AVAV4107 Flight Controls Handout PDF

AVAV4107 Flight Controls Handout AVAV 4107 FLIGHT CONTROLS Handout The material contained in this manual is for TRAINING PURPOSES ONLY Page 1 of 66...

AVAV4107 Flight Controls Handout AVAV 4107 FLIGHT CONTROLS Handout The material contained in this manual is for TRAINING PURPOSES ONLY Page 1 of 66 09 June, 2016  AVAV4107 Flight Controls Handout TABLE OF CONTENTS Introduction 3 System operation 5 Mechanical assistance 7 Flight Control Systems 9 Primary flight controls 10 Elevator 10 Aileron 13 Rudder 18 Secondary flight controls 20 Hydro-mechanical controls 32 Artificial feel devices 34 Full power controls 35 Pneumatic powered controls 36 Electrically powered controls 36 Fly by wire controls 36 Analog fly by wire controls 41 Digital fly by wire controls 42 Supplemental Flight Control Systems 47 Yaw Damper 47 Mach/Speed Trim 49 Rudder Limiter 51 Gust Lock 52 Stall warning 53 Maintenance practices 55 Balancing 55 Rigging 58 Page 2 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Flight Controls H1 Aircraft axis Flight dynamics is the science of air vehicle orientation and control in three dimensions. The three critical flight dynamics parameters are the angles of rotation in three dimensions about the vehicle's center of mass, known as pitch, roll and yaw. Aircraft engineers develop control systems for a vehicle's orientation (attitude) about its center of mass. The control systems include actuators, which exert forces in various directions, and generate rotational forces or moments about the centre of gravity of the aircraft, and thus rotate the aircraft in pitch, roll, or yaw. For example, a pitching moment is a vertical force applied at a distance forward or aft from the centre of gravity of the aircraft, causing the aircraft to pitch up or down. Roll, pitch and yaw refer, in this context, to rotations about the respective axes starting from a defined equilibrium state. The equilibrium roll angle is referenced from wings level or zero bank angle. Yaw is known as "heading" without slipping or skidding i.e. coordinated. The equilibrium pitch angle in aircraft usually refers to angle of attack, rather than orientation. However, common usage ignores this distinction between equilibrium and dynamic cases. A fixed-wing aircraft increases or decreases the lift generated by the wings when it pitches nose up or down by increasing or decreasing the angle of attack (AOA). The roll angle is also known as bank angle on a fixed wing aircraft, which usually "banks" to change the horizontal direction of flight. However simply rolling the wings will not immediately cause an aircraft to change direction. The uncorrected input will begin a slip (uncoordinated turn) first. Because an aircraft is usually streamlined from nose to tail to reduce drag, it is advantageous to keep the sideslip angle near zero however; there are instances when an aircraft may be deliberately "side slipped". Page 3 of 66 09 June, 2016  AVAV4107 Flight Controls Handout ROLL PITCH Page 4 of 66 09 June, 2016  AVAV4107 Flight Controls Handout YAW An aircraft: Rolls ABOUT its longitudinal axis. Pitches ABOUT its lateral axis. Yaws ABOUT its vertical axis. System operation A conventional fixed-wing aircraft flight control system consists of flight control surfaces, the respective cockpit controls, connecting linkages, and the necessary operating mechanisms to control an aircraft's direction in flight. Aircraft engine controls are also considered as flight controls as they change speed. The fundamentals of aircraft controls are explained in flight dynamics. This section centers on the operating mechanisms of the flight controls. Page 5 of 66 09 June, 2016  AVAV4107 Flight Controls Handout De Havilland Tiger Moth elevator and rudder cables Mechanical or manually-operated flight control systems are the most basic method of controlling an aircraft. They were used in early aircraft and are currently used in small aircraft where the aerodynamic forces are not excessive. Very early aircraft, such as the Wright Flyer I, Blériot XI and Fokker Eindecker used a system of wing warping where no conventionally hinged control surfaces were used on the wing, and sometimes not even for pitch control as on the Wright Flyer I and original versions of the 1909 Etrich Taube, which only had a hinged/pivoting rudder in addition to the warping-operated pitch and roll controls. A manual flight control system uses a collection of mechanical parts such as pushrods, tension cables, pulleys, counterweights, and sometimes chains to transmit the forces applied to the cockpit controls directly to the control surfaces. Turnbuckles are often used to adjust controls and set cable tension. Increases in the control surface area required by large aircraft or higher loads caused by high airspeed aircraft lead to a large increase in the forces needed to move them. Consequently complicated mechanical gearing arrangements were developed to extract maximum mechanical advantage in order to reduce the forces required from the pilots. This arrangement can be found on bigger or higher performance propeller aircraft such as the Fokker 50. Page 6 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Fokker F50 Mechanical Assistance Servo Tab Some mechanical flight control systems use servo tabs that provide aerodynamic assistance. A servo tab (sometimes called a Flettner tab after its inventor Anton Flettner) is a small hinged device installed on an aircraft control surface to provide aerodynamic assist for the movement of the control surface. The flight control mechanisms move these tabs, aerodynamic forces in turn move, or assist the movement of the control surfaces reducing the amount of mechanical forces needed. This arrangement was used in early piston-engine transport aircraft and in early jet transports. Servo tabs move in the opposite direction of the control surface. The tab has a leverage advantage, being located closer to the trailing edge of the surface and thus can lever the control surface in the opposite direction. This has the effect of reducing the control force required by the pilot to move the controls. In the case of some aircraft the servo tab is the only control that is connected to the pilot's stick or wheel. The pilot moves the wheel which moves the servo tab and then the servo tab, using its mechanical advantage, moves the elevator or aileron, which is otherwise free-floating. The Boeing 737 incorporates a system, whereby in the event of total hydraulic system failure, it automatically and reverts to being controlled via servo-tab. Page 7 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Servo tab Bristol Britannia Anti-servo tabs An anti-servo tab works in the opposite way to a servo tab. It deploys in the same direction as the control surface, making the movement of the control surface more difficult and requires more force applied to the controls by the pilot. This may seem counter-productive, but it is commonly used on aircraft where the controls are too light or the aircraft requires additional stability in that axis of movement. The anti-servo tab serves to artificially increase stability and also make the controls heavier in feel to the pilot. Page 8 of 66 09 June, 2016  AVAV4107 Flight Controls Handout On some aircraft with all-flying stabilators an anti-servo tab acts as a trimming device. In this use some manufacturers term it a "balance tab" or "anti-balance tab". Anti-servo tab Flight Control Systems Stabilizer Stabilizers are either fixed or adjustable but are the “rigid” portion of the directional stability in this style of pitch control system. The fixed stabilizer is not adjustable. Aircraft with this style of horizontal stab must provide other forms of trim The adjustable horizontal stabilizer provides pitch (longitudinal) trim. The stabilizer is controlled by stabilizer trim which is adjusted either electrically (control switches located on the control horn of both pilots control wheels) or manually by rotating the stabilizer trim control wheels. Either stabilizer trim control method drives the stabilizer trim indicator. The trim actuator is a gearbox, jackscrew and ballnut assembly which drives the stabilizer front spar to provide stabilizer motion. The gearbox is driven by an electric actuator, an autopilot actuator, and a cable drum. Page 9 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Movable Horizontal Stabilizer Elevator Elevators are flight control surfaces, usually at the rear of an aircraft, which control the aircraft's orientation by changing the pitch of the aircraft, and so also the angle of attack of the wing. In simplified terms, they make the aircraft nose-up or nose-down (Ascending and descending are more a function of the wing). An increased wing angle of attack will cause a greater lift to be produced by the profile of the wing, and a slowing of the aircraft speed. A decrease in angle of attack will produce an increase in speed. Elevators are commonly one of two styles, firstly the conventional horizontal stabilizer (fixed or adjustable) and attached elevator or secondly a stabilator. The rear wing (stabilizer) to which elevators are attached have the opposite effect to a wing. They usually create a downward pressure which counters the unbalanced moment due to the airplane's center of gravity being located ahead of the center of pressure. An elevator decreases or Page 10 of 66 09 June, 2016  AVAV4107 Flight Controls Handout increases the downward force created by the rear wing. An increased downward force, produced by up elevator, forces the tail down and the nose up so lift increases (aircraft climbs) and the aircraft speed is reduced if power is not added (i.e. the wing will operate at a higher angle of attack, which produces a greater lift coefficient). A decreased downward force at the tail, produced by down elevator, allows the tail to rise and the nose to lower. The resulting lower wing angle of attack provides a lower lift coefficient, so the craft must move faster (either by adding power or going into a