M o d u l e H e l i c o p t e r A e r o d y n a m i c s, S t r u c t u r e s a n d S y s t e m s

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1 M o d u l e H e l i c o p t e r A e r o d y n a m i c s, S t r u c t u r e s a n d S y s t e m s F l i g h t C o n t r o l S y s t e m s

2 T a b l e o f c o n t e n t s I. INTRODUCTION AXIS: Plane of rotation: Axis of rotation: Flight control components: PROPERTIES: MAIN PILOT COMMANDS: OTHER PILOT COMMANDS: II. CYCLIC CONTROL GENERAL: Tip path plane - TPP: Swashplates principle: FUNCTIONS OF CYCLIC CONTROL: Main functions of cyclic control: Additional functions: III. COLLECTIVE CONTROL COLLECTIVE CONTROL AND PITCH ANGLE: THROTTLE CONTROL: GOVERNOR AND CORRELATORS: IV. SWASHPLATE SWASHPLATE ASSEMBLY: Swashplate components: Swashplate description: SWASHPLATE OPERATION: Collective input operation: Cyclic operation: SPIDER CONTROL MECHANISM: V. YAW CONTROL ANTI-TORQUE CONTROL: TAIL ROTOR: Heading control: Tail rotor commands:

3 Fenestron tail rotor systems: Tail rotor control systems: TWIN NON-COAXIAL ROTORS: Tandem rotors: Side-by-side rotors: Twin intermeshing rotors: Twin coaxial rotors: BLEED AIR NOTAR HELICOPTERS: NOTAR concept: NOTAR system operation: VI. MAIN ROTOR HEAD DESIGN AND OPERATION FEATURES INTRODUCTION: COMPARISON OF MAIN ROTOR TYPES: Rigid rotor systems: Semirigid rotor systems: Fully articulated rotor system: Combination rotor system: Bearingless Main Rotors (BMR technology): VII. BLADE DAMPERS: FUNCTION AND CONSTRUCTION ROTOR VIBRATIONS: Low frequency vibrations: Medium and high frequency vibrations: HELICOPTER GROUND AND AIR RESONANCE: Ground resonance phenomenon: Air resonance phenomenon: Ground (and air) resonance solutions: DAMPER FUNCTION: Hydraulic dampers: Elastomeric dampers: Variable dampers: Fluid filled dampers: DAMPER CONTRUCTION: Hydraulic dampers: Elastomeric inertial dampers: Embedded inertial dampers: VIII. ROTOR BLADES: MAIN AND TAIL ROTOR BLADE CONSTRUCTION AND ATTACHMENT GENERAL:... 91

4 BLADE DESIGNS: BLADE STRUCTURES AND MATERIALS: Forces acting on blades: Types of blades: Wood blades: Metal blades: Composite blades: NACA design airfoils: ONERA design airfoils: ROTOR TRANSMISSION: Main purposes: Clutch: Belt drive clutch: Centrifugal clutch: MAIN ROTOR HEADS: Hooke joint teetering head: Bell 206 teetering head: Tri-hinge head: Door-hinge hub: Semi-teetering head with elastomeric element: MAIN ROTOR CONSTRUCTION AND ATTACHMENT: Rotor head of the BO-105: Rotor head of Lockheed AH-56 Cheyenne: Rotor head of Westland Lynx: Rotor head of Bell-412: Rotor head of EC-725: Sikorsky S-64 Skycrane example: TAIL ROTOR CONSTRUCTION AND ATTACHEMENT: Antitorque systems: Antitorque fenestron systems: Robinson R22 antitorque example: IX. TRIM CONTROL, FIXED AND ADJUSTABLE STABILIZERS TRIM CONTROL: Introduction: Magnetic brakes: Electrical trim actuators: FIXED AND ADJUSTABLE STABILIZERS:

5 2.1. Fixed stabilizers: Adjustable stabilizers: X. SYSTEM OPERATION: MANUAL, HYDRAULIC, ELECTRICAL AND FLY-BY-WIRE MANUAL SYSTEM OPERATION: Introduction: Screw jacks (or jackscrews): Cables, pulleys, chains and push rods: HYDRAULIC SYSTEM OPERATION: Hydro-mechanical: ELECTRO-MECANICALSYSTEM OPERATION (FLY-BY-WIRE): XI. ARTIFICIAL FEEL ARTIFICIAL FEEL SYSTEMS: A simple spring feel units: Q feel units: MAIN ROTOR CONTROL SYSTEM FEEL GRADIENT UNITS: TAIL ROTOR CONTROL SYSTEM FEEL YAW PEDAL DAMPERS: FEEL INTEGRATION WITH AUTOFLIGHT SYSTEMS: XII. BALANCING AND RIGGING MAIN ROTOR BLADE ALIGNMENT: UNBALANCED SEMIRIGID ROTOR SYSTEMS: Lateral: Chordwise: Spanwise: Combined: VERTICAL VIBRATIONS: Extreme low frequency vibrations: Low frequency vibrations: Medium frequency vibrations: High frequency vibrations: MAIN ROTOR BLADE TRACKING: Electronic blade tracker: Reflector tracking: Strobe light:

6 I. INTRODUCTION 1. AXIS: The flight control systems allow the operation on the three axis device: Roll control (longitudinal axis), Pitch control (lateral axis), and Yaw control or direction (vertical axis)

7 1.1. Plane of rotation: A plane formed by the average tip path of the blades is known as the plane of rotation. The plane of rotation is at a right angle to the axis of rotation Axis of rotation: An imaginary line that passes through a point on which a body rotates is called the axis of rotation. Its rotation is at a right angle to the plane of rotation. Disc area (A): The span length of one blade is used as the radius. The area of the hub in the disc area is not included since it doesn't make any lift (but is negligible). The disc area is the total space in the area of the circle formed by the rotating rotor blades. The following formula is used to figure disc area:

8 where: A: disc area, π = , R: radius. 2 A = πr 1.3. Flight control components: Flight control on helicopters is obtained by changing the configuration of the main and tail rotors. The lift force is generated by the blades of the main rotor that are usually in a number between two and six. To manoeuvre a helicopter three controls are used; a collective pitch lever, cyclic pitch stick and yaw pedals

9 Collective pitch lever: Movement of the collective pitch lever will increase or decrease the pitch angle of all main rotor blades, by the same amount and at the same time. Increasing the pitch on all main rotor blades will increase the total rotor thrust, and decreasing it will have the opposite effect. Cyclic pitch lever: The cyclic pitch stick is used to tilt the main rotor disc, forwards, backwards and to the left or the right, or some combination of these. This will provide a thrust in the direction in which the disc is tilted, and will cause the helicopter to move in that direction. The tilting of the main rotor disc is achieved by independently adjusting the pitch on individual rotor blades causing them to move upwards or downwards. When cyclic pitch inputs are made the main rotor blades will be subject to an increase or decrease in their pitch angle as they rotate, and so the disc remains tilted in the direction selected by the pilot. Governor/FADEC controls: It is normal for modern helicopter engines to remain at a fixed flight idle speed, which is controlled by a fuel governor or computer controlled FADEC (Full Authority Digital Engine Control system), although some older helicopters, and some of those below the 5700 Kg MTOM weight category, provide a hand throttle twist grip on the collective pitch lever. In governed or FADEC systems an increase or decrease in the power required is automatically achieved, in systems using a hand throttle it is necessary for the pilot to make adjustments to the engine RPM in response to control inputs, obviously the governed or FADEC systems are more accurate and relieve the pilot of the additional workload imposed by the need to maintain engine, and therefore rotor RPM. Yaw pedals: The yaw pedals increase the pitch angle of the tail rotor blades, collectively only, as tail rotors do not require cyclic pitch inputs. The tail rotor is used to cancel out the torque reaction caused by the main rotor. An increase in main rotor collective pitch will produce more torque, and will therefore require more thrust from the tail rotor to oppose the resultant torque reaction. In addition to allowing the pilot to counteract torque reaction, the yaw pedals provide a means by which the helicopter can yaw; the nose moves to left or right. To yaw the helicopter against the torque reaction more thrust will be required, therefore more pitch is applied to the tail rotor blades, producing the necessary thrust. To yaw the aircraft in the same direction as the torque reaction, it will merely be necessary to reduce the tail rotor pitch and allow the helicopter to be rotated by the torque reaction force. From this it can be seen that the helicopter controls are very interactive, for example in the hover an increase in main rotor collective pitch will cause an increase in torque, and therefore torque reaction and additional thrust will be required from the tail rotor to oppose any tendency for the torque reaction to yaw the helicopter.

10 Electronic systems: Many helicopters, other than the most basic types, will incorporate electronic systems within the basic control systems to provide automatic stabilising and, in larger type s automatic pilot. Autostab: Automatic stabilising systems, sometimes referred to as autostab, are used to ensure that the helicopter remains at a fixed height, heading and speed, regardless of any disturbing influences, such as wind gusts. These systems ensure that the helicopter remains stable, without the need for continuous inputs from the pilot, thereby reducing pilot workload and fatigue. Autopilot: Many of the larger helicopters have a full autopilot system, where a pre-selected flight plan can be flown with a minimum of inputs from the pilot, thereby further reducing cockpit workload

11 2. PROPERTIES: Flying control systems are regulated by the FARs, PARTs, and must comply with the following standards: Sense : The aircraft must move in the direction signified by the control input, e.g. control column back, and pitch nose-up. Rigidity: The control system must be strong enough to withstand any operating loads without excessive distortion, e.g. airloads on the control surfaces (irreversibility). Stability: The control surfaces must remain where selected by the pilot and must not be affected by signals which are not self initiated, e.g. vibration and aerodynamic loads. Safety : Passengers, cargo and loose articles must safeguard the control system against jamming, chafing, and interference. Guards must therefore be fitted where appropriate to provide the necessary protection. Fail-Safe : The control system must be duplicated or be capable of manual operation in the event of hydraulic power failure. 3. MAIN PILOT COMMANDS: Cyclic-pitch lever -- A helicopter pilot controls the pitch, or angle, of the rotor blades with two inputs: the cyclic- and collective-pitch levers, often just shortened to the cyclic and the collective. The cyclic, or "stick," comes out of the floor of the cockpit and sits between the pilot's legs, enabling a person to tilt the craft to either side or forward and backward. Collective-pitch lever -- The collective-pitch lever is responsible for up-and-down movements. For example, during takeoff, the pilot uses the collective-pitch lever to increase the pitch of all the rotor blades by the same amount. Foot pedals -- A pair of foot pedals controls the tail rotor. Working the pedals affects which way the helicopter points, so pushing the right pedal deflects the tail of the helicopter to the left and the nose to the right; the left pedal turns the nose to the left

12 4. OTHER PILOT COMMANDS: Tail boom -- The tail boom extends out from the rear of the fuselage and holds the tail rotor assemblies. In some models, the tail boom is nothing more than an aluminum frame. In others, it's a hollow carbon-fiber or aluminum tube. Anti-torque tail rotor -- Without a tail rotor, the main rotor of a helicopter simply spins the fuselage in the opposite direction. It's enough to make your stomach heave just thinking about all that endless circling. Thankfully, Igor Sikorsky had the idea to install a tail rotor to counter

13 this torque reaction and provide directional control. In twin-rotor helicopters, the torque produced by the rotation of the front rotor is offset by the torque produced by a counter rotating rear rotor. Landing skids -- Some helicopters have wheels, but most have skids, which are hollow tubes with no wheels or brakes. A few models have skids with two ground-handling wheels

14 II. CYCLIC CONTROL 1. GENERAL: The cyclic pitch control tilts the main rotor disc by changing the pitch angle of the rotor blades in their cycle of rotation. When the main rotor disc is tilted, the horizontal component of lift moves the helicopter in the direction of tilt. The rotor disc tilts in the direction that pressure is applied to the cyclic pitch control. If the cyclic is moved forward, the rotor disc tilts forward; if the cyclic is moved aft, the disc tilts aft, and so on. Cyclic pitch control location

15 The cyclic pitch stick is positioned centrally in front of the pilot and co-pilots seats, and is used to tilt the disc, causing the helicopter to move horizontally in any direction. The cyclic can pivot in all directions. The cyclic pitch stick is used to tilt the main rotor disc, forwards, backwards and to the left or the right, or some combination of these. This will provide a thrust in the direction in which the disc is tilted, and will cause the helicopter to move in that direction. The tilting of the main rotor disc is achieved by independently adjusting the pitch on individual rotor blades causing them to move upwards or downwards. When cyclic pitch inputs are made the main rotor blades will be subject to an increase or decrease in their pitch angle as they rotate, and so the disc remains tilted in the direction selected by the pilot Tip path plane - TPP: The tip path plane, or TPP, is the plane connecting the rotor blade tips as they rotate. While hovering, the thrust vector of a helicopter is oriented upwards, perpendicular to the tip path plane. In order for the helicopter to travel forward, this thrust vector needs to be rotated slightly in the forward direction

