable takeoff and landing weights. A heavier gross weight will result 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 airplane to clear an obstacle that normally would not have been seriously considered during takeoffs under more favorable conditions.

The detrimental effects of overloading on performance are not limited to the immediate hazards involving takeoffs and landings. Overloading has an adverse effect on all climb and cruise performance that leads to overheating during climbs added wear on engine parts, increased fuel consumption, slower cruising speeds, and reduced range.

The manufacturers of modern airplanes furnish weight and balance data with each airplane produced. Generally, this information may be found in the FAA approved Airplane Flight Manual or Pilot's Operating Handbook (AFM/POH). With the advancements in airplane design and construction in recent years has come the development of "easy to read charts" for determining weight and balance data. Increased performance and load carrying capability of these airplanes require strict adherence to the operating limitations prescribed by the manufacturer.

Deviations from the recommendations can result in structural damage or even complete failure of the airplane's structure. Even if an airplane is loaded well within the maximum weight limitations, it is imperative that weight distribution be within the limits of center of gravity 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 airplane. The pilot is often completely unaware of the weight and balance limitations of the airplane being flown and of the reasons for these limitations. In some airplanes, it is not possible to fill all seats, baggage compartments, and fuel tanks, and still remain within approved weight or balance limits. As an example, in several popular four-place airplanes the fuel tanks may not be filled to capacity when four occupants and their baggage are carried. In a certain two-place airplane, no baggage may be carried in the compartment aft of the seats when spins are to be practiced.

WEIGHT AND BALANCE

Often a pilot regards the airplane's weight and balance data as information of interest only to engineers, dispatchers, and operators of scheduled and nonscheduled air carriers. Along with this idea, the reasoning is that the airplane was weighed during the certification process and that this data is valid indefinitely, regardless of equipment changes or modifications. Further, this information is mistakenly reduced to a workable routine or "rule of thumb" such as: "If I have three passengers, I can load only 100 gallons of fuel; four passengers—70 gallons."

Admittedly, this rule of thumb is adequate in many cases, but as the subject "Weight and Balance" suggests, the concern is not only with the weight of the airplane but also the location of its center of gravity (CG). The importance of the CG should have become apparent in the discussion of stability, control ability, and performance. If all pilots understood and respected the effect of CG on an airplane, then one type of accident would be eliminated from the records: "PRIMARY CAUSE OF ACCIDENT— AIRPLANE CENTER OF GRAVITY OUT OF REARWARD LIMITS AND UNEQUAL LOAD DISTRIBUTION RESULTING IN AN UNSTABLE AIRPLANE. PILOT LOST CONTROL OF AIRPLANE ON TAKEOFF AND CRASHED."

The reason airplanes are so certificates are obvious when one gives it a little thought. For instance, it is of added value to the pilot to be able to carry extra fuel for extended flights when the full complement of passengers is not to be carried. Further, it is unreasonable to forbid the carriage of baggage when it is only during spins that its weight will adversely affect the airplane's flight characteristics. Weight and balance limits are placed on airplanes for two principal reasons:

1. Because of the effect of the weight on the airplane's primary structure and its performance characteristics; and
2. Because of the effect the location of this weight has on flight characteristics, particularly in stall and spin recovery and stability.

LOAD FACTORS AND FLIGHT MANEUVERS - VG DIAGRAM

The flight operating strength of an airplane is presented on a graph whose horizontal scale is based on load factor. The diagram is called a Vg diagram—velocity versus "g" loads or load factor.

Each airplane has its own Vg diagram that is valid at a certain weight and altitude. The lines of maximum lift capability (curved lines) are the first items of importance on the Vg diagram. The subject airplane in the illustration is capable of developing no more than one positive "g" at 62 m.p.h., the wing level stall speed of the airplane.

Since the maximum load factor varies with the square of the airspeed, the maximum positive lift capability of this airplane is 2 "g" at 92 m.p.h., 3 "g" at 112 m.p.h., 4.4 "g" at 137 m.p.h., and so forth. Any load factor above this line is unavailable aerodynamically; i.e., the subject airplane cannot fly above the line of maximum lift capability (it will stall). Essentially 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 subject airplane is flown at a positive load factor greater than the positive limit load factor of 4.4, structural damage will be possible. When the airplane 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. First, 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 airplane; any airspeed less 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 to occur at this condition. The maneuver speed is a valuable reference point, since an airplane 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 airplane is below the maneuver speed.