descent) to produce the required lift. The setting of the elevator thus determines the airplane's trim speed - a given elevator position has only one speed at which the aircraft will maintain a constant stable condition. Elevator (Pitch) Control In some aircraft pitch-control surfaces are in the front, ahead of the wing; this type of configuration is called a canard (the French word for duck). The Wright Brothers' early aircraft were of this type. Page 11 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Piaggio Avanti P180 Beechcraft Starship The canard type is more efficient, since the forward surface usually is required to produce upward lift (instead of downward force as with the usual empennage) to balance the net pitching moment. The main wing is also less likely to stall, as the forward control surface is configured to stall before the wing, causing a pitch down and reducing the angle of attack of the main wing. Stabilator A stabilator is a single piece horizontal surface that pivots about a point 1/3 aft of its leading edge (center of pressure). Because of the requirement for balance stabilators often have a balance weight on the end of a long tube connected to the front of the spar (pivot point) of the stabilator. Due to the ease with which they move and the large moving surface area, stabilators are often fitted with an anti-servo tab on the trailing edge to increase the control pressure to the pilot. Page 12 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Stabilator (Pitch) Control Aileron (Roll) Control Ailerons are hinged flight control surfaces attached to the trailing edge of the wing of a fixed-wing aircraft. The ailerons are used to control the aircraft in roll, which results in a change in heading due to the tilting of the lift vector. The two ailerons are typically interconnected so that one goes down when the other goes up: the down going aileron increases the lift on its wing while the up going aileron reduces the lift on its wing, producing a rolling moment about the aircraft's longitudinal axis. The word aileron is French for "little wing". Page 13 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Aileron Control System Frise Aileron An unwanted side effect of aileron operation is called adverse yaw—a yawing moment in the opposite direction to the roll. Using the ailerons to roll an aircraft to the right produces a yawing motion to the left. As the aircraft rolls, adverse yaw is caused primarily by the change in drag on the left and right wing. The rising wing generates increased lift which causes increased induced drag. The descending wing generates reduced lift which reduces induced drag. The difference in drag on each wing produces the Page 14 of 66 09 June, 2016  AVAV4107 Flight Controls Handout adverse yaw. There is also often an additional adverse yaw contribution from a difference in profile drag between the up-aileron and down-aileron. Adverse yaw is effectively compensated by the use of the rudder, which results in a side force on the vertical tail which opposes the adverse yaw by creating a favourable yawing moment. On some aircraft the rudder and aileron are interconnected by a spring, when the aileron is moved the spring pulls on the rudder. Another method of compensation is differential ailerons, which have been designed and engineered such that the down going aileron deflects less than the up going one. In this case the opposing yaw moment is generated by a difference in profile drag between the left and right wingtips. Frise ailerons accentuate this profile drag imbalance by protruding beneath the wing of an upward-deflected aileron, most often by being hinged slightly behind the leading edge and near the bottom of the surface. When the lower section of the leading edge protrudes slightly below the wings under surface as the aileron is deflected upwards, a substantial increase in profile drag occurs on that side. Ailerons may also be designed to use a combination of these methods. With ailerons in the neutral position the wing on the outside of the turn develops more lift than the opposite wing due to the variation in airspeed across the wing span, and this tends to cause the aircraft to continue to roll. Once the desired angle of bank (degree of rotation about the longitudinal axis) is obtained, the pilot uses opposite aileron to prevent the angle of bank from increasing due to this variation in lift across the wing span. This minor opposite use of the control must be maintained throughout the turn. The pilot also uses a slight amount of rudder in the same direction as the turn to counteract adverse yaw and to produce a "coordinated" turn where the fuselage is parallel to the flight path. A simple gauge on the instrument panel called the slip indicator, also known as "the ball", indicates when this coordination is achieved. Page 15 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Ailerons are the trailing-edge control surface nearest the wing tip (although on some airliners they can also be found at the wing root). On this parked Piper Cherokee, the left aileron has deflected downwards, and the right, upwards. Aileron spades These are flat metal plates, usually attached to the aileron lower surface, ahead of the aileron hinge, by a lever arm. They reduce the force needed by the pilot to deflect the aileron and are often seen on aerobatic aircraft. As the aileron is deflected upward, the spade produces a downward aerodynamic force, which tends to rotate the whole assembly so as to further deflect the aileron upward. The size of the spade and its lever arm determine how much force the pilot needs to apply to deflect the aileron. Aileron balance weights To prevent control surface flutter (Aeroelastic flutter), the center of lift of the control surface should be behind the center of gravity of that surface. To achieve this, lead weights may be added to the front of the aileron. In some aircraft the aileron construction may be too heavy to allow this system to work without huge weight increases. In this case, the weight may be added to a lever arm to move the weight well out in front to the aileron body. These balance weights are tear drop shapes (to reduce drag) which make them appear quite different from spades, although both project forward and below the aileron. Page 16 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Types of ailerons Frise Ailerons Engineer Leslie George Frise (1897–1979) developed an aileron shape which is often used due to its ability to counteract adverse yaw. The Frise aileron is pivoted at about its 25 to 30% chord line and near its bottom surface. When the aileron is deflected up to make its wing go down, the leading edge of the aileron dips into the airflow beneath the wing. The movement of the leading edge into the airflow helps to move up the trailing edge and decrease the stick force. As well, increasing drag on the down moving wing to reduce adverse yaw. Differential ailerons By careful design of the mechanical linkages, the up aileron can be made to deflect more than the down aileron (e.g. US patent 1565097). This helps reduce the likelihood of a wing tip stall when aileron deflections are made at high angles of attack. The idea is that the loss of lift associated with the up aileron carries no penalty while the increase in lift associated with the down aileron is minimized. The de Havilland Tiger Moth classic British biplane is one of the best- known aircraft, and one of the earliest, to use differential ailerons. Combinations with other control surfaces Page 17 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Flaperon - A control surface that combines an aileron and flap is called a flaperon. A single surface on each wing serves both purposes: used as an aileron, the flaperons left and right are actuated differentially; when used as a flap, both flaperons are actuated downwards. When a flaperon is actuated downwards (i.e. used as a flap) there is enough freedom of movement left to be able to still use the aileron function. Spoileron - A further form of roll control, common on modern jet transport aircraft, utilises spoilers in conjunction with ailerons. Elevon - On a delta-winged aircraft, the ailerons are combined with the elevators to form an elevon. Several modern fighter aircraft may have no ailerons on the wings at all, and combine roll control with an all moving tailplane. This is a rolling tail. Aileron Droop On some aircraft (notably the De Havilland Beaver) both ailerons move down proportionately to the flaps even though they are still controlled independently. They don’t move as far as the flaps but serve to increase the camber along the entire length of the wing rather than just ahead of the flaps. Photo by Bill Lindsay Page 18 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Rudder (Yaw) Control On an aircraft, the rudder is a directional control surface along with the elevator and ailerons that control pitch and roll. The rudder is usually attached to the fin (or vertical stabilizer) which allows the pilot to control yaw in the vertical axis, i.e. change the horizontal direction in which the nose is pointing. The rudder is controlled by the foot pedals, often in both the pilots and co-pilots positions. In practice, both aileron and rudder control input are used together to turn an aircraft, the ailerons imparting roll, the rudder imparting yaw, and also compensating for a phenomenon called adverse yaw. Adverse yaw is readily seen if the simplest types of ailerons alone are used for a turn. The downward moving aileron acts like a flap, generating more lift for one wing, and therefore more drag. Since the 1930s, many aircraft have used Frise ailerons or differential ailerons, which compensate for the adverse yaw and require little rudder input in regular turns. Initially, this drag yaws the Page 19 of 66 09 June, 2016  AVAV4107 Flight Controls Handout aircraft in the direction opposite to the desired course. A rudder alone will turn a conventional fixed wing aircraft, but much more slowly than if ailerons are also