16 1.2. Swashplates principle: Like all other flight controls it operation is purely instinctive, moving the cyclic pitch stick forwards will tilt the disc forwards, and the helicopter will move forwards, moving the cyclic stick rearwards has the opposite effect. Cyclic control and swashplate

17 Since tilting the rotor hub or rotor shaft is impractical, an alternative means of rotating the tip path plane - TPP - is needed. Most modern helicopters use a system of swashplates. Seen in the following diagram, the swashplate system is composed of upper and lower swashplates. The brown portion of the diagram, including the lower swashplate, remains stationary relative to the helicopter. The upper swashplate (in grey) rotates with the rotor, while remaining parallel to the lower swashplate. By utilizing what is called cyclic control, the swashplates can be angled so as to vary the pitch of the blades depending on their azimuth angle. As the swashplates are tilted in the proper direction, there is an increased lift on the aft portion of the rotor, causing the blades to flap up, which in turn causes the TPP to rotate forwards. As the TPP rotates forwards, the thrust does as well, imparting a forward acceleration to the helicopter. 2. FUNCTIONS OF CYCLIC CONTROL: 2.1. Main functions of cyclic control: The rotor disc tilts in the direction that pressure is applied to the cyclic pitch control. If the cyclic is moved forward, the rotor disc tilts forward; if the cyclic is moved aft, the disc tilts aft, and so on. Because the rotor disc acts like a gyro, the mechanical linkages for the cyclic control rods are rigged in such a way that they decrease the pitch angle of the rotor blade approximately 90 before it reaches the direction of cyclic displacement, and increase the pitch angle of the rotor blade approximately 90 after it passes the direction of displacement: an increase in pitch angle increases angle of attack; a decrease in pitch angle decreases angle of attack. For example, if the cyclic is moved forward, the angle of attack decreases as the rotor blade passes the right side of the helicopter and increases on the left side. This results in maximum downward deflection of the rotor blade in front of the helicopter and maximum upward deflection behind it, causing the rotor disc to tilt forward. Movement of the cyclic stick to left or right will cause the helicopter to move in that direction. The cyclic pitch stick is pivoted at its lower end and is connected to two push/pull tubes, one transmitting left/right (roll) movements and the other fore and aft (pitch) movements. A yoke assembly allows these movements to be made independently so that only roll or pitch inputs may be made without causing movement of the other, however, simultaneous roll and pitch movements can be made if required. In fore and aft movement, one push/pull tube transmits movements to the control mixing unit

18 Side to side movements operate two push/pull rods which operate in opposite directions, when the cyclic stick is moved to the left one rod will move forwards and the other rearwards, when the stick is moved to the right the opposite will occur. This is required because there are two lateral (roll) main rotor actuators which must operate in opposition to achieve lateral control, whereas, fore and aft pitch movements are achieved by only one main rotor actuator, that uses the fixed or non-rotating scissors as a datum point about which movements are made. Cyclic inputs on stationary and rotating swashplate

19 Cyclic transmission principle

20 The action is carry out thanks to the cyclic pitch stick

21 The action is carry out thanks to the cyclic pitch stick Note: On some helicopters, the control rods were routed internally up through the main rotor mast to protect them. On those helicopters, the cyclic inputs come down from the top of the mast and the swashplate is under the transmission, where it is all covered and protected from wires (Enstrom)

22 2.2. Additional functions: The cyclic stick grip will incorporate switches for operation of important systems; these are normally associated with control trimming, autoflight or autostabilisation, cargo release and communication systems

23 III. COLLECTIVE CONTROL The collective pitch control (or simply collective or thrust lever ), located on the left side of the pilot s seat, changes the pitch angle of all main rotor blades simultaneously, or collectively, as the name implies. Raising the collective pitch control increases the pitch angle the same amount on all blades

24 1. COLLECTIVE CONTROL AND PITCH ANGLE: The collective pitch lever or stick is located by the left side of the pilot's seat and is operated with the left hand. This lever moves up and down pivoting about the aft end and, through a series of mechanical linkages, changes the pitch angle of the main rotor blades. The collective is used to make changes to the pitch angle of the main rotor blades and does this simultaneously, or collectively, as the name implies. As the collective pitch control is raised, there is a simultaneous and equal increase in pitch angle of all main rotor blades; as it is lowered, there is a simultaneous and equal decrease in pitch angle. This is done through a series of mechanical linkages and the amount of movement in the collective lever determines the amount of blade pitch change. An adjustable friction control helps prevent inadvertent collective pitch movement. Changing the pitch angle on the blades changes the angle of attack on each blade. With a change in angle of attack comes a change in drag, which affects the RPM of the main rotor. As the pitch angle increases, angle of attack increases, drag increases and rotor RPM decreases. Decreasing pitch angle decreases both angle of attack and drag, while rotor RPM increases. In order to maintain a constant rotor RPM, which is essential in helicopter operations, a proportionate change in power is required to compensate for the change in drag. This is accomplished with the throttle control or a governor, which automatically adjusts engine power. 2. THROTTLE CONTROL: The function of the throttle is to regulate engine RPM. If the governor system does not maintain the desired RPM when the collective is raised or lowered, or if those systems are not installed, the throttle must be moved manually with the twist grip in order to maintain RPM. The throttle control is much like a motorcycle throttle, and works in virtually the same way. Twisting the throttle to the left increases RPM; twisting the throttle to the right decreases RPM. Note: A twist grip throttle is usually mounted on the end of the collective lever. The throttles on some turbine helicopters are mounted on the overhead panel or on the floor in the cockpit

25 Twist grip throttle 3. GOVERNOR AND CORRELATORS: A governor is a sensing device that senses rotor and engine RPM and makes the necessary adjustments in order to keep rotor RPM constant. In normal operations, once the rotor RPM is set, the governor keeps the RPM constant, and there is no need to make any throttle adjustments. Governors are common on all turbine helicopters (as it is a function of the fuel control system of the turbine engine), and used on some piston powered helicopters. A correlator is a mechanical connection between the collective lever and the engine throttle. When the collective lever is raised, power is automatically increased; when lowered, power is decreased. This system maintains RPM close to the desired value, but still requires adjustment of the throttle for fine tuning. Some helicopters do not have governors and require coordination of all collective and throttle movements. When the collective is raised, the throttle must be increased; when the collective is lowered, the throttle must be decreased

26 As with any aircraft control, large adjustments of either collective pitch or throttle should be avoided. All corrections should be made through the use of smooth pressure. In piston helicopters, the collective pitch is the primary control for manifold pressure, and the throttle is the primary control for RPM. However, the collective pitch control also influences RPM, and the throttle also influences manifold pressure; therefore, each is considered to be a secondary control of the other s function. Both the tachometer (RPM indicator) and the manifold pressure gauge must be analyzed to determine which control to use. If the RPM is and the manifold pressure is Solution Increasing the throttle increases manifold pressure and RPM. Lowering the collective pitch decreases manifold pressure and increases RPM. Raising the collective pitch increases manifold pressure and decreases RPM. Reducing the throttle decreases manifold pressure and RPM. Relationship between RPM, manifold pressure, collective, and throttle

27 Collective transmission principle

28 Collective inputs on a stationary and rotating swashplate

29 IV. SWASHPLATE A major part of a helicopter flight control system is the mechanism used to transfer control inputs from the non-rotating parts of the system, to those that are rotating. There are 2 commonly used methods used to achieve this: Swashplate, Spider control. Of these, the swashplate is perhaps the most common device used for main rotors, and the spider for tail rotors, although a there are some helicopters that use spider control for main rotors too. 1. SWASHPLATE ASSEMBLY: The purpose of the swashplate is to transmit control inputs from the collective and cyclic controls to the main rotor blades. It consists of two main parts: the stationary swashplate and the rotating swashplate. The swashplate consists of two primary elements through which the rotor mast passes. The swashplate includes a rotating and non-rotating plate, normally referred to as stars because of their shape, they may be made from steel, titanium or light alloy, with the choice of material being dependent upon the inservice loads that the swashplate will have to withstand. The non-rotating star is mounted onto the main rotor gearbox shaft by a large spherical ball, housed in its centre, sometimes referred to as a uniball. This ball is free to run up and down a slider sleeve, fitted around the shaft. This non-rotating disc, often referred to as the "stationary star," is attached by a bearing surface to a second disc, often referred to as the "rotating star," which turns with the rotor and is mechanically linked to the rotor blade pitch horns. The rotating star is fitted to the non-rotating star on a bearing, and so can freely rotate about it. The rotating and non-rotating swashplates move as a single entity in the horizontal plane, and any change in the horizontal angle of the non-rotating swashplate, will be transferred to the rotating swashplate, which will move to the same angle. The rotor blade pitch horns are placed approximately 90 ahead of or behind the blade on which they control the pitch change

30 If this were not done, gyroscopic precession would cause the movement of the helicopter to be 90 out of phase with the movement of the cyclic pitch stick, that is, if the cyclic stick were displaced to the right, the helicopter would move forward; if the cyclic stick were displaced forward, the helicopter would move to the left, and so on. Swashplate detail on Bell UH

31 1.1. Swashplate components: Swashplate components In this figure, collective and cyclic control inputs are transmitted to the stationary swashplate by control rods causing it to tilt or to slide vertically. The pitch links attached from the rotating swashplate to the pitch horns on the rotor hub transmit these movements to the blades Swashplate description: The stationary swashplate is mounted around the main rotor mast and connected to the cyclic and collective controls by a series of pushrods. It is restrained from rotating but is able to tilt in all directions and move vertically. The rotating swashplate is mounted to the stationary swashplate by means of a bearing and is allowed to rotate with the main rotor mast. Both swashplates tilt and slide up and down as one unit. The rotating swashplate is connected to the pitch horns by the pitch links. The purpose of the swashplate is to transmit control inputs from the collective and cyclic controls to the main rotor blades. It consists of two main parts: the stationary swashplate and the rotating swashplate

32 The stationary swashplate is mounted around the main rotor mast and connected to the cyclic and collective controls by a series of pushrods. It is restrained from rotating but is able to tilt in all directions and move vertically. The rotating swashplate is mounted to the stationary swashplate by means of a bearing and is allowed to rotate with the main rotor mast. Both swashplates tilt and slide up and down as one unit. The rotating swashplate is connected to the pitch horns by the pitch links

33 Swashplate cross-section details Recall: Collective and cyclic control inputs are transmitted to the stationary swashplate by control rods causing it to tilt or to slide vertically. The pitch links attached from the rotating swashplate to the pitch horns on the rotor hub transmit these movements to the blades

34

35 2. SWASHPLATE OPERATION: Regardless of actual arrangement and manufacturers preference all swash plates transfer control inputs to the main rotor in the same manner, description of the operating sequence is most clearly achieved by use of simple illustrations Collective input operation: When a collective input is made all 3 actuators will extend or retract by the same amount. The movement of the actuators will raise or lower the swash plate assembly, causing the spherical bearing at its centre to move up or down on the slider shaft. The rotor blades will get a simultaneous and equal collective pitch change, as illustrated, increasing or decreasing total rotor thrust. The illustration shows this effect when the collective pitch lever is raised, and there is an equal and simultaneous pitch increase on all rotor blades

36 2.2. Cyclic operation: When the cyclic pitch stick is moved a differential input will be made at the mixing unit and this will be transmitted to the actuators, causing the actuators to rise or lower. The non-rotating star will tilt, and this action is passed on to the main rotor blades, changing their pitch independently causing the rotor disc to tilt because of phase lag the tilt of the swash plate and rotor disc will not be in the same direction. As the rotating star rotates, there will be a continuous change in blade pitch depending upon the cyclic pitch stick position