Next, 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 airplane; any airspeed less does not provide negative lift capability sufficient to damage the airplane from excessive flight loads.

The limit airspeed (or redline speed) is a design reference point for the airplane—the subject airplane is limited to 225 m.p.h. If flight is attempted beyond the limit airspeed, structural damage or structural failure may result from a variety of phenomena. Thus, the airplane in flight is limited to a regime of airspeeds and g's which do not exceed the limit (or redline) speed, do not exceed the limit load factor, and cannot exceed the maximum lift capability. The airplane must be operated within this "envelope" to prevent structural damage and ensure that the anticipated service lift of the airplane 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 airplane.

LOAD FACTORS AND FLIGHT MANEUVERS - ROUGH AIR

All certificate airplanes 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 that exceed the load limits.
Most airplane flight manuals now include turbulent air penetration information. Operators of modern airplanes, capable of a wide range of speeds and altitudes, are benefited by this added feature both in comfort and safety. In this connection, it is to be noted 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.

In summary, it must be remembered that load factors induced by intentional acrobatics, abrupt pull-ups from dives, high-speed stalls, and gusts at high airspeeds all place added stress on the entire structure of an airplane. Stress on the structure involves forces on any part of the airplane. There is a tendency for the uninformed to think of load factors only in terms of their effect on spars and struts. Most structural failures due to excess load factors involve rib structure within the leading and trailing edges of wings and tail group. The critical area of fabric-covered airplanes is the covering about one-third of the chord aft on the top surface of the wing.

The cumulative effect of such loads over a long period of time may tend to loosen and weaken vital parts so that actual failure may occur later when the airplane is being operated in a normal manner.

LOAD FACTORS AND FLIGHT MANEUVERS - CHANDELLES AND LAZY EIGHTS

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. Generally, the better the maneuver is performed, the less extreme will be the load factor induced. A chandelle or lazy eight, in which the pull-up produces a load factor greater than 2 G's will not result in as great a gain in altitude, and in low-powered airplanes it may result in a net loss of altitude.

The smoothest pull-up possible, with a moderate load factor, will deliver the greatest gain in altitude in a chandelle and will result in a better overall performance in both chandelles and lazy eight's.

Further, it will be noted that recommended entry speed for these maneuvers be generally near the manufacturer's design maneuvering speed, thereby allowing maximum development of load factors without exceeding the load limits.

LOAD FACTORS AND FLIGHT MANEUVERS - 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 airplanes.

The only way this stall can be induced at 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 airplane with a stalling speed of 60 knots) will produce a load factor of 3 G's. Further, only a very narrow margin for error can be allowed for acrobatics in light airplanes. To illustrate how rapidly the load factor increases with airspeed, a high-speed stall at 112 knots in the same airplane would produce a load factor of 4 G's.

LOAD FACTORS AND FLIGHT MANEUVERS - SPINS

Since a stabilized spin is not essentially different from a stall in any element other than rotation, the same load factor considerations apply as those that apply to stall recovery. Since spin recoveries usually are 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 will usually be found to be about 2.5 G's.
The load factor during a spin will vary with the spin characteristics of each airplane but is usually found to be slightly above the 1 G of level flight. There are two reasons this is true:

1. The airspeed in a spin is very low, usually within 2 knots of the unaccelerated stalling speeds; and
2. The airplane pivots, rather than turns, while it is in a spin.

LOAD FACTORS AND FLIGHT MANEUVERS - STALLS

The normal stall entered from straight level flight, or an unaccelerated straight climb, will 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; and the pilot has the feeling of "floating free in space." In the event snapping the elevator control forward effects recovery, negative load factors, 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 sometimes are induced. Inadvertently these may be further increased 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 airplane structures and may produce recurrent or secondary stalls by increasing the angle of attack 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 G's. A higher load factor should never be necessary unless recovery has been effected with the airplane's nose near or beyond the vertical attitude, or at extremely low altitudes to avoid diving into the ground.