used in conjunction. Use of rudder and ailerons together produces co-ordinated turns, in which the longitudinal axis of the aircraft is in line with the arc of the turn, neither slipping (under-ruddered), nor skidding (over-ruddered). Improperly ruddered turns at low speed can precipitate a spin which can be dangerous at low altitudes. Sometimes pilots may intentionally operate the rudder and ailerons in opposite directions in a manoeuvre called a forward slip. This may be done to overcome crosswinds and keep the fuselage in line with the runway, or to more rapidly lose altitude by increasing drag, or both. Some aircraft have a “V” tail. In this case the two stabs serve as both a rudder and elevator and are called ruddivator. The advantage is one less stab creating drag, the disadvantage is that the aircraft is more susceptible to both pitch and yaw due to turbulence. Beechcraft Bonanza Eclipse Any aircraft rudder is subject to considerable forces that determine its position via a force or torque balance equation. In extreme cases these forces can lead to loss of rudder control or even destruction of the rudder. In large transport aircraft the largest achievable angle of a rudder in flight is called its blow down limit; it is achieved when the force from the air or Page 20 of 66 09 June, 2016  AVAV4107 Flight Controls Handout blow down equals the maximum available hydraulic pressure. The faster the aircraft is, the lower the limit. Secondary Flight Controls Trim tab Trim tabs are small surfaces connected to the trailing edge of a larger control surface, such as a rudder, elevator or aileron and is used to control the trim of the controls, i.e. to counteract aerodynamic forces and stabilise the aircraft in a particular desired attitude without the need for the operator to constantly apply a control force. This is done by adjusting the angle of the tab relative to the larger surface. Changing the setting of a trim tab adjusts the neutral or resting position of a control surface (such as an elevator or rudder). As the desired position of a control surface changes (corresponding mainly to different speeds), an adjustable trim tab will allow the operator to reduce the manual force required to maintain that position—to zero. Thus the trim tab acts as a servo tab. Because the center of pressure of the trim tab is further away from the axis of rotation of the control surface than the center of pressure of the control surface, the movement generated by the tab can match the movement generated by the control surface. The position of the control surface on its axis will change until the movements from the control surface and the trim surface balance each other. Page 21 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Lift Augmentation Flaps High lift devices Flaps are hinged surfaces on the trailing edge of the wings of a fixed-wing aircraft. As flaps are extended, the stalling speed of the aircraft is reduced, which means that the aircraft can fly safely at lower speeds (especially during takeoff and landing). Flaps are also used on the leading edge of the wings of some high-speed jet aircraft, where they may be called Krueger flaps Extending flaps increases the camber of the wing airfoil, thus raising the maximum lift coefficient. This increase in maximum lift coefficient allows the aircraft to generate a given amount of lift with a lower speed. Therefore, extending the flaps reduces the stalling speed of the aircraft. Page 22 of 66 09 June, 2016  AVAV4107 Flight Controls Handout The left wing of an Airbus A319-100. The three canoe-shapes are flap track fairings to hide and streamline the flap driving mechanisms. The flaps (two on each side, on the A319) lay directly above the flap track fairings. Extending flaps also increases drag. This can be beneficial in the approach and landing phase because it helps to slow the aircraft. Another useful side effect of flap deployment is a decrease in aircraft pitch angle. This provides the pilot with a greater view over the nose of the aircraft and allows a better view of the runway during approach and landing. A fully extended flap before landing Page 23 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Some trailing edge flap systems increase the plan form area of the wing in addition to changing the camber. In turn, the larger lifting surface allows the aircraft to generate a given amount of lift with a lower speed, thus further reducing stalling speed. Although this effect is very similar to increasing the lift coefficient, raising the plan form area of the wing does not itself raise the lift coefficient. The Fowler flap is an example of a flap system that increases the plan form area of the wing in addition to increasing the camber (lift coefficient). Extending the flaps also increases the drag coefficient of the aircraft. Therefore, for any given weight and airspeed, flaps increase the drag force. Flaps increase the drag coefficient of an aircraft because of higher induced drag caused by the distorted span wise lift distribution on the wing with flaps extended. Some flaps increase the plan form area of the wing and, for any given speed, this also increases the parasitic drag component of total drag. Depending on the aircraft type, flaps may be partially extended for takeoff. With general aviation aircraft, the use of flaps for takeoff may be optional. This depends on the manufacturer's procedures in the Airplane Flight Manual for a specific takeoff method (e.g., short field, soft field, normal, etc.). Flaps may be partially extended on takeoff to increase the amount of lift generated at a given airspeed, as well as to reduce the stalling speed of the airplane. Together, these two effects help an airplane lift off in a shorter distance at a lower drag penalty than that incurred by a full flap deflection. Flaps are usually fully extended for landing to give the aircraft a lower stalling speed so the approach to landing can be flown more slowly, allowing the aircraft to land in a shorter distance. The higher lift and drag associated with fully extended flaps allows a steeper and slower approach for landing. This demonstrates the combined benefit of the higher lift and drag coefficients of fully extended flaps. There are many types of flaps or combinations, some flap systems include: Krueger flap: hinged flap on the leading edge. Often called a "droop". Plain flap: rotates on a simple hinge. Split flap: upper and lower surfaces are separate, the lower surface operates like a plain flap, but the upper surface stays immobile or moves only slightly. Page 24 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Gouge flap: a cylindrical or conical aerofoil section which rotates backwards and downwards about an imaginary axis below the wing, increasing wing area and chord without affecting trim. Invented by Arthur Gouge for Short Brothers in 1936. Fowler flap: slides backwards before hinging downwards, thereby increasing both camber and chord, creating a larger wing surface better tuned for lower speeds. It also provides some slot effect. The Fowler flap was invented by Harlan D. Fowler. Fairey-Youngman flap: moves body down before moving aft and rotating. Slotted flap: a slot (or gap) between the flap and the wing enables high pressure air from below the wing to re-energize the boundary layer over the flap. This helps the airflow to stay attached to the flap, delaying the stall. Blown flaps: systems that blow engine air over the upper surface of the flap at certain angles to improve lift characteristics such as this Antonov. Antonov An-72 Page 25 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Page 26 of 66 09 June, 2016  AVAV4107 Flight Controls Handout As the graph above shows slotted and fowler provide greater lift before they increase the drag significantly. Leading edge slats Slats are aerodynamic surfaces on the leading edge of the wings of fixed- wing aircraft which, when deployed, allow the wing to operate at a higher angle of attack. A higher coefficient of lift is produced as a product of angle of attack and speed, so by deploying slats an aircraft can fly more slowly or take off and land in a shorter distance. They are usually used while landing or performing manoeuvres which take the aircraft close to the stall, but are usually retracted in normal flight to minimize drag. Types include: Page 27 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Automatic - the slat lies flush with the wing leading edge until reduced aerodynamic forces allow it to extend by way of springs when needed. Fixed - the slat is permanently extended. This is sometimes used on specialist low-speed aircraft (these are referred to as slots) or when simplicity takes precedence over speed. Powered - the slat extension can be controlled by the pilot. This is commonly used on airliners. Leading edge slats, have a similar function as the trailing-edge flaps. Note that a Krueger flap and a leading-edge slat differ in how they are extended. A slat provides a separation of the leading edge from the rest of the wing for air to pass from the bottom of the surface to the top, delaying boundary layer separation, whereas a Krueger flap does not because it only increases the wing area and wing curvature. Some aircraft slats are automatic and deploy when the airspeed is low enough that the impact air no longer holds the leading edge against the wing and the slat is forced forward with spring pressure. Drooped Leading edge Extended Retected Page 28 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Kruger Flap Retracted Extended Leading Edge Slats Retracted Extended Page 29 of 66 09 June, 2016  AVAV4107 Flight Controls Handout The position of the leading edge slats on an airliner (Airbus A310-300). In this picture, the slats are extended, note also the deployed trailing edge flaps. Leading edge slots Slots are leading edge devices that are similar to slats and perform a similar function however, slots are fixed in position. Their purpose is the same - to encourage the boundary layer to stay attached further aft on