37 The illustration shows the effect when there is a change in fore and aft (pitching) movement, in which case the non-rotating scissors acts as a datum point, and the fore and aft actuator moves the swash plate about it. In the case of a lateral (rolling) input one lateral actuator will rise and the other will fall. 3. SPIDER CONTROL MECHANISM: Although this type of control layout varies considerably from the swashplate system it performs exactly the same functions and the controls in the cockpit are the same. There is a separate control run for each of the primary movements of the pilot s; a cyclic control for lateral movement and cyclic control for fore and aft movement and a collective control. In a typical system each of these is connected separately to a control beam thus eliminating the need for a mixer unit to give collective control, but it does require a special type of control linkage to achieve cyclic movement. Blade pitch changes are affected by a vertical spindle in a sliding sleeve connected to the blades by spider arms, the spindle and sleeve moving inside the hollow rotor shaft from the gear box. As the vertical spindle is connected to the blades that are rotating, it follows that the vertical spindle is rotating within the rotor drive shaft; therefore to achieve a link-up of the rotating side of the system and the non-rotating side there is a rotating/non-rotating link with universal movement to give both collective and cyclic movement. (See chapter VIII Rotor blades)

38 V. YAW CONTROL Helicopters fly and move horizontally using one or more sets of spinning rotor blades. Each of these blades is an airfoil shape which, when moving through the air, force air downwards and in doing so are themselves forced upwards. There are a number of ways that this torque force is counteracted in rotor design: Tail rotor; Twin non-coaxial rotors; Contra-rotating coaxial rotors; Bleed air NOTAR (NO TAil Rotor). 1. ANTI-TORQUE CONTROL: The anti-torque pedals, located on the cabin floor by the pilot s feet, control the pitch, and therefore the thrust, of the tail rotor blades. The main purpose of the tail rotor is to counteract the torque effect of the main rotor. Anti-torque pedals Rewind about 350 years to Newton and his third law of motion: for every action there is an equal and opposite reaction. A helicopter creates lift by spinning its main rotor blade(s) in one direction; this action has an equal and opposite effect on the helicopter fuselage, making it 'want' to spin in the opposite

39 direction. Since the torque on the aircraft varies with changes in engine power used by the main rotor, the tail rotor thrust must also be varied. The pedals are connected to the pitch change mechanism on the tail rotor gearbox and allow the pitch angle on the tail rotor blades to be increased or decreased. Besides counteracting torque of the main rotor, the tail rotor is also used to control the heading of the helicopter while hovering or when making hovering turns. Hovering turns are commonly referred to as pedal turns. This Newton law applies to the helicopter fuselage and its rotation in the opposite direction of the main rotor blades unless counteracted and controlled. To make flight possible and to compensate for this torque, most helicopter designs incorporate an anti-torque rotor or tail rotor. The anti-torque pedals allow the pilot to control the pitch angle of the tail rotor blades, which in forward flight puts the helicopter in longitudinal trim and, while at a hover, enables the pilot to turn the helicopter 360. The anti-torque pedals are connected to the pitch change mechanism on the tail rotor gearbox and allow the pitch angle on the tail rotor blades to be increased or decreased. Since the torque on the aircraft varies with changes in engine power used by the main rotor, the tail rotor thrust must also be varied. The pedals are connected to the pitch change mechanism on the tail rotor gearbox and allow the pitch angle on the tail rotor blades to be increased or decreased. 2. TAIL ROTOR: The most common configuration is the combination of one main rotor and one tail rotor. The tail rotor will compensate the torque which is produced by the main rotor. The tail rotor is also responsible for the control of the helicopter along the vertical axis, during hover flight Heading control: The thrust of the tail rotor depends on the pitch angle of the tail rotor blades. This pitch angle can be positive, negative, or zero. A positive pitch angle tends to move the tail to the right. A negative pitch angle moves the tail to the left, while no thrust is produced with a zero pitch angle. With the right pedal moved forward of the neutral position, the tail rotor either has a negative pitch angle or a small positive pitch angle. The farther forward the right pedal is displaced, the larger the negative pitch angle. The closer the right pedal is to the neutral position, the more positive the pitch angle, and somewhere in between, it has a zero pitch angle. As the left pedal is moved forward of the neutral position, the positive pitch angle of the tail rotor increases until it becomes maximum with full forward displacement of the left pedal. If the tail rotor has a negative pitch angle, tail rotor thrust is working in the same direction as the torque of the main rotor. With a small positive pitch angle, the tail rotor does not produce sufficient thrust to overcome the torque effect of the main rotor during cruise flight. Therefore, if the right pedal is displaced forward of neutral during cruising flight, the tail rotor thrust does not overcome the torque effect, and the nose yaws to the right. With the anti-torque pedals in the neutral position, the tail rotor has a medium positive pitch angle. In medium positive pitch, the tail rotor thrust approximately equals the torque of the main rotor during cruise flight, so the helicopter maintains a constant heading in level flight. The tail rotor is used to control the heading of the helicopter while hovering or when making hovering turns, as well as counteracting the torque of the main rotor

40 If the left pedal is in a forward position, the tail rotor has a high positive pitch position. In this position, tail rotor thrust exceeds the thrust needed to overcome torque effect during cruising flight so the helicopter yaws to the left. The above explanation is based on cruise power and airspeed. Since the amount of torque is dependent on the amount of engine power being supplied to the main rotor, the relative positions of the pedals required to counteract torque depend upon the amount of power being used at any time. The maximum positive pitch angle of the tail rotor is generally somewhat greater than the maximum negative pitch angle available. This is because the primary purpose of the tail rotor is to counteract the torque of the main rotor

41 The capability for tail rotors to produce thrust to the left (negative pitch angle) is necessary, because during autorotation the drag of the transmission causes the nose to yaw to the left, or in the same direction the main rotor is turning. Tail rotor configuration By far the most common method for yaw control was first used by Igor Sikorsky in Sikorsky's configuration featured a single main rotor with a small tail rotor mounted at the end of a long tail boom, as exemplified by the AH-64 Apache. While the main rotor generates the lift and thrust that make flight possible, the purpose of the tail rotor is to counteract the torque effect. The tail rotor works much like any propeller or rotor. It spins at a high rate of speed, accelerating air in a specific direction, to create a force in the opposite direction that counteracts the force of the main rotor. The pilot of the helicopter can alter the properties of the tail rotor to increase or decrease this force thereby causing the helicopter to yaw to either the left or the right. To maximize the force of the tail rotor, the rotor is usually mounted on a considerably long tail boom. There are two notable disadvantages of this arrangement: First, the long tail boom creates an aircraft with a large "footprint," meaning it needs more space for manoeuvring and storage. Second, while the helicopter is on the ground, the spinning tail rotor is usually low to the ground so that it presents a hazard, often a fatal one, to ground crew Tail rotor commands:

42 From the neutral position, applying right pedal causes the nose of the helicopter to yaw right and the tail to swing to the left. Pressing on the left pedal has the opposite effect: the nose of the helicopter yaws to the left and the tail swings right. With the anti-torque pedals in the neutral position, the tail rotor has a medium positive pitch angle. In medium positive pitch, the tail rotor thrust approximately equals the torque of the main rotor during cruise flight, so the helicopter maintains a constant heading in level flight. A vertical fin or

43 stabilizer is used in many single-rotor helicopters to help aid in heading control. The fin is designed to optimize directional stability in flight with a zero tail rotor thrust setting. The size of the fin is crucial to this design. If the surface is too large, the tail rotor thrust may be blocked. Helicopters that are designed with tandem rotors do not have an anti-torque rotor. These helicopters are designed with both rotor systems rotating in opposite directions to counteract the torque rather than a tail rotor Fenestron tail rotor systems: The fenestron tail rotor is another relatively new technique that was patented by Eurocopter. The fenestron, pictured below, is essentially the same concept as a conventional tail rotor. Both systems feature spinning blades that generate a thrust force to cancel out the tendency of a helicopter fuselage to rotate. Typical tail rotor fenestron However, the fenestron rotor differs from a conventional rotor by adding several more blades. Whereas a conventional tail rotor seldom has more than four blades, a typical fenestron includes eight to thirteen blades. Compared to conventional tail rotor blades, the fenestron blades are also much smaller and spin at higher speeds. Furthermore, these blades are mounted within a shroud that forms part of the vertical tail fin of the helicopter

44 This configuration turns the rotor into a ducted fan whose blade tips are protected from the external air. The only significant drawback to ducted fans like the fenestron is that the shroud adds weight that offsets at least some of the improvements in performance. Regardless of the weight penalty, noise reduction was one of the primary reasons a ducted fan system was adopted for the now-cancelled RAH-66 Comanche stealth helicopter pictured above. RAH-66 Comanche The primary advantage of this ducted fan arrangement is to reduce the turbulence and vortex shedding that occurs on rotor and propeller blades. In so doing, the rotor becomes more aerodynamically efficient by reducing drag, and noise as well as vibration is also significantly reduced. The fenestron offers safety advantages too since the shroud helps protect the rotor from striking outside objects

45 Tail rotor strikes against trees, power lines, and other obstructions are one of the most common causes of helicopter crashes, so reducing the rotor's vulnerability to damage is important. The shroud also reduces the danger tail rotors have traditionally posed to ground crew operating near helicopters during takeoff and landing. Since the majority of the noise generated by a helicopter tends to come from the tail rotor, designers adopted a ducted fan to reduce noise levels and make the helicopter more difficult to detect. This rotor system was essentially a copy of the French Fenestron yet was referred to by the name Fantail in the US Tail rotor control systems: Tail rotor control systems can be of two types: cable systems or push/pull tube systems. In earlier helicopters cable systems were used because of the need to reduce the weight of the longer system, cable systems offering a weight saving of approximately 25-30% over tube systems. However, cable systems suffer from several disadvantages, they require strengthening of the structure because of the relatively high cable tensions, and the steel cables will expand and contract at a different rate to the light alloy structure. Whilst the latter was overcome with the use of cable tension regulators, cable systems still required more maintenance, and were prone to developing faults. Many modern helicopters, especially the larger types, now use push/pull tube systems for tail rotor control. Cable systems: The majority of cable systems use push/pull tubes from the yaw pedals to a cable quadrant, from here control cables are used to transfer control inputs through the fuselage and tail boom structure. In very early helicopters the cables were wound around a cable spool, usually 1½ to 2½ turns, which was connected directly to a mechanical screw-jack that turned the motion through 90 and provided the movement to the tail rotor, via a spider mechanism, although this provided a purely manual control system and was generally only used on the light helicopter types. An alternative to this, especially where hydraulic controls were used, was to position another quadrant just before the tail rotor, and connect it via a push/pull tube to the tail rotor control mechanism, or hydraulic actuator. One of the quadrants would be a cable tension regulator, ensuring consistent cable tensions. Cables used in a system that used a cable spool could either be of the continuous loop type or would have nipples swaged on to the ends of the cable. Those systems using 2 quadrants would comprise of 2 cables, having swaged nipples at each end to ensure positive retention at the quadrants. Push/pull tube systems:

46 In this system, the tubes transfer control inputs from the yaw pedals right through to the tail rotor control mechanism or actuator. Push/pull tube systems are more difficult to route than cables, and are comprised of many more components, many of which could potentially develop faults, but there is less possibility of lost motion developing rapidly within the system, as could be the case if control cables lost tension. 3. TWIN NON-COAXIAL ROTORS: 3.1. Tandem rotors: If you are going to build a large helicopter for lifting heavy loads, then it makes sense to use two separate rotors to generate more lifting force. If these rotors are separate, such as those on the CH-46 Sea Knight, then they are called "non-coaxial," meaning they do not share the same axis. Tandem twin non-coaxial rotors The arrangement as a tandem rotor is mainly used with big helicopters. Because of the opposite rotation of the rotors, the torque of each single rotor will be neutralized. The construction of the control system is much more complicated, compared to a helicopter with a tail rotor. The control along the vertical axis during hover flight is done by bending the rotor discs against each other

47 Tandem CH-46 Sea Knight helicopter Since these helicopters have two rotors that are usually equal in size, the simple solution to the yaw control problem is to spin the rotors in opposite directions. Thus, the yaw forces created by each rotor cancel each other out. Yaw control is achieved by tilting the rotors in opposite directions (to the left and right in the case of the CH-46). The two rotors are linked by a transmission that ensures the rotors are synchronized and do not hit each other, even during an engine failure. Tandem rotor designs achieve yaw by applying opposite left and right cyclic to each rotor, effectively pulling both ends of the helicopter in opposite directions. To achieve pitch, opposite collective is applied to each rotor; decreasing the lift produced at one end, while increasing lift at the opposite end, effectively tilting the helicopter forward or back. Advantages of the tandem-rotor system are a larger centre-of-gravity range and good longitudinal stability. Disadvantages of the tandem-rotor system are a complex transmission and the need for two large rotors Side-by-side rotors: The arrangement of two rotors side by side was never very popular. This is a large, twin-rotor cargo helicopter with 4 seats and plenty of space for hauling debris, landed capsules, rovers, spacecraft and whatnot. It is based on the real largest helicopter ever built