the wing. Page 30 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Globe swift wing The Fieseler Fi 156 Storch had permanently extended slots on its leading edges (fixed slats). Operation The chord of the slat is typically only a few percent of the wing chord. The slats may extend over the outer third of the wing, or they may cover the entire leading edge. Many early aerodynamicists, including Ludwig Prandtl believed that slats work by inducing a high energy stream to the flow of the main airfoil thus re-energizing its boundary layer and delaying stall. In reality, the slat does not give the air in the slot high velocity (it actually reduces its velocity) and also it cannot be called high-energy air since all the air outside the actual boundary layers has the same total head. The actual effects of the slat are: Page 31 of 66 09 June, 2016  AVAV4107 Flight Controls Handout The slat effect: The velocities at the leading edge of the downstream element (main airfoil) are reduced due to the circulation of the upstream element (slat) thus reducing the pressure peaks of the downstream element. The circulation effect: The circulation of the downstream element increases the circulation of the upstream element thus improving its aerodynamic performance. The dumping effect: The discharge velocity at the trailing edge of the slat is increased due to the circulation of the main airfoil thus alleviating separation problems or increasing lift. Off the surface pressure recovery: The deceleration of the slat wake occurs in an efficient manner, out of contact with a wall. Fresh boundary layer effect: Each new element starts out with a fresh boundary layer at its leading edge. Thin boundary layers can withstand stronger adverse gradients than thick ones. Spoilers Spoilers, general description Spoilers reduce lift on the wing. If raised on only one wing, they aid roll control, causing that wing to drop. If the spoilers rise symmetrically in flight, the aircraft can either be slowed in level flight or can descend rapidly without an increase in airspeed. When the spoilers rise on the ground at high speeds, they reduce the wing's lift, which puts more of the aircraft's weight on the wheels. The flight spoilers are available both in flight and on the ground. However, the ground spoilers can only be raised when the weight of the aircraft is on the landing gear, usually activated by a sensor. When the spoilers deploy on the ground, they decrease lift and make the brakes more effective. In flight, a ground-sensing switch on the landing gear prevents deployment of the ground spoilers. Page 32 of 66 09 June, 2016  AVAV4107 Flight Controls Handout The inner workings of the spoiler during the landing of an Airbus A320. The spoiler during the landing of an Airbus A321. In aeronautics, a spoiler (sometimes called a lift dumper) is a device intended to reduce lift in an aircraft. Spoilers are plates on the top surface of a wing which can be extended upward into the airflow and spoil it. By doing so, the spoiler creates a carefully controlled stall over the portion of the wing behind it, greatly reducing the lift of that wing section. Spoilers differ from airbrakes in that airbrakes are designed to increase drag making Page 33 of 66 09 June, 2016  AVAV4107 Flight Controls Handout little change to lift, while spoilers greatly reduce lift making only a moderate increase in drag. Spoilers are used by gliders to control their rate of descent and thus achieve a controlled landing at a desired spot. An increased rate of descent could also be achieved by lowering the nose of an aircraft, but this would result in an excessive landing speed. However spoilers enable the approach to be made at a safe speed for landing. Airliners too, are usually fitted with spoilers. Spoilers are sometimes used when descending from cruise altitudes to assist the aircraft in descending to lower altitudes without picking up speed. Their use is often limited, however, as turbulent airflow which develops behind them causes noticeable noise and vibration, which may cause discomfort to extra- sensitive passengers. The spoilers may also be differentially operated to provide roll control. On landing, however, the spoilers are nearly always used at full effect to assist in slowing the aircraft. The increase in form drag created by the spoilers directly assists the braking effect. However, the real gain comes as the spoilers cause a dramatic loss of lift and hence the weight of the aircraft is transferred from the wings to the undercarriage, allowing the wheels to be mechanically braked with much less chance of skidding. Reverse thrust is also often further used to help slow the aircraft on landing. Spoilers as control surfaces Some aircraft use spoilers in combination with or in lieu of ailerons for roll control, primarily to reduce adverse yaw when rudder input is limited by higher speeds. For such spoilers the term spoileron has been coined. In the case of a spoileron, in order for it to be used as a control surface, it is raised on one wing, thus decreasing lift and increasing drag, causing roll and yaw. Page 34 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Mitsubishi MU2 uses spoilers only Methods of control Hydro-mechanical Controls The complexity and weight of mechanical flight control systems increase considerably with the size and performance of the aircraft. Hydraulically powered control surfaces help to overcome these limitations. With hydraulic flight control systems, the aircraft's size and performance are limited by economics rather than a pilot's muscular strength. At first, only-partially boosted systems were used in which the pilot could still feel some of the aerodynamic loads on the control surfaces (feedback). A hydro-mechanical flight control system has two parts: The mechanical circuit, which links the cockpit controls with the hydraulic circuits. Like the mechanical flight control system, it consists of rods, cables, pulleys, and sometimes chains. The hydraulic circuit, which has hydraulic pumps, reservoirs, filters, pipes, valves and actuators. The actuators are powered by the hydraulic pressure generated by the pumps in the hydraulic circuit. The actuators convert hydraulic pressure into control surface movements. The pilot's movement of a control causes the mechanical circuit to open the matching servo valve in the hydraulic circuit. The hydraulic circuit powers Page 35 of 66 09 June, 2016  AVAV4107 Flight Controls Handout the actuators which then move the control surfaces. As the actuator moves, the servo valve is closed by a mechanical feedback linkage - one that stops movement of the control surface at the desired position. This arrangement was found in the older-designed jet transports and in some high- performance aircraft. Artificial feel devices With purely mechanical flight control systems, the aerodynamic forces on the control surfaces are transmitted through the mechanisms and are felt directly by the pilot. This gives tactile feedback of airspeed and aids flight safety. With hydro mechanical flight control systems however, the load on the surfaces cannot be felt and there is a risk of overstressing the aircraft through excessive control surface movement. To overcome this problem artificial feel systems are used. Page 36 of 66 09 June, 2016  AVAV4107 Flight Controls Handout The following drawing is a typical spring and cam operated feel unit. The cam is attached to input shaft and rotates with control input to the flight control system. As the cam rotates the cam follower roller is forced up the edge of the cam, this causes the feel and centering springs to stretch and create a force against the control input. As the control input force is reduced the centering springs force the roller into the detent of the cam centering the control system. The control trim actuator is attached to the mechanism to adjust the neutral point of the centering mechanism. This will adjust the neutral point to move the control system and trim the control system, the cockpit and control surface move. Trim and Centering Mechanism A hydro-pneumatic system is used to give a variable feel to some elevator systems. This uses air-data from a pitot system to vary the amount of feel relative to the airspeed, therefore as air speed increases the amount of opposition to pilot input increases. The Boeing 737 uses this system. Full power controls Page 37 of 66 09 June, 2016  AVAV4107 Flight Controls Handout The true hydro-mechanical control system does not have any mechanical back up capabilities. To compensate for this feature the aircraft uses multiple hydraulics systems. A hydraulic system powered by engine driven pumps and electrical hydraulic pumps. Pneumatically powered controls Some aircraft such as the B747 use aircraft pneumatic air to operate the leading edge flap system. This pneumatic energy comes from the aircraft pneumatic system. Below is an example of pneumatically powered leading edge flaps. Page 38 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Electrically powered flight controls The use of electrical motors to operate flight controls is very common. Electrically powered flaps and control surface trim is used on many general aviation aircraft. Larger aircraft use electric motors to operate control surface trim, and as alternate power sources for trailing edge and leading edge flap systems. Auto pilot systems interface with the flight control systems through electrical servo motors. Fly-by-wire control systems A fly-by-wire (FBW) system replaces manual flight control of an aircraft with an electronic interface. The movements of flight controls are converted to electronic signals transmitted by wires (hence the fly-by-wire term), and flight control computers determine how to move the actuators at each control surface to provide the expected response. Commands from the computers are also input without the pilot's knowledge to stabilize the aircraft and perform other tasks. Fly-by-optics, also known as fly-by-light, is a further development using fiber optic cables. Page 39 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Development of Fly by Wire Systems Mechanical and hydro-mechanical flight control systems are relatively heavy and require careful routing of flight control cables through the aircraft by systems of pulleys, cranks, tension cables and hydraulic pipes. Both systems often require redundant backup to deal with failures, which again increases weight. Furthermore, both have limited ability to compensate for changing aerodynamic conditions. Dangerous characteristics such as stalling, spinning and pilot-induced oscillation (PIO), which depend mainly on the stability and structure of the aircraft concerned rather than the control system itself, can still occur with the systems. The term "fly-by-wire" implies a purely electrically-signaled control system. However, it is used in the general sense of computer-configured controls, where a computer system is interposed between the operator and the final control actuators or surfaces. This modifies the manual inputs of the pilot in accordance with control parameters. Page 40 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Side-sticks, center sticks, or conventional flight control yokes can be used to fly FBW aircraft. While the side-stick offers the advantages of being lighter, mechanically simpler, and unobtrusive, The Boeing Company's aerospace engineers decided that the lack of visual feedback (none given by side-sticks) is a significant problem, and so they designed conventional control yokes in the Boeing 777 and the brand-new Boeing 787. This same approach has been used for the Embraer 170/190 jets however; most Airbus airliners are operated with side-sticks. Fly-by wire systems are by their nature quite complex; however their operation can be explained in relatively simple terms. When a pilot moves the control column (or sidestick), a signal is sent to a computer, this is analogous to moving a game controller, the signal is sent through multiple wires (channels) to ensure that the signal reaches the computer. When there are three channels being used this is known as 'Triplex'. The computer receives the signals, performs a calculation (adds the signal voltages and divides by the number of signals received to find the mean average voltage) and adds another channel. These four 'Quadruplex' signals are then sent to the control surface actuator and the surface begins to move. Potentiometers in the actuator send a signal back to the computer (usually a negative voltage) reporting the position of the actuator. When the actuator reaches the desired position the two signals (incoming and outgoing) cancel each other out and the actuator stops moving (completing a feedback loop). Automatic Stability Systems Fly-by-wire control systems allow aircraft computers to perform tasks without pilot input. Automatic stability systems operate in this way. Gyroscopes fitted with sensors are mounted in an aircraft to sense movement changes in the pitch, roll and yaw axes. Any movement (from straight and level flight for example) results in signals to the computer, which automatically moves control actuators to stabilize the aircraft. Manoeuver Load Alleviation The manoeuver load alleviation function is achieved by two ailerons and three spoilers on each wing. These surfaces are deflected upwards on both wings. Page 41 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Maneuver Load Alleviation (MLA) function is to reduce the loads applied to the wing structure during high g manoeuver to prevent overstress. Turbulence Damping The purpose of the Turbulence Damping Function is to damp the structural loads induced by atmospheric turbulence. The longitudinal and lateral Turbulence Damping commands are computed by Flight Control Computers as a function of the vertical and lateral acceleration information. They are added to the normal computer command and transmitted to the associated elevator servo-controls and yaw damper system for take-off, approach and landing. Safety and redundancy Aircraft systems may be quadruplexed (four independent channels) to prevent loss of signals in the case of failure of one or even two channels. High performance aircraft that have FBW controls (also called CCVs or Control-Configured Vehicles) may be deliberately designed to have low or even negative stability in some flight regimes, the rapid-reacting CCV controls compensating for the lack of natural stability. Pre-flight safety checks of a fly-by-wire system are often performed using Built-In Test Equipment (BITE). On programming the system, either by the pilot or ground crew, a number of control movement steps are automatically performed. Any failure will be indicated to the crews. Some aircraft, the Panavia Tornado for example, retain a very basic hydro- mechanical backup system for limited flight control capability on losing electrical power, in the case of the Tornado this allows rudimentary control of the stabilators only for pitch and roll axis movements. Weight saving A FBW aircraft can be lighter than a similar design with conventional controls. Partly due to the lower overall weight of the system components; and partly because the natural stability of the aircraft can be relaxed, slightly for a transport aircraft and more for a maneuverable fighter, which means that the stability surfaces that are part of the aircraft structure can therefore be made smaller. These include the vertical and horizontal Page 42 of 66 09 June, 2016  AVAV4107 Flight Controls Handout stabilizers. If these structures can be reduced in size, airframe weight is reduced. The advantages of FBW controls were first exploited by the military and then in the commercial airline market. The Airbus series of airliners used full-authority FBW controls beginning with their A320 series (though some limited FBW functions existed on A310). Boeing followed with their 777 and later designs. Electronic fly-by-wire systems can respond flexibly to changing aerodynamic conditions, by tailoring flight control surface movements so that aircraft response to control inputs is appropriate to flight conditions. Electronic systems require less maintenance, whereas mechanical and hydraulic systems require lubrication, tension adjustments, leak checks, fluid changes, etc. Furthermore, putting circuitry between pilot and aircraft can enhance safety; for example the control system can try to prevent a stall, or it can stop the pilot from over stressing the airframe. The main concern with fly-by-wire systems is reliability. While traditional mechanical or hydraulic control systems usually fail gradually, the loss of all flight control computers could immediately render the aircraft uncontrollable. For this reason, most fly-by-wire systems incorporate either redundant computers (triplex, quadruplex etc.), some kind of mechanical or hydraulic backup or a combination of both. A "mixed" control system such as the latter is not desirable and modern FBW aircraft normally avoid it by having more independent FBW channels, thereby reducing the possibility of overall failure to minuscule levels that are acceptable to the independent regulatory and safety authority responsible for aircraft design, testing and certification before operational service. History F-8C Crusader digital fly-by-wire testbed Page 43 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Electronic signalling of the control surfaces was tested in the 1950s. This replaced long runs of mechanical and hydraulic connections with electrical ones. The first non-experimental aircraft that was designed and flown (in 1958) with a fly-by-wire flight control system was the Avro Canada CF-105 Arrow (analog computer). A feat not repeated with a production aircraft until the Concorde in 1969. This system also included solid-state components and system redundancy, was designed to be integrated with a computerised navigation and automatic search and track radar, was flyable from ground control with data uplink and downlink, and provided artificial feel (feedback) to the pilot. The first digital fly-by-wire aircraft to take to the air (in 1972) was an F-8 Crusader, which had been modified electronically by the National Aeronautics and Space Administration of the United States as a test aircraft, a feat mirrored in the USSR by the Sukhoi T-4. At about the same time in the United Kingdom a trainer variant of the British Hawker Hunter fighter was modified at the British Royal Aircraft Establishment with fly-by- wire flight controls for the right-seat pilot. This was test-flown, with the left- seat pilot having conventional flight controls for safety reasons, and with the capability for him to override and turn off the fly-by-wire system. Analog systems All "fly-by-wire" flight control systems eliminate the complexity, the fragility, and the weight of the mechanical circuit of the hydro mechanical or electromechanical flight control systems. Fly-by-wire replaces those with electronic circuits. The control mechanisms in the cockpit now operate signal transducers, which in turn generate the appropriate electronic commands. These are next processed by an electronic controller, either an analog one, or more modernly, a digital one. Aircraft and spacecraft autopilots are now part of the electronic controller. The hydraulic circuits are similar except that mechanical servo valves are replaced with electrically-controlled servo valves, operated by the electronic controller. This is the simplest and earliest configuration of an analog fly-by- wire flight control system. In this configuration, the flight control systems must simulate "feel". The electronic controller controls electrical feel Page 44 of 66 09 June, 2016  AVAV4107 Flight Controls Handout devices that provide the appropriate "feel" forces on the manual controls. This was used in Concorde, the first production fly-by-wire airliner. In more sophisticated versions, analog computers replaced the electronic controller. The canceled 1950s Canadian supersonic interceptor, the Avro Canada CF-105 Arrow, employed this type