48 3.3. Twin intermeshing rotors: MIL MI-12 Like the tandem rotor, this configuration doesn't need a tail rotor, because the torque is compensated by the opposite rotation of the rotors. This system was developed during the early days of helicopter flying, but fell into oblivion

49 The principles by which intermeshing rotors operate are the same as those previously discussed for other twin rotor helicopters. The difference is that the rotors are mounted very close together and actually intermesh; the blades on one rotor travel through the rotor disk of the other rotor. Of course, the intermeshing of the rotors is carefully timed to keep the rotors from chopping each other to pieces. Twin intermeshing rotors Today this kind of arrangement of the rotor has been rediscovered and used with the Kaman K-MAX, a single seat helicopter, mainly used for external load transportation. A subgroup of twin non-coaxial rotors is the twin intermeshing rotor system, seen only on a few helicopters made by the Kaman Company, including the H-43 Husky and K-MAX

50 H-43 Husky helicopter 3.4. Twin coaxial rotors: Another yaw control strategy pioneered by the Russian manufacturer Kamov utilizes two rotors that share the same axis (coaxial), an example being the Ka-50. Much like the non-coaxial design, the rotors rotate in opposite directions, each one counteracting the yaw force of the other. Yaw control is achieved by increasing the blade pitch on one rotor while decreasing the pitch on the other. The result is a differential in torque, resulting in a yawing motion. The most notable advantage of the twin coaxial arrangement is that it is very compact

51 Although the twin rotor configuration tends to be rather tall, the lack of a long tail boom results in a very short fuselage that takes up much less space. For this reason, Kamov designs have proven very popular for shipboard use with the Russian Navy. Twin coaxial rotor configuration The control along the vertical axe occurs as a result of different lifts of the two rotor discs. Depending on which rotor produces more lift, the helicopter will turn to the left or right, because of the torque. For these helicopters it is not possible t o reach a high cruising speed, because the drag is too large. Only after the development of the rigid rotor, has it been possible to build the two rotors closer together and reduce the drag considerably

52 Kamov KA

53 4. BLEED AIR NOTAR HELICOPTERS: 4.1. NOTAR concept: NOTAR system circulation control

54 By far the most recent yaw control strategy developed for helicopters is the NO TAil Rotor (NOTAR) design. NOTAR was first developed by McDonnell Douglas during the 1980s and applied to a modified OH-6 Cayuse. The concept proved so successful that the company began marketing dedicated NOTAR spin-offs of the successful MD 500 family including the MD 520N. The NOTAR system utilizes boundary layer control to provide anti-torque from the main rotor. Two circulation control slots run along the starboard side of the tailboom, shown in figure. The result is that the tail boom acts like a wing flying in the downwash of the main rotor. This configuration produces 70 percent of the anti-torque in the hovering flight condition. Directional control is obtained using a rotating thruster and twin rudders at the end of the tail boom. In forward flight, the vertical stabilizers provide the majority of anti-torque, with the thruster once again used for directional control. Pilot inputs control the thruster nozzle and left vertical stabilizer. The right vertical stabilizer moves under computer control from a gyro that senses the angular rotations of the aircraft in flight. A variable-pitch fan pressurizes the tailboom with a low pressure high volume of air. The fan is driven by a short shaft from the main engine. Consequently, the entire NOTAR assembly is comparable to a conventional tail rotor design. Though the NOTAR arrangement still requires a long tail boom, the need for a spinning tail rotor is eliminated. This reduces the danger to ground crew, and also allows the pilot to maneuver into positions that he or she normally would not even consider. For example, NOTAR allows a pilot to stick the tail boom into a tree. Try that with a standard tail rotor, and it will be time for an unscheduled landing. We hope that this in-depth discussion satisfies any questions you might have about helicopter tail rotors. Of course, someone might come up with a better, simpler way of counteracting the force of the main rotor in a few years. Then again, someone might perfect flying saucer technology too, making helicopters a thing of the past. As for me, I'll put my money on helicopters being around for a while. The NOTARs large screened air inlet to the fan causes reduced interaction effects than the conventional tail design since it is located near the rotor hub and subsequently has lower power requirements. For helicopters with tail rotors that sit at or above the main rotor (which is the case for many larger helicopters), the difference in power requirements between the NOTAR design and the conventional design is negligible since the tail rotor now has a reduced interaction effect with the main rotor NOTAR system operation: The pilot and co-pilot yaw pedals control both the position of the jet-thruster and the pitch angle of the ducted fan, thereby allowing the correct amount of anti-torque force to be applied, without drawing too much power from the transmission system, as would be the case if the ducted fan was set at constant pitch and speed

55 MD-500 helicopter Additionally, the yaw pedals also control the angle of the vertical stabilisers fitted to the rear of the tail boom, which have a maximum deflection of 29 degrees, which unload the jet-thruster and ducted fan during forward flight, making more power available for the main rotor, with a reduction in fuel consumption

56 In the hover, the Coanda effect produced by the circulation control slots provides the majority of the anti-torque force, whilst during forward flight this is produced by the vertical stabilisers and jet-thruster. During autorotation with the engine shut down directional control is provided by the vertical stabilisers. Ducted fan: A large ducted fan, driven by the main rotor transmission, is mounted in the rear of the fuselage; this provides a flow of low pressure air through a large diameter hollow tail boom, which is made from composite materials. On the end of the tail boom is a variable jet-thruster controlled by the pilot and co-pilot yaw pedals, this thrusters opens and closes in response to control inputs to provide a greater or lesser anti-torque force. Coanda effect: The large diameter hollow tail boom has 2 slots manufactured into its side, known as circulation control slots. A percentage of the air flowing through the boom exits through these slots and causes the air flowing in the downwash from the main rotor to adhere to one side longer than the other, in effect producing a vertical aerofoil section, and thus producing a side force to oppose torque reaction

57 VI. MAIN ROTOR HEAD DESIGN AND OPERATION FEATURES 1. INTRODUCTION: The rotor system is the rotating part of a helicopter which generates lift. The rotor consists of a mast, hub, and rotor blades. The mast is a hollow cylindrical metal shaft which extends upwards from and is driven and sometimes supported by the transmission. At the top of the mast is the attachment point for the rotor blades called the hub. The rotor blades are then attached to the hub by any number of different methods. Main rotor systems are classified according to how the main rotor blades are attached and move relative to the main rotor hub. There are three basic classifications: semirigid, rigid, or fully articulated. Some modern rotor systems, such as the bearingless rotor system, use an engineered combination of these types. The choice of which main rotor head type is used on individual helicopter types is made by consideration of several factors, which is mainly based upon aerodynamic considerations; there are however, both advantages and disadvantages for each group, which are discussed in the following paragraphs. 2. COMPARISON OF MAIN ROTOR TYPES: The most common main rotor types in use today are the fully articulated and semi-rigid, but these definitions were produced in the earliest days of helicopter manufacture. In recent times it has been the practice to utilise features from one main rotor type on another, and the actual group that any modern main rotor head falls into can now be more difficult to define, especially since the introduction of elastomeric components, which are now very common. In general terms, it is easiest to define the three groupings by the constructional arrangement of the main rotor head, which are generally defined as follows: Rigid: A main rotor head that has no facility for the rotor blades to flap or drag. Semirigid: A main rotor head incorporating a flapping hinge, or other device that allows flapping. Fully articulated: A main rotor incorporating flapping hinges and drag or lead/lag hinges. Each of these main rotor head types will, of course, incorporate feathering bearings, to allow for rotor blade pitch changes, and therefore these are excluded from the definitions. Whilst these are workable definitions they are by no means always accurate. For example the Westland Lynx helicopter has a main rotor described as being semirigid, and yet it provides a limited facility for the rotor blades to flap and drag

58 In any particular case it will be the manufacturer who decides which of the three generic groups the actual main rotor head belongs to, and this information will be found in the Aircraft Maintenance Manual (AMM) Rigid rotor systems: In a rigid rotor system, the blades, hub, and mast are rigid with respect to each other. The rigid rotor system is mechanically the simplest rotor. There are no vertical or horizontal hinges so the blades cannot flap or drag, but they can be feathered. Rigid main rotor blades Operating loads from flapping and lead/lag forces must be absorbed by bending rather than through hinges. By flexing, the blades themselves compensate for the forces which previously required rugged hinges. The result is a rotor system that has less lag in the control

59 response, because the rotor has much less oscillation. The rigid rotor system also negates the danger of mast bumping inherent in semirigid rotors. Four-blade rigid main rotor. The rigid rotor system shown in next figure is mechanically simple, but structurally complex because operating loads must be absorbed in bending rather than through hinges. In this system, the blade roots are rigidly attached to the rotor hub. In this example, blades are comprised of glass fiber reinforced material. The hub is a single piece of forged rigid titanium

60 They cannot flap or lead/lag, but they can be feathered. As advancements in helicopter aerodynamics and materials continue to improve, rigid rotor systems may become more common because the system is fundamentally easier to design and offers the best properties of both semirigid and fully articulated systems. The rigid rotor system is very responsive and is usually not susceptible to mast bumping like the semirigid or articulated systems because the rotor hubs are mounted solid to the main rotor mast. This allows the rotor and fuselage to move together as one entity and eliminates much of the oscillation usually present in the other rotor systems. Four-blade rigid main rotor. The rigid rotor includes a reduction in the weight and drag of the rotor hub and a larger flapping arm, which significantly reduces control inputs. Without the complex hinges, the rotor system becomes much more reliable and easier to maintain than the other rotor configurations. A disadvantage of this system is the quality of ride in turbulent or gusty air

61 The rigid rotor includes a reduction in the weight and drag of the rotor hub and a larger flapping arm, which significantly reduces control inputs. Without the complex hinges, the rotor system becomes much more reliable and easier to maintain than the other rotor configurations. A disadvantage of this system is the quality of ride in turbulent or gusty air. Westland Lynx rigid rotor system

62 Main rotor blade attachment joint on rigid main rotor system Because there are no hinges to help absorb the larger loads, vibrations are felt in the cabin much more than with other rotor head designs. There are several variations of the basic three rotor head designs. The bearingless rotor system is closely related to the articulated rotor system, but has no bearings or hinges. This design relies on the structure of blades and hub to absorb stresses

63 The main difference between the rigid rotor system and the bearingless system is that the bearingless system has no feathering bearing; the material inside the cuff is twisted by the action of the pitch change arm. Nearly all bearingless rotor hubs are made of fiber-composite materials. Advantages: Simplicity of design reduces maintenance activities and the potential for faults to develop. Generally rigid main rotor systems offer a smaller cross sectional area to the airflow than similar sized rotor heads of the other two types, and therefore drag is less. Control response is very rapid and accurate. Disadvantages: A more complex control system is required to ensure aircraft stability, especially in forward flight, as there is no capacity for flapping to equality other than by rotor blade flexing. Rotor blades must be of a very strong design to withstand the forces in all modes of flight. Rigid rotor systems can be very susceptible to sudden wind gusts and side winds Semirigid rotor systems: A semirigid rotor system is usually composed of two blades that are rigidly mounted to the main rotor hub. The main rotor hub is free to tilt with respect to the main rotor shaft on what is known as a teetering hinge. The teetering hinge allows the main rotor hub to tilt, and the feathering hinge enables the pitch angle of the blades to change. A semirigid rotor system allows for two different movements, flapping and feathering. This system is normally comprised of two blades, which are rigidly attached to the rotor hub. Flapping: The hub is then attached to the rotor mast by a trunnion bearing or teetering hinge and is free to tilt with respect to the main rotor shaft. This allows the blades to see-saw or flap together. As one blade flaps down, the other flaps up. Since there is no vertical drag hinge, lead/lag forces are absorbed and mitigated by blade bending. Feathering: Feathering is accomplished by the feathering hinge, which changes the pitch angle of the blade. Since there is no vertical drag hinge, lead-lag forces are absorbed through blade bending. Mast bumping: Helicopters with semirigid rotors are vulnerable to a condition known as mast bumping which can cause the rotor of gravity (COG) is below where it is attached to the mast. This flap stops to shear the mast. The mechanical design of the semirigid rotor system dictates downward flapping of the blades must have some physical limit. Mast bumping is the result of excessive rotor flapping. Each rotor