of system. Analog computers also allowed some customization of flight control characteristics, including relaxed stability. This was exploited by the early versions of F-16, giving it impressive maneuverability. Digital systems The Airbus A320, first airliner with digital fly-by-wire controls A digital fly-by-wire flight control system is similar to its analog counterpart. However, the signal processing is done by digital computers and the pilots literally can "fly-via-computer". This also increases the flexibility of the flight control system, since the digital computers can receive input from any aircraft sensor (such as the altimeters and the pitot tubes. This also increases the electronic stability, because the system is less dependent on the values of critical electrical components as in an analog controller. The computers sense position and force inputs from pilot controls and aircraft sensors. They solve differential equations to determine the appropriate command signals that move the flight controls to execute the intentions of the pilot. The programming of the digital computers enables flight envelope protection. In this aircraft, designers precisely tailor an aircraft's handling characteristics, to stay within the overall limits of what is possible given the Page 45 of 66 09 June, 2016  AVAV4107 Flight Controls Handout aerodynamics and structure of the aircraft. For example, the computer in flight envelope protection mode can try to prevent the aircraft from being handled dangerously by preventing pilots from exceeding pre-set limits on the aircraft's flight-control envelope, such as those that prevent stalls and spins, and which limit airspeeds and g forces on the airplane. Software can also be included that stabilize the flight-control inputs to avoid pilot-induced oscillations. Since the flight-control computers continuously "fly" the aircraft, pilot's workloads can be reduced. Also, in military and naval applications, it is now possible to fly military aircraft that have relaxed stability. The primary benefit for such aircraft is more maneuverability during combat and training flights and the so-called "carefree handling" because stalling, spinning, and other undesirable performances are prevented automatically by the computers. Typical Electronic Flight Control System Modern electronic flight control systems are electrically controlled and hydraulically driven. Cockpit controls are a joy stick/control wheel for pilots to input to the computer systems. The computer systems are primary and secondary. The primary computer system establishes the order of the control laws, Page 46 of 66 09 June, 2016  AVAV4107 Flight Controls Handout speed brake and ground spoiler control. The laws are Normal, Alternate, and Direct. The secondary computer system establishes control laws for direct control laws, including yaw damper function, rudder trim, rudder travel limit, and pedal travel limit. One computer primary or secondary is capable of controlling the aircraft and assuring safe flight and landing. Normal operation, one primary computer is declared to be the master. It processes orders and send them to the other computers, which will then execute them on their related servo control. The normal law provides, three axis control, flight envelop protection, and manoeuver load protection. For failures of the type that affect the flight control system, or peripherals, there are three possible reconfiguration levels; alternate law, direct law, and mechanical. With these failures the protective features will be degrades or lost completely. The electronic flight control system protects the airframe in the following ways. Load factor limitation which limits the max gravitational loads. Pitch attitude protection which limits max nose up. High angle of attack which protects against stall and wind shear. (This protection has priority over all other protections.) High speed protection which applied a nose up order to reduce airspeed. Low energy protection which increases thrust to recover from low speed (energy) situation. Maneuver load alleviation which redistributes the lift over the wing to relieve structural loads on the outer wing surfaces (bending moment). Turbulence damping function which dampens structural loads induced by atmosphere turbulence. Mechanical backup is used to control the aircraft in the case of a complete loss of electrical power. In the pitch mode the stabilizer is trimmed by cockpit controls. This form of trim has priority over electrical control. In the lateral mode the rudder pedals are used to operate the rudder, with a backup from the yaw damper unit. Page 47 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Lift augmentation is provided by flaps and slats. The surfaces are electrically controlled and hydraulically operated. Flaps and slats are provided with protection for the following; Asymmetry protection between left and right wing. Flaps disconnect detection system which detects attachment failure and inhibit flap operation. Flap load relief system which automatically retracts the flaps to a pre- set position to prevent structural damage to the flaps from high air loads. Slats alpha prevents slat retraction at high angles of attack (stall condition). Applications A Dassault Falcon 7X, the first business jet with digital fly-by-wire controls The Space Shuttle Orbiter has an all-digital fly-by-wire control system. This system was first exercised (as the only flight control system) during the glider unpowered-flight "Approach and Landing Tests" that began on the Space Shuttle Enterprise during 1977. During 1984, the Airbus Industries Airbus A320 became the first airliner to fly with an all-digital fly-by-wire control system. During 2005, the Dassault Falcon 7X became the first business jet with fly-by-wire controls. Page 48 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Redundancy If one of the flight-control computers crashes, or is damaged in combat, or suffers from "insanity" caused by electromagnetic pulses, the others overrule the faulty one (or even two of them), they continue flying the aircraft safely, and they can either turn off or re-boot the faulty computers. Any flight-control computer whose results disagree with the others is ruled to be faulty, and it is either ignored or re-booted. (In other words, it is voted- out of control by the others.) In addition, most of the early digital fly-by-wire aircraft also had an analog electrical, a mechanical, or a hydraulic back-up flight control system. The Space Shuttle has, in addition to its redundant set of four digital computers running its primary flight-control software, a fifth back-up computer running a separately developed, reduced-function, software flight-control system - one that can be commanded to take over in the event that a fault ever affects all of the computers in the other four. This back-up system serves to reduce the risk of total flight-control-system failure ever happening because of a general-purpose flight software fault that has escaped notice in the other four computers. For airliners, flight-control redundancy improves their safety, but fly-by-wire control systems also improve economy in flight because they are lighter, and they eliminate the need for many mechanical, and heavy, flight-control mechanisms. Furthermore, most modern airliners have computerized systems that control their jet engine throttles, air inlets, fuel storage and distribution system, in such a way to minimize their consumption of jet fuel. Thus, digital control systems do their best to reduce the cost of flights. Airbus/Boeing Airbus and Boeing commercial airplanes differ in their approaches in using fly-by-wire systems. In Airbus airliners, the flight-envelope control system always retains ultimate flight control, and it will not permit the pilots to fly outside these performance limits. However, in the event of multiple failures of redundant computers, the A320 does have mechanical back-up system for its pitch trim and its rudder. The A340-600 has a purely electrical (not electronic) back-up rudder control system, and beginning with the new A380 airliner, all flight-control systems have back-up systems that are Page 49 of 66 09 June, 2016  AVAV4107 Flight Controls Handout purely electrical through the use of a so-called "three-axis Backup Control Module" (BCM) With the Boeing 777 model airliners, pilots can completely override the computerized flight-control system to permit the aircraft to be flown beyond its usual flight-control envelope during emergencies. Airbus's strategy, which began with the Airbus A320, has been continued on subsequent Airbus airliners. Aircraft with fly-by-wire flight controls require computer controlled flight control modes that are capable of determining the operational mode (computational law) of the aircraft. A reduction of electronic flight control can be caused by the failure of a computational device, such as the flight control computer or an information providing device, such as the ADIRU (Air data inertial reference unit). Electronic flight control systems (EFCS) also provide augmentation in normal flight, such as increased protection of the aircraft from overstress or providing a more comfortable flight for passengers by recognizing and correcting for turbulence and providing yaw damping. Two aircraft manufacturers produce commercial passenger aircraft with primary flight computers that can perform under different flight control modes (or laws). The most well-known are the Normal, Alternate, Direct and Mechanical Laws of the Airbus A320-A380. Boeing's fly-by-wire system is used in the Boeing 777. Boeing also has two other commercial aircraft the 787 and the 747-8, which use fly-by-wire controls. This newer generation of aircraft use the lighter weight electronic systems to increase safety and performance while lowering aircraft weight. Since these systems can also protect the aircraft from overstress situations, the designers can therefore reduce over-engineered components, further reducing weight Supplemental