64 system design has a maximum flapping angle. If flapping exceeds the design value, the static stop will contact the mast. It is the violent contact between the static stop and the mast during flight that causes mast damage or separation. This contact must be avoided at all costs. Bell 230 semirigid rotor

65 Mast bumping is directly related to how much the blade system flaps. In straight and level flight, blade flapping is minimal, perhaps 2 under usual flight conditions. Flapping angles increase slightly with high forward speeds, at low rotor RPM, at high-density altitudes, at high gross weights, and when encountering turbulence. Manoeuvring the aircraft in a sideslip or during low-speed flight at extreme COG positions can induce larger flapping angles. BELL 230 semirigid main rotor

66 Underslung rotor: The underslung rotor system mitigates the lead/lag forces by mounting the blades slightly lower than the usual plane of rotation, so the lead and lag forces are minimized. As the blades cone upward, the centre of pressures of the blades are almost in the same plane as the hub. Whatever stresses are remaining bend the blades for compliance. If the semirigid rotor system is an underslung rotor, the centre of gravity (COG) is below where it is attached to the mast. This underslung mounting is designed to align the blades centre of mass with a common flapping hinge so that both blades centres of mass vary equally in distance from the centre of rotation during flapping. The rotational speed of the system tends to change, but this is restrained by the inertia of the engine and flexibility of the drive system. Only a moderate amount of stiffening at the blade root is necessary to handle this restriction. Simply put, under slinging effectively eliminates geometric imbalance. Advantages: The elimination of individual flapping hinges and dragging hinges simplifies construction, and eliminates many faults, particularly those that increase vibration levels, associated with fully articulated rotor heads. Because of the simplicity of design, maintenance can be made more simple and less time consuming. The rotor blades are fixed to the main rotor hub and do not usually rely upon centrifugal force for rigidity. The reduction in components also reduces the weight and drag producing characteristics when compared to the larger and more complex fully articulated main rotor heads. Disadvantages: Sudden wind gusts can cause instability, because of an inability, or limited ability for individual rotor blades to flap. Because of the lack of flapping and dragging hinges, greater bending forces are applied to the rotor blade roots and attachments. The rotor blades must be designed with sufficient strength to withstand these greater forces, increasing their weight. Teetering type semi-rigid main rotor heads need to be underslung to minimize Coriolis effect when the rotor see saws Fully articulated rotor system: In a fully articulated rotor system, each rotor blade is attached to the rotor hub through a series of hinges, which allow the blade to move independently of the others. These rotor systems usually have three or more blades. The blades are allowed to flap, feather, and lead or lag independently of each other. Horizontal (flapping) hinge: The horizontal hinge, called the flapping hinge, allows the blade to move up and down. This movement is designed to compensate for dissymmetry of lift. The flapping hinge may be located at varying distances from the rotor hub, and there may be more than one hinge

67 Fully articulated rotor systems allow each blade to lead/lag (move back and forth in plane), flap (move up and down about an inboard mounted hinge) independent of the other blades, and feather (rotate about the pitch axis to change lift). The flapping hinge for the blade permits this motion and is balanced by the centrifugal force of the weight of the blade, which tries to keep it in the horizontal plane. The centrifugal force is nominally constant; however, the flapping force is affected by the severity of the maneuver (rate of climb, forward speed, aircraft gross weight). As the blade flaps, its COG changes. These changes the local moment of inertia of the blade with respect to the rotor system and it speeds up or slows down with respect to the rest of the blades and the whole rotor system. This is accommodated by the lead/lag or drag hinge. This is also known as the conservation of angular momentum. An in-plane damper typically moderates lead/lag motion. Fully articulated flapping hub Vertical (drag) hinge: The vertical hinge, also called the lead-lag or drag hinge, allows the blade to move back and forth. This movement is called lead-lag, dragging, or hunting. Dampers are usually used to prevent excess back and forth movement around the drag hinge. The purpose of the drag hinge and dampers is to compensate for the acceleration and deceleration caused by Coriolis Effect. Each blade can also be feathered, that is, rotated around its spanwise axis. Feathering the blade means changing the pitch angle of the blade. By changing the pitch angle of the blades you can control the thrust and direction of the main rotor disc

68 Fully articulated main rotor principle Each blade of a fully articulated rotor system can flap, drag, and feather independently of the other blades

69 Fully articulated rotor movements As the rotor spins, each blade responds to inputs from the control system to enable aircraft control. The centre of lift on the whole rotor system moves in response to these inputs to effect pitch, roll, and upward motion. The magnitude of this lift force is based on the collective input, which changes pitch on all blades in the same direction at the same time. The location of this lift force is based on the pitch and roll inputs from the pilot. Fully articulated rotor systems are found on helicopters with more than two main rotor blades

70 Bell 427 fully articulated rotor system

71 Note: Bell 427 fully articulated rotor system is often referred to as soft in plane ; each blade operates independently and leads, lags, and flaps in a controlled manner due to elastomeric construction. Lead/lag hinge allows the rotor blade to move back and forth in plane

72 Drag hinge allows the rotor blade to move back and forth in plane Elastomeric bearings: Newer rotor systems use elastomeric bearings, arrangements of rubber and steel that can permit motion in two axes. Besides solving some of the above mentioned kinematic issues, these bearings are usually in compression, can be readily inspected, and eliminate the maintenance associated with metallic bearings. Elastomeric bearings are naturally fail-safe and their wear is gradual and visible. The metal-to-metal contact of older bearings and the need for lubrication is eliminated in this design

73 2.4. Combination rotor system: EC-725 main rotor with elastomeric bearings Modern rotor systems may use the combined principles of the rotor systems mentioned above. Some rotor hubs incorporate a flexible hub, which allows for blade bending (flexing) without the need for bearings or hinges. These systems, called Flextures, are usually constructed from composite material. Elastomeric bearings may also be used in place of conventional roller bearings. Elastomeric bearings are bearings constructed from a rubber type material and have limited movement that is perfectly suited for helicopter applications. Flextures and elastomeric bearings require no lubrication and, therefore, require less maintenance. They also absorb vibration, which means less fatigue and longer service life for the helicopter components. Combination rotor systems, such as Bell s soft-in-plane, use composite material and elastomeric bearing to reduce complexity and maintenance. These systems increase the reliability

74 Bell 412 Arapaho (CH-146 Griffon) The terms soft-in-plane and stiff-in-plane are use to compare lead-lag frequency, in the plane of rotation, to the shaft rotational frequency. If it's soft-inplane the lead-lag frequency is normally less than the shaft rotational frequency (soft follower of the shaft rotational) In this example, the CH-146 Griffon has four-blade hingeless flex-beam soft-in-plane main rotor. Hub is composed of two titanium flex-beam yokes, four steel spindles with grip lugs for blade attachment, and elastomeric bearings and dampers. A mechanical pendulum dampers is mounted on leading and trailing edges to absorb the vibrations

75 Agusta Westland AW-139 AW-139 has a fully articulated main rotor with five composite blades, five elastomeric bearings and five hydraulic dampers Bearingless Main Rotors (BMR technology): During the past three decades, the helicopter industry has invested a very substantial amount of resources in the development of production hingeless and bearingless rotor systems. Bearingless rotor systems, such as the Eurocopter systems, have contact surfaces or load points made of elastomeric composite components that deform and twist to allow blade movement. The hingeless (bearingless) rotor system functions much as the articulated system does, but uses elastomeric bearings and composite flextures to allow flapping and lead lag movements of the blades in place of conventional hinges

76 Its advantages are improved control response with less lag and substantial improvements in vibration control. It does not have the risk of ground resonance associated with the articulated type unless the landing gear system needs servicing. The hingeless rotor system is also considerably a more expensive system. Eurocopter EC-135 main rotor (Spheriflex BMR technology) Most of these components are on condition life items versus metal components which must be changed at certain times due to metal fatigue. The composite components are designed so that even if a portion fails, the aircraft can make a safe landing

77 Hingeless rotored helicopters, such as the MBB BO-105, and the Westland Lynx have been in production for almost 25 years. However, successful bearingless rotored helicopters have gone into production only during the last decade. Typical examples are the MD-900 Explorer, the Comanche bearingless main rotor, the Eurocopter EC135. Eurocopter's starflex and spheriflex bearingless main rotor hubs ensure very fast response to pitch changes, while also offering excellent manoeuvrability and stability. Used across Eurocopter's helicopter product line, they have a fail-safe design through the application of composite materials, and are practically maintenance-free

78 VII. BLADE DAMPERS: FUNCTION AND CONSTRUCTION 1. ROTOR VIBRATIONS: With the many rotating parts found in helicopters, some vibration is inherent. You need to understand the cause and effect of helicopter vibrations because abnormal vibrations cause premature component wear and may even result in structural failure. With experience, you learn what vibrations are normal versus those that are abnormal and can then decide whether continued flight is safe or not. Helicopter vibrations are categorized into low, medium, or high frequency Low frequency vibrations: Low frequency vibrations ( cycles per minute) usually originate from the main rotor system. The vibration may be felt through the controls, the airframe, or a combination of both. Furthermore, the vibration may have a definite direction of push or thrust. It may be vertical, lateral, horizontal, or even a combination. Normally, the direction of the vibration can be determined by concentrating on the feel of the vibration, which may push you up and down, backwards and forwards, or from side to side. The direction of the vibration and whether it is felt in the controls or the airframe is an important means for the mechanic to troubleshoot the source. Some possible causes could be that the main rotor blades are out of track or balance, damaged blades, worn bearings, dampers out of adjustment, or worn parts Medium and high frequency vibrations: Medium frequency vibrations (1,000-2,000 cycles per minute) and high frequency vibrations (2,000 cycles per minute or higher) are normally associated with out of- balance components that rotate at a high RPM, such as the tail rotor, engine, cooling fans, and components of the drive train, including transmissions, drive shafts, bearings, pulleys, and belts. Most tail rotor vibrations can be felt through the tail rotor pedals as long as there are no hydraulic actuators, which usually dampen out the vibration. Any imbalance in the tail rotor system is very harmful, as it can cause cracks to develop and rivets to work loose. Piston engines usually produce a normal amount of high frequency vibration, which is aggravated by engine malfunctions such as spark plug fouling, incorrect magneto timing, carburettor icing and/or incorrect fuel/air mixture. Vibrations in turbine engines are often difficult to detect as these engines operate at a very high RPM. 2. HELICOPTER GROUND AND AIR RESONANCE: In whirling, the kinetic energy (KE) of the hub is constantly changing because the circular motion requires a constant change of velocity. It follows that the kinetic energy of the blades must also be changing constantly

79

80 2.1. Ground resonance phenomenon: The blade KE variation is due to motion of the blade centre of gravity (COG) plus that due to in-plane rotation of the moment of inertia of the blade about the null point. Essentially whirling is a continuous interplay of blade and hub energy and in the absence of friction at the hinges and any aerodynamic effect it could continue indefinitely. Figure is in stationary co-ordinates. Figure (a) shows an articulated rotor turning anticlockwise and whirling forwards whereas (b) shows the same rotor which is still turning anticlockwise but which is whirling backwards. A rotor turning anticlockwise with forward whirling. The reaction from the hull opposes the whirling. A rotor turning anticlockwise with backward whirling. The reaction from the hull which is the same as in (a) is now in the same direction as the whirling and amplifies it. This is the mechanism of ground resonance Air resonance phenomenon: Air resonance is a condition primarily relevant to helicopters having hingeless rotors. In articulated rotors the natural frequency of the blade about the lagging hinge is low and lag dampers are in any case present to prevent ground resonance. Furthermore the motion of the hull is decoupled to some extent by the articulation. In a hingeless rotor the blades lag by flexing and the stiffness is higher, leading to a higher natural frequency of lagging motion. The relatively stiff connection between the rotor and the hull means that large rotor moments can excite hull flexing. If such hull flexing is resonant then an unstable system could exist. A bending mode of the hull could result in a lateral motion at the rotor head that is similar to the rocking experienced in ground resonance Ground (and air) resonance solutions: In the absence of preventive measures, a helicopter on the ground with progressive backward rotor whirling is a mechanical oscillator. Given the huge amount of energy stored in the rotor, once started the whirling will increase in amplitude until something breaks. Hull rocking (ground) resonance can only occur if there is a reaction from the ground, hence the name of the phenomenon. This also explains why a machine can fly safely but disintegrates on landing, as has happened on a number of occasions. There are a number of solutions to ground resonance which will be explored. It will be seen from a consideration that damping any changes in the angle θ will be highly effective hence the use of dragging dampers in the traditional fully articulated rotor head. In many cases damping is provided in the undercarriage to dissipate landing impacts and this damping can augment but not replace the damping in the head