Flight Control Systems Page 50 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Yaw Damper A yaw damper is a device used on many aircraft (usually jets and turboprops) to dampen (reduce) the rolling and yawing oscillations known as Dutch roll. It involves yaw rate sensors and a processor that provides a signal to an actuator connected to the rudder. The use of the yaw damper helps to provide a better ride for passengers and on some aircraft is a required piece of equipment to ensure that the aircraft stability remains within certification values. Another function of the yaw damper is to automatically help to correct for the yaw that occurs due to asymmetric thrust in the event of an engine failure. The Yaw Damper System uses the following inputs to dampen undesirable sideslip and rolling motions: (1) Air Data Computer (2) Inertial Reference Unit (3) Modal Suppression Accelerometer The Yaw Damper System, commands drive yaw damper servos, which control the rudder mechanical linkage and power control actuators. Rudder displacement commands are proportional to yaw rate. Turn coordination is also provided by the Yaw Damper System. Roll attitude inputs are used to compute turn coordination commands. These commands drive the yaw damper servos, which control the rudder through power control actuators. The computers monitor system operation and indicate system faults through automatic and manually initiated system testing. Yaw Damper Servos The yaw damper servos use electrical command inputs from the yaw damper modules. The electrical input commands control hydraulic flow to actuator pistons, which provide a mechanical output. This output is connected in series with manual and autopilot rudder inputs supplied to rudder power control actuators, which move the rudder. The maximum rudder authority for the yaw damping is less than the normal rudder travel. The yaw damper servo-actuators form the main electrical inputs of the rudder control. They transmit the rudder deflection orders to the rudder servo control through the differential mechanism. Page 51 of 66 09 June, 2016  AVAV4107 Flight Controls Handout These are controlled by the lateral control laws which include turn coordination and Dutch roll damping in manual or in autopilot mode. Mach/speed trim Mach/speed trim is used to compensate for the effects Mach tuck. Mach tuck is the result of an aerodynamic stall due to an over-speed condition, rather than the more common stalls resulting from boundary layer separation due to insufficient airspeed, increased angle of attack, excessive load factors, or a combination of those causes. As the aircraft's wing approaches its critical Mach number, the aircraft is traveling below Mach 1.0. However, the accelerated airflow over the upper surface of the cambered wing exceeds Mach 1.0 and a shock wave is created at the point on the wing where the accelerated airflow has gone supersonic. While the air ahead of the shock wave is in laminar flow, a boundary layer separation is created aft of the shock wave, and that section of the wing fails to produce lift. The image above illustrates this concept. Page 52 of 66 09 June, 2016  AVAV4107 Flight Controls Handout In most aircraft susceptible to Mach tuck, the camber at the wing root, the section of the wing closest to the fuselage, is more pronounced than that of the wing tip. This design ensures that in a standard stall the root will stall before the tips. This allows the pilot to recognize the stall while still maintaining control of the ailerons to enhance stall recovery. However, this also means that when an airfoil exceeds its critical Mach number, the shock wave, and resulting stall condition, will begin to form at the root. A second design element that leads to Mach tuck is that many aircraft which will approach the speed of sound are designed with swept wings. The center of pressure of a wing is an imaginary point where the summation of all lifting forces across the wing's surface can be resolved into a single lift vector. When the wing root stalls, the center of pressure of the (reduced) lift being generated by the wing is shifted towards the wing tip. With a swept wing, this also means that the center of pressure travels aft (because it is traveling out from the wing root and therefore backwards as the wing sweeps). When the center of pressure moves aft, its movement rearwards compared to the unmoving center of mass of the aircraft will generate a force which will act to depress the nose of the aircraft; this nose down pitching moment is “Mach tuck." As the wing becomes more affected by the shock wave the center of pressure will continue to travel aft, thereby causing a significantly higher nose-down force and requiring a nose-up input or trim to maintain level flight. Although Mach tuck develops gradually, if it is allowed to progress significantly, the center of pressure can move so far rearward that there is no longer enough elevator authority available to counteract it, and the airplane enters a steep, sometimes unrecoverable dive. In addition, until the aircraft goes supersonic, as the shock wave goes towards the rear, because of the faster flow there the top shockwave will impinge upon the horizontal stabilizer and elevator control surfaces further back than the lower shockwave; this can greatly exacerbate the nose down tendencies. The horizontal stabilizer at the tail of the aircraft generates a downward force, so loss of effective horizontal stabilizer area will reduce this downward force, so the tail will pitch up and the nose will pitch down. If the shock wave affects the elevators, it may reduce their effectiveness, making it impossible for the pilot to alter the aircraft's pitch. Page 53 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Finally, there is a related condition that can exacerbate Mach tuck. If enough of the wing surface becomes engulfed in the shock wave, the wing will not produce enough lift to support the aircraft, and a standard stall will occur. This often fatal combination of over speed and aerodynamic stall can most easily be avoided by not allowing the effects of Mach tuck to develop beyond its incipient stage. This is best accomplished by retarding the throttle, extending speed brakes, and if possible, extending the landing gear. Any actions, which would increase aerodynamic drag and thus reduce airspeed below critical Mach, will prevent further aggravation of the condition. Dealing with Mach tuck All supersonic aircraft experience some degree of Mach tuck. Historically, recovery from a Mach tuck in subsonic aircraft has not always been possible. In some cases, as the aircraft descends, the air density increases and the extra drag will slow the aircraft and control will return. For aircraft such as supersonic fighters/bombers or supersonic transports such as Concorde that spend long periods in supersonic flight, Mach tuck is often compensated for by moving fuel between tanks in the fuselage to change the position of the centre of mass. This minimizes the amount of trim required and keeps the changing location of the center of pressure within acceptable limits. Supersonic and subsonic aircraft often have an all-moving tailplane (a stabilator) rather than separate elevator control surfaces. This avoids the shock wave making the control surfaces pitch downwards. Transport aircraft use a system called Mach/speed trim to compensate for the feature. This system is a feature of the Auto Flight system. The Mach trim system engages automatically if the aircraft is in the Mach tuck region, air data computer information is valid, system power is present and the computed actuator position and actual position agree. This will allow the auto flight system to apply an aircraft nose up trim to compensate for the shift of the center of lift caused by the increase of air speed. As the air speed decreases the auto flight system reduces the pitch up signal. Through this operation the aircraft attitude remains constant. Rudder limiter Page 54 of 66 09 June, 2016  AVAV4107 Flight Controls Handout The function of the rudder limiter is to provide an automatic means of varying the amount of travel of the rudder control system as a function of air speed. This accomplished by providing airspeed signals to the rudder limiter actuator which in turn drives the rudder limiter mechanism to vary the amount the mechanical control system moves the power control actuators. Airspeed is from the air data system and is processed by a computer to determine the correct amount of rudder travel against the air speed. Rudder limiter is not the same as the blow down limit discussed previously. Gust lock Gust lock on a rudder. A gust lock on an aircraft is a mechanism that locks control surfaces in place preventing random movement and possible damage of the surface from wind while parked. Gust locks may be internal (part of the control system) or external as in the rudder photo above. Page 55 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Safety A gust lock can pose a serious safety hazard if its removal is omitted before an aircraft's takeoff because it renders the flight control inoperative. The prototype of the B-17's crash on October 30, 1935 The very first example built of the Boeing B-17 Flying Fortress, the initial Model 299 aircraft, was lost in just this way on October 30, 1935, when its self-contained gust locks were left engaged, with the resulting crash killing Boeing chief test pilot Leslie Tower, and United States Army Air Corps test pilot Ployer Peter Hill. Prince Gustaf Adolf of Sweden, the American singer and actress Grace Moore and 21 others were killed in 1947 during the crash of a KLM flight at Copenhagen Airport due to the flight crew forgetting to disengage the gust lock on the tail fin of the aircraft. For this reason, gust lock disengagement and a control function check is a very important step on the takeoff checklist. As well, all locks must be in plain view and easily identifiable by red streamers or flags usually marked “remove before flight”. Some internal