81 In modern helicopters employing damping, ground resonance is virtually unknown provided the dampers are kept in good order. These dampers may be hydraulic, similar to automobile dampers, which work by forcing oil through a small orifice, or elastomeric, which work by dissipating heat in hysteretic flexing. The latter have the advantage of needing no maintenance. Oil filled dampers will lose effectiveness if the oil leaks. Given the destructive nature of ground resonance, it is a good idea to examine the dampers as part of the pre-flight inspection. By moving the blade on its dragging hinge, the resistance of the damper can be felt and the oil can be heard rushing through the damping orifice. All of the blades should feel and sound the same. If one blade feels different the damper may have some air in it. As the air is forced through the damping orifice the sound will change. Whilst one weak damper may not cause ground resonance, it may result in an increase in vibration in forward flight. It is also useful to learn the characteristic of the machine s padding on start-up. If the rotor dampers are satisfactory, but there is unusual padding, the undercarriage oleos may need attention. A smoother rotor start may result if all of the blades are first moved to their rearward damper travel limit. Unusual padding may also result if the machine is parked on a slope when gravity will tend to take the blades out of pattern during the early stages of starting. 2. DAMPER FUNCTION: The main function of the damper is to slow the tendency of the blade to lag when the rotor is first started up or to lead when the rotor is slowed or the brake applied. Dampers are usually used to prevent excess back and forth movement around the drag hinge. The purpose of the drag hinge and dampers is to compensate for the acceleration and deceleration caused by Coriolis effect. They absorb the jolts and attenuate the blades vibrations during their revolution. There were several types, but some are not used any more, like the damper with disc, friction or hydraulics with pallets. On the other hand, the piston hydraulic dampers are still largely widespread. In certain cases, dampers are doubled of blade spacing wires, on old design machines. The damper help keep the blades equally spaced and reduce the possibility of ground resonance Hydraulic dampers: Hydraulic dampers rely upon throttled flow of fluids that generate damping forces proportional to the square of velocity. This can result in extremely high damping forces when amplitudes increase

82 Hydraulic dampers are prone to leakage problems and have a short life due to a large number of moving components and seal wear resulting in higher maintenance cost. The weight and size of these dampers result in another penalty due to high parasitic load and aerodynamic drag. Fully-articulated rotor and damper 2.2. Elastomeric dampers: Elastomeric dampers employing bonded elastomeric material exhibit viscoelastic behavior under dynamic conditions, dissipating energy through hysteresis. These materials respond non-linearly to the amplitude of motion, frequency of motion, and temperature. Compared with conventional

83 hydraulic dampers, elastomeric dampers are lighter in weight and have fewer parts. They do not have the sliding seals of hydraulic dampers and are not affected by sand and dust. Elastomeric dampers exhibit gradual deterioration, detectable by visual inspection. These elastomers are rapidly gaining popularity as a solution for designing lag dampers with high damping capability, and have been utilized for articulated rotor systems; examples include the Eurocopter EC-135, the Boeing AH-64 Apache, the Boeing, CH-47F Chinook, the Bell model 412, the Sikorsky RAH-66 Comanche, and the McDonnell Douglas Explorer. An elastomeric damper is nonlinear and highly dependent on frequency, temperature and loading conditions such as preload and excitation amplitudes. The damping of an elastomer has been shown to degrade substantially at low amplitudes, causing undesirable limit cycle oscillations. Consequently, alternative methods for augmenting aeromechanical stability are explored. Elastomeric inertial dampers

84 Inertial dampers (Eurocopter EC-135 main rotor) Advantages of elastomeric dampers: Elastomeric dampers are generally self-lubricating, Elastomeric dampers cannot seize, They are not subject to faults such as brinelling, pitting or galling that conventional bearing suffer from,

85 Less vibration and shock is transmitted to adjacent components, Boots and seals are not required as no lubrication is required, They contain fewer parts, and are generally bonded together, eliminating a potential loose article hazard, They have better resistance to environmental conditions, Generally elastomeric components have a life of at least 5 times that of conventional components, They can be designed to take all loads and motion at a set point, eliminating the need for several bearings to be used in one location. Disadvantages of elastomeric dampers: The cost of elastomeric components tend to be much greater than similar conventional components, The size is dependent upon the load to be taken, and elastomeric components can often be larger than the conventional component they replace, They are susceptible to attack by strong chemicals and solvents, and require careful handling Variable dampers: Current practice of adding passive damping may be improved to handle large dynamic range of the blade with several peaks of vibration resonance. To minimize extra-large damping forces that may damage the control system of blade, passive dampers should have relatively small damping coefficients, which in turn limit the effectiveness. By providing variable damping, a much larger damping coefficient to suppress the vibration can be realized. If the damping force reaches the maximum allowed threshold, the damper will be automatically switched into the mode with smaller damping coefficient to maintain near-constant damping force. Furthermore, the proposed control system will also have a fail-safe feature to guarantee the basic preformation of a typical passive damper. The proposed control strategy to avoid resonant regions in the frequency domain is to generate variable damping force in combination with the supporting stiffness to manipulate the restoring force and conservative energy of the controlled blade system. Two control algorithms are developed and verified by a prototype variable damper, a digital controller and corresponding algorithms. Primary experiments show good potentials for the proposed variable damper: about 66% and 82% reductions in displacement at 1/3 length and the root of the blade respectively

86 2.4. Fluid filled dampers: Main rotor with variable dampers Fluid filled dampers combine benefits of bonded elastomeric dampers such as simplicity, lightweight, reliability, energy storage capacity, multi-axis spring rate capability and maintenance free operation with the broader range of dynamic capability provided by non-toxic, non-corrosive fluids. These dampers are exclusively manufactured by the Lord Corporation under the trade name of fluidlastic. Fluidlastic dampers enjoy several advantages over hydraulic dampers in the sense that they do not require dynamic seals, extremely close tolerances, plated surfaces and polished finishes on the components

87 As they are hermetically sealed, they are not affected by sand and dust, are not prone to leakage, and are thus designed to be maintenance-free. Recent application of Fluidlastic includes the Bell 430, the NH90 and the RAH-66 Comanche. 3. DAMPER CONTRUCTION: 3.1. Hydraulic dampers: In the case of modern hydraulic shock absorbers, heating up the viscous fluid and compressing the nitrogen gas inside the shock body is what absorbs and dissipates the kinetic energy generated during suspension travel. Hydraulic damper (cross-section) The oil in the shock is heated during a compression event, where the shock piston and attached valve shims are pushed down into the shock body, forcing the fluid through orifices in the valves. This generates heat within the fluid, which is then dissipated to the shock body, where it can be transferred to the ambient atmosphere. The same heat generation, absorption and dissipation process happens in reverse on the rebound stroke, where the suspension travels back to its neutral position

88 3.2. Elastomeric inertial dampers: Internal part of a hydraulic damper As a result of blade lag motion, the damper mass oscillates in the lag direction and the fluid in the tuning port is pumped through the inner chamber. Fluid motion creates a force which reduces the effective stiffness of the damper. The fluid force increases as the frequency of the system increases

89 3.3. Embedded inertial dampers: Embedded inertial dampers may be promising for lag damping of rotor blades. In addition, embedded inertial dampers may utilize part of the leading edge weight of the blade and simplify the rotor hub considerably. Embedded mechanical devices have been successfully integrated into full scale rotor blades. An embedded inertial damper will be subject to similar loads and geometric constraints as existing embedded devices. An ideal embedded chordwise inertial damper for helicopter blade lag damping would have both a high static stiffness and a low dynamic stiffness

90 Embedded inertial damper

91 VIII. ROTOR BLADES: MAIN AND TAIL ROTOR BLADE CONSTRUCTION AND ATTACHMENT 1. GENERAL: A rotor is a whole made up of a hub (or rotor head) on which are fixed the blades. The rotor diameter choice depends on the use which will be made. Amongst other things, it is a function of the blades number, the horse-power available and speed of traverse. A rotor of low diameter decreases the weight and complexity, and increases the traverse speed. On the other hand, it requires a higher power available for the hovering. The reverse occurs for a rotor of large diameter which, moreover, offers a larger kinetic energy in autorotation. The number of blades depends primarily on the wing load (report/ratio of the mass to be raised on the rotor disc surface). A light helicopter will be able to have two or three blades, whereas a large device has five or six of them. 2. BLADE DESIGNS: The adopted designs of blades can be: rectangular, trapezoidal, mixed. The rotor blade is an airfoil designed to rotate about a common axis to produce lift and provide directional control for a helicopter. It is often referred to as a rotary wing. The design and construction of a rotor blade vary with the manufacturer, although they all strive to manufacture the most efficient and economical lifting device. The particular helicopter design places certain requirements on the main rotor blades, which influence their design and construction. Most rotor blades are designed as symmetrical airfoils to produce a stable aerodynamic pitching characteristic. Aerodynamic stability is achieved when the centre of gravity, centre of pressure, and blade-feathering axis all act at the same point. The blade is more stable in flight because these forces continue to act at almost the same point as the blade changes pitch. At present only one Army helicopter is equipped with an unsymmetrical airfoil. This unsymmetrical airfoil blade is capable of producing greater lift than a symmetrical airfoil blade of similar dimensions. Aerodynamic stability is achieved by building a 3 upward angle into the trailing edge section of the blade. This prevents excessive centre-of pressure travel when the rotor blade angle of attack is changed

92 Typical rotor blades

93 According to the figure above, we notice that the blade root end is notched because, in translation, it appears at this place a zone of unhooking and another one of reversed flow (comparable zones with those appearing in autorotation with translation). A blade is a shaped surface intended to turn in an appreciably horizontal plane. Its characteristics are comparable with those of a plane wing, put aside the lengthening which is larger. The lengthening of a blade is the ratio between its depth e on its length R Max. e A = 100% R Max

94 3. BLADE STRUCTURES AND MATERIALS: 3.1. Forces acting on blades: The matters choice and the construction methods depend on the techniques of the moment but are especially based on the efforts which the blades must undergo. Let us see in a diagrammatic way the resultant of the blade elementary efforts: Efforts due to gravity (downwards inflection), Traction efforts counterbalanced by the efforts of opposed direction trail, Centrifugal loads (wrenching), Lift efforts (inflection upwards), Inertial efforts due to the step variation, Aerodynamic efforts of damping due to the vertical beats (alternating bending),

95 Beats in the rotation plan due to the trail variations and the forces of Coriolis (alternating bending) Types of blades: Materials: A variety of material is used in the construction of rotor blades; aluminum, steel, brass, and fiberglass are most common. The first helicopter rotor blades were constructed out of laminated wood and fabric. One of the major drawbacks of using wood to construct the rotor blades is that wood absorbs moisture, which changes the mass of the rotor blade. Wooden rotor blades were used up until the 1960s, until they were replaced by steel and aluminium. Advantages of steel and aluminium rotor blades are that they're cheaper and easier to produce, and that they do not suffer from moisture absorption. However, disadvantages include a low strength to density ratio and a poor resistance to fatigue. Today, composite materials are principally used for rotor blades. Blades construction: Blades are either full, or dig (blades out of box). In the blades construction, it is necessary to avoid the processes resisting tiredness badly (notch effects, corrosion concentration, stress concentration). To this end, the assembly by bolts or rivets is avoided as much as possible; bonding is often preferred. A particularly delicate area of the blade is the fitting of fastener. The assembly must carefully be studied for resumption as progressive as possible of the efforts (fittings in bevel, precise borings of the bolts holes, plastic film interposition between metal surfaces likely to present contact corrosion). The bending stresses, alternate and permanent, are very high to approximately 20 % of the ray. If the blade inking is done in the vicinity of this section, the precautions must be reinforced. A typical metal blade has a hollow, extruded aluminum spar which forms the leading edge of the blade. Aluminum pockets bonded to the trailing edge of the spar assembly provide streamlining. An aluminum tip cap is fastened with screws to the spar and tip pocket. A steel cuff bolted to the root end of the spar provides a means of attaching the blade to the rotor head. A stainless steel abrasion strip is adhesive-bonded to the leading edge