locks prevent engine start or throttle movement until they are removed. Stall warning and safety devices Aeroplanes can be equipped with devices to prevent or postpone a stall or to make it less (or in some cases more) severe, or to make recovery easier. An aerodynamic twist can be introduced to the wing with the leading edge near the wing tip twisted downward. This is called Page 56 of 66 09 June, 2016  AVAV4107 Flight Controls Handout washout and causes the wing root to stall before the wing tip. This makes the stall gentle and progressive. Since the stall is delayed at the wing tips, where the ailerons are, roll control is maintained when the stall begins. A stall strip is a small sharp-edged device which, when attached to the leading edge of a wing, encourages the stall to start there in preference to any other location on the wing. If attached close to the wing root it makes the stall gentle and progressive; if attached near the wing tip it encourages the aircraft to drop a wing when stalling. A stall fence is a flat plate in the direction of the chord to stop separated flow progressing out along the wing Vortex generators, tiny strips of metal or plastic placed on top of the wing near the leading edge that protrudes past the boundary layer into the free stream. As the name implies, they energize the boundary layer by mixing free stream airflow with boundary layer flow thereby creating vortices, this increases the inertia of the boundary layer. By increasing the inertia of the boundary layer airflow separation and the resulting stall may be delayed. An anti-stall strake is a leading edge extension which generates a vortex on the wing upper surface to postpone the stall. A stick pusher is a mechanical device which prevents the pilot from stalling an aeroplane. It pushes the elevator control forwards as the stall is approached, causing a reduction in the angle of attack. Generically, a stick pusher is known as a stall identification device or stall identification system. A stick shaker is a mechanical device which shakes the pilot's controls to warn of the onset of stall. A stall warning is an electronic or mechanical device which sounds an audible warning and/or illuminates a light as the stall speed is approached. The majority of aircraft contain some form of this device that warns the pilot of an impending stall. An Angle-Of-Attack (AOA) Indicator, also called a Lift Reserve Indicator, is a pressure differential instrument that integrates airspeed and angle of attack into one instantaneous, continuous readout. An AOA indicator provides a visual display of the amount of available lift throughout its slow speed envelope regardless of the many variables which act upon an aircraft. This indicator is immediately responsive to changes in speed, angle of attack and wind conditions and automatically compensates for aircraft weight, altitude, and temperature. Page 57 of 66 09 June, 2016  AVAV4107 Flight Controls Handout An angle of attack limiter or an "alpha" limiter is a flight computer that automatically prevents pilot input from causing the plane to rise over the stall angle. Some alpha limiters can be disabled by the pilot. Stall warning systems often involve inputs from a broad range of sensors and systems to include a dedicated angle of attack sensor. Blockage, damage, or non-operation of stall and angle of attack (AOA) probes can lead to the stall warning becoming unreliable and cause the stick pusher, over speed warning, autopilot and yaw damper to malfunction. Angle of attach indicator (AOA) Maintenance Practices Balancing The reason for balancing of flight controls is to prevent the phenomenon of flutter. A flutter is caused by the movement of the control surface lagging behind the rise and fall of the forward portion of the wing as it flexes, thus tending to increase the oscillations. Mass balancing of the controls prevents this type of flutter. For ailerons, the positioning of the mass balance weight must be closer to the wing tip. Alternately, the weight can be distributed along the length of the aileron in the leading-edge. Page 58 of 66 09 June, 2016  AVAV4107 Flight Controls Handout There are many reasons for this, but the most common are excessive hinge gap or excessive "slop" in the pushrod connections and control horns. One of the most common causes of flutter is loose linkage between the servo and control surface. This is easily corrected but often overlooked or ignored. It only takes a moment to jiggle your control surfaces by hand to see how much movement there is and determine its origin. Servo gears get worn, clevises get worn, and control horns get worn. Flutter can occur at any time on any control surface if the conditions are correct. All are subject to flutter but the typical pecking order is: AILERONS, ELEVATOR, and then RUDDER. Mass balance weights are used when aerodynamic smoothness is required. It is generally required that the center of gravity of the control be located ahead of the hinge point for the control. All balancing is at the direction of the manufacturer. Page 59 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Page 60 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Balance horns are used to provide an aerodynamic balance for the ailerons, elevators, and rudders in which the area ahead of the hinge is concentrated on one part of the surface in the form of a horn. This horn produces a balancing moment, thus reducing the amount of force required to move the controls or the control’s CG (center of gravity) forward of the hinge to reduce the likelihood of control flutter. In addition to the factors mentioned above there are other variables to control surface balancing. One of these is painting the aircraft. If the control surfaces are to be painted they will have to re-weighed and balanced in accordance with the aircraft maintenance manual. Trapped water in composite structures, loose or vibrating trim tabs can all develop into flutter. When these units are repaired, care must be taken to ensure they are re-weighed and balanced. Page 61 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Rigging General Flight control rigging is important, when the flight controls are rigged correctly, the aircraft will perform correctly. The flight control surfaces will move as commanded by the pilot or auto-flight systems. As described previously, there are many different types of flight control operating systems, cables, push/pull rods, electrical, and hydraulically operated actuators. Cable operated flight control systems depend on the movement of wire cable to actuate the control surface. If cable tension is not correct the system will not operate correctly, too light and the cables may slip out of the pulley grooves, too heavy the controls will be difficult to move and there will be excessive wear in the control system. Airframes will change size with temperature and pressure; this will make getting the correct cable tension difficult. On large aircraft, cable tension regulators compensate for dimensional changes in the airframe and maintain a relatively constant cable tension. Others require that the tension be set according to the ambient temperature, calculating the value from a table provided by the manufacturer. Control cables are manufactured from carbon steel or corrosion resistant steel but all use corrosion resistant steel swage fittings. Flexible cable of 7x7 or extra flexible cable of 7x19 designations is most commonly used. These cables will meet either a MIL-SPEC or a manufactures specification. Page 62 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Aircraft Control Cable Types Page 63 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Cable Tensiometer Cable tensions are measured using a Cable Tensiometer. Several different brands of tensiometers are available. The style pictured above measures using different riser blocks for different size cables. The reading then has to be compared on a chart to see what the tension (in lbs.) is for that size cable. Rigid control Push-pull rod systems are used mainly on helicopters, and some fixed wing aircraft. The tubes are manufactured from heat treated seamless aluminum with threaded rod ends riveted into the ends. Push-pull rods are often mistakenly referred to as torque tubes. Rigging Procedures Using the applicable aircraft maintenance manual, adjust the cable tension, or push-rod length to achieve the correct control surface position. As all aircraft types are different it is mandatory that maintenance manuals are used. Some aircraft use inclinometers to position control surfaces; others Page 64 of 66 09 June, 2016  AVAV4107 Flight Controls Handout use direct measurement of the deflection of the control surface to set the correct rig position from an index plate. Still others use rig pins to position parts of the control system and adjustments are made in the length of the rods so that eye to eye length is correct. Typical Index Plate Hydraulic actuators are used to operate large flight control surfaces. These actuators have adjustable rod end or head end fittings. This allows the adjustment of the actuator when the control inputs have been adjusted, mechanical or electronic. In general flight control rigging is accomplished using the following; Rigging pins are used to ensure the controls to prevent movement during the rigging procedure. Adjust turnbuckles to reposition the control surface and provide the correct cable tension. Adjust push-pull tubes to reposition the control surface Verify control surface position by use of rigging boards, inclinometers, protractors, or tape measures Safety entire system by lock-wiring turnbuckles, cotter pinning castle nuts, checking self-locking nuts are in safety. Remove rigging pins. Verify control system movement to check that the control input loads are within the required tolerance. Page 65 of 66 09 June, 2016  AVAV4107 Flight Controls Handout Independent inspection provides a double check of the flight controls to ensure their correct operation. END Page 66 of 66 09 June, 2016