96 3.3. Wood blades: Wood the most used was: balsa, the spruce, the mahogany tree and the birch. The fibrous texture of wood confers on the blades a great flexibility which makes them less prone to the phenomena of tiredness than the metal blades (fittings of fastener put aside); moreover, the bonded assembly is extremely easy. On the other hand, the variations in temperature and relative humidity cause deformations and frequent unbalancing, fabrication, were moreover artisanal. These blades are not used any more Metal blades: The materials using the composition of these blades are mainly the alloys of aluminum and steels, often supplemented by the plastics. The member is the centre piece of the blade; it generally forms the leading edge and is often obtained by spinning (or another process of the same kind) so as to have a fibrous texture authorizing the alternate constraints. The trailing edge either is brought back to the spar, or integrated in the blade. It is sometimes reinforced by veins forming a box; it is often supplemented of filling light material (metal or plastic). The blade skins are formed around and bonded to the spars, which in most cases form the leading edge of the blades. Metal blade skins are supported from the inside with aluminum honeycomb, ribs, and some smaller blades which have no bracing or support inside themselves. Metal tail rotor blades are usually of two constructional methods, the first uses a D section extruded spar and ribs to form the blade shape, and the second a solid spar and honeycomb filling. The illustration shows a tail rotor blade that uses the first of the two common constructional methods. It has an open, D shaped extruded spar that runs from the blade tip, to the root, where it forms the attachment fitting. A small light alloy plate is bonded to the inside of the spar lip to complete the D shape and provide torsional strength and rigidity. Thin light alloy ribs are bonded to the rear face of the spar, providing shape and resisting compressive loads felt on the skin. The light alloy skin is then bonded onto the whole structure, the use of bonding adhesives ensures an uninterrupted smooth skin that enhances aerodynamic efficiency and reduces drag. The leading edge is covered with a much tougher metal strip, such as stainless steel, nickel or tungsten, to provide an anti-erosion sheathing. Like main rotor blades the tail rotor blade must be balanced, in this example chord wise balance weights are fitted at the canted closing rib, close to the attachment fitting and span wise weights are fitted at the tip. These are then covered by a light alloy tip cap, which is normally riveted into place using a blind riveting technique

97 Metal blade principle

98 3.5. Composite blades: It is a recent principle of construction; in this case, the blades are entirely out of composite materials enabling them to have a very great flexibility. The spar, essential element, is obtained starting from impregnated glass fibers of resin which go from the root at the end of the blade. The coating is obtained same manner; the fibers are generally cross. The trailing edge is filled with expanded foam (or another fill material). The blade skins are formed around and bonded to shaped titanium spars. The blade skins are supported inside with aluminum nomex honeycomb. The space around the spar is filled with foam plastic. Generally, composite rotor blade uses a hybrid composite material, made from glass and carbon fibres, these hybrid materials are designed to use the advantageous properties of each material to provide a stronger, lighter, or more load resistant material, than would be the case if only one fibre type was used. Typical composite blade The leading edge is protected by a stainless steel (titanium) anti-erosion sheathing, which, like the main rotor blades, incorporates a heating element to provide anti-icing during cold weather operation. The blade shape is provided by honeycomb, which is bonded to the unidirectional composite spar, all of which is covered by a cross ply composite skin. In this particular design the pitch change horn is set into bushes fitted to the tail rotor blade, unlike many other tail rotor blades, which are secured to a feathering bearing assembly that receives the control inputs. Balance weights are fitted to the angles closing strip close to the blade root for chord wise balance, and at the tip for span wise balance, and once again, a detachable fairing is used to cover these weights

99 3.6. NACA design airfoils: The NACA airfoils are airfoil shapes for aircraft wings developed by the National Advisory Committee for Aeronautics (NACA). The shape of the NACA airfoils is described using a series of digits following the word "NACA". The parameters in the numerical code can be entered into equations to precisely generate the cross-section of the airfoil and calculate its properties. The NACA four-digit wing sections define the profile by: First digit describing maximum camber as percents of the chord. Second digit describing the distance of maximum camber from the airfoil leading edge in tens of percents of the chord. Last two digits describing maximum thickness of the airfoil as percents of the chord. Four-digit series airfoils by default have maximum thickness at 30% of the chord (0.3 chords) from the leading edge. Example 1: NACA 0015 airfoil: The NACA 0015 airfoil is symmetrical, the 00 indicating that it has no camber. The 15 indicates that the airfoil has a 15% thickness to chord length ratio: it is 15% as thick as it is long

100 Example 1: NACA 2412 airfoil: Symmetrical NACA 0015 airfoil The NACA 2412 airfoil is non-symmetrical; it has a maximum camber of 2% located 40% (0.4 chords) from the leading edge with a maximum thickness of 12% of the chord. Non-symmetrical NACA 2412 airfoil 3.7. ONERA design airfoils: The ONERA airfoils are airfoil shapes for aircraft wings developed by the French Office National d Etudes et Recherche Aéronautique. The shape of the ONERA airfoils is described using a series of digits following the letters "OA"

101 Example 1: OA212 airfoil: It is an ONERA/Aerospatiale rotorcraft airfoil (Constructed from patent and smoothed) Max thickness 12% at 31.8% chord. Max camber 2.3% at 31.8% chord. First digit describing the type of airfoil: ex: OA2 = convex plan. Last two digits describing maximum thickness of the airfoil as percent of the chord. The OA212 airfoil is non-symmetrical; the first 2 digit indicates that it has 2% camber. The 12 last digits indicate that the airfoil has a 12% thickness to chord length ratio: it is 12% as thick as it is long. Non-symmetrical 0A212 airfoil Example 2: Specifications for AS Dauphin 365N airfoil: Blade weight: 42,3 kg. Length: 5275 mm. Twist: See curve below. Chord: 385 mm in the current part and 405 mm towards the extremity. Airfoil «convex plan» OA2: Constant (OA212) of reference 1000 to reference 3690 Evolutionary (OA212 OA209 OA207) of reference 3690 until the extremity

102

103 AS Dauphin 365 N3 airfoil details

104 4. ROTOR TRANSMISSION: The transmission system transfers power from the engine to the main rotor, tail rotor, and other accessories during normal flight conditions. The main components of the transmission system are the main rotor transmission, tail rotor drive system, clutch, and freewheeling unit. The freewheeling unit allows the main rotor transmission to drive the tail rotor drive shaft during autorotation. Because it is part of the transmission system, the transmission lubricates it to ensure free rotation. Helicopter transmissions are normally lubricated and cooled with their own oil supply. A sight gauge is provided to check the oil level. Some transmissions have chip detectors located in the sump. These detectors are wired to warning lights located on the pilot s instrument panel that illuminate in the event of an internal problem. Some chip detectors on modern helicopters have a burn off capability and attempt to correct the situation without pilot action. If the problem cannot be corrected on its own, the pilot must refer to the emergency procedures for that particular helicopter Main purposes: The primary purpose of the main rotor transmission is to reduce engine output RPM to optimum rotor RPM. This reduction is different for the various helicopters. As an example, suppose the engine RPM of a specific helicopter is 2,700. A rotor speed of 450 RPM would require a 6:1 reduction. A 9:1 reduction would mean the rotor would turn at 300 RPM. In helicopters with horizontally mounted engines, another purpose of the main rotor transmission is to change the axis of rotation from the horizontal axis of the engine to the vertical axis of the rotor shaft. This is a major difference in the design of the aircraft power plant and power train whereas the aircraft propeller is mounted directly to the crankshaft or to shaft that is geared to the crankshaft. The importance of main rotor RPM translates directly to lift. RPM within normal limits produces adequate lift for normal manoeuvring. Therefore, it is imperative not only to know the location of the tachometers, but also to understand the information they provide. If rotor RPM is allowed to go below normal limits, the outcome could be catastrophic

105 5.2. Clutch: Rotor transmissions In a conventional aircraft, the engine and propeller are permanently connected. However, in a helicopter there is a different relationship between the engine and the rotor

106 Because of the greater weight of a rotor in relation to the power of the engine, as compared to the weight of a propeller and the power in an aircraft, the rotor must be disconnected from the engine when the starter is engaged. Idler or manual clutch

107 A clutch allows the engine to be started and then gradually pick up the load of the rotor. Freewheeling turbine engines do not require a separate clutch since the air coupling between the gas producer turbine and the power (takeoff) turbine functions as an air clutch for starting purposes. When the engine is started, there is little resistance from the power turbine. This enables the gas producer turbine to accelerate to normal idle speed without the load of the transmission and rotor system dragging it down. As the gas pressure increases through the power turbine, the rotor blades begin to turn, slowly at first and then gradually accelerate to normal operating RPM. On reciprocating and single-shaft turbine engines, a clutch is required to enable engine start. Air, or wind milling starts, is not possible. The two main types of clutches are the centrifugal clutch and the idler or manual clutch. How the clutch engages the main rotor system during engine start differs between helicopter design. Piston powered helicopters have a means of engaging the clutch manually just as a manual clutch in an automobile. This may be by means of an electric motor that positions a pulley when the engine is at the proper operating condition (oil temperature and pressure in the appropriate range), but it is controlled by a cockpit mounted switch Belt drive clutch: Some helicopters utilize a belt drive to transmit power from the engine to the transmission. A belt drive consists of a lower pulley attached to the engine, an upper pulley attached to the transmission input shaft, a belt or a set of V-belts, and some means of applying tension to the belts. The belts fit loosely over the upper and lower pulley when there is no tension on the belts. Some aircraft utilize a clutch for starting. This allows the engine to be started without requiring power to turn the transmission. One advantage this concept has is that without a load on the engine starting may be accomplished with minimal throttle application. However, caution should also be used during starting, since rapid or large throttle inputs may cause overspeeds. Once the engine is running, tension on the belts is gradually increased. When the rotor and engine tachometer needles are superimposed, the rotor and the engine are synchronized, and the clutch is then fully engaged. Advantages of this system include vibration isolation, simple maintenance, and the ability to start and warm up the engine without engaging the rotor. When the clutch is not engaged, engines are very easy to overspeed, resulting in costly inspections and maintenance Centrifugal clutch: The centrifugal clutch is made up of an inner assembly and an outer drum. The inner assembly, which is connected to the engine driveshaft, consists of shoes lined with material similar to automotive brake linings. At low engine speeds, springs hold the shoes in, so there is no contact with the outer drum, which is attached to the transmission input shaft. As engine speed increases, centrifugal force causes the clutch shoes to move outward and begin sliding against the outer drum. The transmission input shaft begins to rotate, causing the rotor to turn slowly at first, but increasing as the friction increases between the clutch shoes and transmission drum

108 5. MAIN ROTOR HEADS: The number of approaches used in rotor head design over the years has been subject to wide variation. In the earliest days the theory of rotor dynamics was not well established and most designs mixed theory with empiricism. As understanding increased, attention turned first to optimization of design for a particular purpose and subsequently to such refinements as the reduction of vibration and of build and maintenance cost Hooke joint teetering head: The fully gimballed teetering head which essentially contains a Hooke joint so that the rotor can tilt in any direction with respect to the mast. The blade grips are mounted on feathering hinges. The feathering axis is parallel to one of the gimbal axes and the attitude of the hub in that axis would be indeterminate were it not for an equalizing linkage that makes the pitch of each blade equal with respect to the hub. This type of head was used on early Bell machines

109 5.2. Bell 206 teetering head: This type of teetering head where only a flapping hinge is fitted and tilting in the other axis is accommodated by the feathering hinges. This type of head can be seen on the Bell 206 Jet Ranger. It is not obvious how this arrangement can act as a universal joint to allow the disc to tilt in any direction relative to the mast. The answer is that the feathering bearings are used to provide the additional degree of freedom Tri-hinge head: A variant of the teetering rotor is one in which the blades have individual flapping hinges so that no bending moments occur due to coning. This allows the blades to be more lightly built than in a conventional teetering system. The effect of the offset between the flapping hinges is eliminated by the action of a third underslung bearing. This approach is known as a tri-hinge head and is used in the Robinson R

110 5.4. Door-hinge hub: This hub is relatively thin in section to reduce drag. The thin section also introduces some flexibility. This type of head was used on the Bell AH-1G

111 5.5. Semi-teetering head with elastomeric element: The semi-teetering head of the Bell 222 contains elastomeric feathering bearings in compression, a thin flexural section and a centring spring on the teeter bearing to give stability in low-g manoeuvres. 6. MAIN ROTOR CONSTRUCTION AND ATTACHMENT:

112 6.1. Rotor head of the BO-105: The only real bearings are for blade feathering. Flapping and dragging movements are accommodated through flexing of the blade shank. The dragging stiffness is supercritical and no dampers are needed. This is also a relatively stiff head in flapping, and pendulum vibration dampers are needed. The BO105 was the first light twin series in the world but also the first to be equipped with a rigid main rotor flapping hinge and drag and composite rotor blades. Trademarked "Rotor System Bölkow", this innovative four-blade system gave the B0-105 superior manoeuvrability (first helicopter able to do loops). Flapping and dragging are permitted by the flexibility of the blade shank

113 BO-105 rotor head

114 6.2. Rotor head of Lockheed AH-56 Cheyenne: Following figure shows another stiff-in-plane head, that of the Lockheed AH-56 Cheyenne. This has door-hinge feathering bearings for low drag and a virtual flapping hinge. The flapping hinge is very stiff indeed and the following rate of the rotor is very high, requiring full-time gyro stabilization. Lockheed Cheyenne head has virtual flapping hinge but is very stiff in dragging

115 6.3. Rotor head of Westland Lynx: Figure shows a soft-in-plane hingeless rotor from the Westland Lynx. Here the flapping flexure is a massive piece of titanium that is only flexible in the context of the enormous forces set up in a rotor head. The flexures are relatively stiff in flapping. Since the rotor can exert large moments on the hull, the Lynx does not need a tall mast; in fact it was a design goal to keep the height down to allow the machine to fit into transport aircraft, and a special low profile gearbox was designed to go with the rotor. With such a low rotor, the Lynx needs the tail boom to be angled down to give blade clearance before turning up to mount the tail rotor. Westland Lynx head has flexible members to allow flapping and dragging. Virtual flapping hinge is stiff. Dragging hinge is soft and drag dampers are needed

116 Westland Lynx AH Rotor head of Bell-412: Figure shows the head of the Bell 412. This has outboard dragging hinges and the drag damping is obtained by an elastomeric block acting on a inward extension of the blade grip which also mounts the pitch arms. The dragging hinges are mounted on flex beams that can twist to act as feathering bearings and bend to allow flapping.two such flex beams are stacked to make a four-blade rotor

117 Bell 412 head having elastomeric drag dampers. Flexural members bend to allow flapping and twist to allow pitch change

118 Bell 412 main rotor assembly

119 6.5. Rotor head of EC-725: Figure shows a head shows a head developed by Aerospatiale witch uses a spherical laminated elastomeric bearing which takes the axial blade thrust whilst allowing feathering, flapping and dragging. Elastomeric blocks are fitted inside the blade grip and these couple with extension arms on the hub. Spheriflex head uses spherical electrometric bearing for flapping, dragging and pitch change

120 The arms are stiff in the dragging plane, but the elastomeric block is in shear for dragging movements, giving a soft-in-plane characteristic and providing the drag damping. In flapping the blocks are in compression where they are quite stiff. Flapping flexibility is provided through a virtual hinge in the arms. The head arms also twist to allow feathering. Eurocopter EC

121 Spheriflex is the simplest technology for rotor blade hubs: Eurocopter Fennec AS Spheriflex technology Extensive use of composite material. SPHERIFLEX rotor head: no lubrication required, small amount of parts. Interchangeable composite main rotor blades, infinite life, no corrosion

122 6.6. Sikorsky S-64 Skycrane example: Description: The main rotor head is fully articulated. The main rotor head is made up of two principal components the hub and the swashplate. The hub consists of an upper and lower plate, hinge assemblies, sleeve and spindle assemblies, and six pressurized dampers. The swashplate consists of a rotating disc and stationary disc connected together

123 Other components of the head are anti-flapping restrainers, adjustable pitch control rod assemblies, and rotating and stationary scissors. Sikorsky S-64 Skycrane description The swashplate and control rods permit movement of flight controls to be transmitted to the blades. The hinge assemblies allow the blades to lead, lag, and flap. The sleeve and spindle assemblies allow each blade to be turned on its span wise axis to change blade pitch

124 The dampers prevent damage to the head by absorbing blade loads during rotor operation. Sikorsky S-64 Skycrane scissors The anti-flapping restrainers restrict flapping motion of the blades when the main rotor head is slowing or stopped. The rotating scissors are connected to the rotating swashplate and to the lower plate of the hub. The stationary scissors are connected to the stationary swashplate and the main gearbox upper housing

125 The anti-flapping restrainers restrict flapping motion of the blades when the main rotor head is slowing or stopped. The rotating scissors are connected to the rotating swashplate and to the lower plate of the hub. The stationary scissors are connected to the stationary swashplate and the main gearbox upper housing. The sleeve and spindle assembly receives lubrication from the individual oil reservoir mounted on the upper plate above the horizontal hinge pin. The oil lubricates the spindle bearings through an external tube from the horizontal hinge pin housing. Centrifugal force and gravity feed the oil into the spindle. Swashplate: The swashplate has two halves: stationary and rotating. Stationary and rotating scissors help to keep their perspective swashplate stable

126 The swashplate assembly pivots around the drive shaft on the uniball. Swashplate description Each MRB has a damper which allows it to lead and lag. The stationary swashplate attaches to the primary servos and transfers inputs to the rotating swashplate. The rotating swashplate rotates with the rotor head and transfers inputs from the stationary swashplate to the pitch horn

127 Control rods: The adjustable control rods extend from the sleeve and spindle horn assembly to the rotating Swashplate. The bearings are allowed inch (0.33 mm) play or a combined allowance of inch (0.533 mm). Blade pitch adjustments are made at the PC link, each notch on the PC moves the blade path about 0.5 inch (1.27 cm). Typical control rod Rotating and stationary scissors: The rotating scissor assemblies are bolted between the lower plate of the main rotor hub and rotating swashplate. They cause the rotating swashplate to rotate with the main rotor hub. The stationary scissors connect the stationary swashplate to the main gearbox. They prevent the stationary swashplate from turning with the main rotor head

128 Droop and flap restrainers: The droop and flap restrainers are spring loaded in the locked position and are disengaged with centrifugal force. They restrict drooping and flapping of blades when the main rotor head is slowing down or static

129 Hydraulic damper installation: The dampers are mounted between the upper and lower plates of the main rotor head to slow hunting of blades during rotation and to absorb rotor engagement loads. The dampers are pressurized with hydraulic fluid from an accumulator installed in the main rotor shaft. Constant pressurization allows the damper piston shaft to be fully extended at all times except when main rotor loads cause the fluid to be bypassed internally, causing the shaft of the piston to withdraw into the damper

130 Typical damper installation

131 Main rotor head installation: Prior to installing the main rotor head inspect the vertical shaft splines and split cone seat. Inspect threaded relief grove for cracks with 5x magnifying glass. If baked resin coating is present no inspection is required. The main gear box shaft has a 3º tilt. When removing or installing the main rotor head this tilt should be removed by filling the nose strut and left hand gear leg with nitrogen. The main rotor head is held onto the shaft by a set of split cones, a retaining nut and a pressure plate

132 Torque procedure: The pressure plate bolts and retaining nut bolts are numbered on both the top and bottom of the main rotor head. Torque retaining nut bolt assemblies to 1045 inch-pounds (87 foot-pounds, 118 Nm.) in increments of 200 inch-pounds (26 Nm.); in the order of etched numbers on retaining nut, until they stabilize. Torque the pressure plate bolts to 2400 inch-lbs (200 foot-lbs or 271 Nm.) in 1000 inch-lbs (113 Nm.), increments with the final torque a 3389 inch-lbs (281 foot-lbs or 383 Nm.) in the order of numbers on pressure plate, until they stabilize. Recheck retaining nut bolt torque, if any bolt slips torque all bolts in sequence to 1045 inch-pounds (118 Nm.)

133 After installing main rotor blades check torque on the retaining nut bolts again. If they move torque top and bottom again as per maintenance manual

134 After 10 hours of flight a torque check of both retaining nut and pressure plate is due and is repeated at 10 hour intervals until the head torque stabilizes. After 3 torque checks the head must be removed and reinstalled after inspecting condition of cones, cone seats, pressure plate and retaining nut. When removing main rotor head the torque should be broken in the same increments as the installation

135 6. TAIL ROTOR CONSTRUCTION AND ATTACHEMENT: Helicopters with a single, main rotor system require a separate antitorque system. This is most often accomplished through a variable pitch, antitorque rotor or tail rotor. Pilots vary the thrust of the antitorque system to maintain directional control whenever the main rotor torques changes, or to make heading changes while hovering. Most helicopters drive the tail rotor shaft from the transmission to ensure tail rotor rotation (and hence control) in the event that the engine quits. Usually, negative antitorque thrust is needed in autorotations to overcome transmission friction. The tail rotor driveshaft is connected to both the main transmission and the tail rotor transmission

136 6.1. Antitorque systems: The antitorque (tail rotor) drive system consists of an antitorque drive shaft and a antitorque transmission mounted at the end of the tail boom. The drive shaft may consist of one long shaft or a series of shorter shafts connected at both ends with flexible couplings. This allows the drive shaft to flex with the tail boom. The tail rotor transmission provides a right angle drive for the tail rotor and may also include gearing to adjust the output to optimum tail rotor RPM. Tail rotors may also have an intermediate gearbox to turn the power up a pylon or vertical fin. Tail rotors operate collectively only therefore, only one control input is required that changes the pitch on all tail rotor blades equally and simultaneously, in other words collectively. Tail rotor heads for aircraft normally consist of a tail rotor hub, tail rotor blade sleeve and spindle assembly, and a control mechanism. It is usual to find that the control mechanism is incorporated into the tail rotor gearbox, with the actuator being bolted to the rear of the gearbox, with its ram running through the tail rotor gearbox output shaft transferring control inputs to the tail rotor by means of a spider, or pitch change beam mechanism. Helicopters with a single main rotor must have some means of balancing the torque reaction due to driving the rotor. Whilst the anti-torque rotor could be mounted anywhere for hovering, in forward flight the most stable location is aft, supported on some kind of structure called a boom. In addition to the antitorque function helicopters need some means of yaw control and the tail rotor also serves that purpose. In order to balance the weight of machinery at the tail, the helicopter cabin usually extends some way forward of the mast. This large forward side area is unstable in yaw and generally some fin area is needed to give directional stability in forward flight. A further consideration is that the main rotor on its own is unstable in pitch in forward flight and a tail plane is usually required. As a result the tail of the conventional helicopter will be a structure supporting a variable pitch tail rotor, its transmission and controls, some fin area and a tail plane. In reality tail booms are also encrusted with antennae, flare launchers, navigation lights, strobe beacons, static vents, registration letters and warning notices, not to mention tail skids and occasionally part of the undercarriage. In a helicopter with more than one rotor the torque may be cancelled in a different way and a great variety of yaw control mechanisms will be found. These multi-rotor yaw mechanisms are considered in Chapter 9. In this chapter the conventional type of tail rotor will be considered, along with a number of alternatives having various advantages and drawbacks. The conventional tail rotor is well understood and relatively inexpensive owing to its wide use. The far aft location of the tail rotor assembly means that it must be lightly built in order to avoid the machine becoming tail heavy. The thrust needed from the tail rotor is much smaller than that needed in the main rotor. These two features conspire to ensure that the tail rotor is considerably more fragile than the main rotor. Unfortunately the fragile tail rotor is in a more exposed location where the pilot cannot see it. During certain manoeuvres, such as quick stops and rearward flight it can get very close to the ground, and it may come off worse in any encounter with vegetation the pilot has not seen. Tail rotor blades are usually fitted with soft aluminium telltales at the tips which will be distorted by any impact and indicate that a close inspection for damage is required

137 6.2. Antitorque fenestron systems: Another form of antitorque system is the fenestron or fan in- tail design. This system uses a series of rotating blades shrouded within a vertical tail. Because the blades are located within a circular duct, they are less likely to come into contact with people or objects. Typical tail rotor fenestron system

138 6.3. Robinson R22 antitorque example:

139 IX. TRIM CONTROL, FIXED AND ADJUSTABLE STABILIZERS 1. TRIM CONTROL: 1.1. Introduction: The general principle of flight with any aircraft is that the aerodynamic, inertial and gravitational forces and moments about three mutually perpendicular axes are in balance at all times. In helicopter steady flight (non-rotating), the balance of forces determines the orientation of the main rotor in space. The balance of moments about the aircraft centre of gravity (CG) determines the attitude adopted by the airframe and when this balance is achieved, the helicopter is said to be trimmed. To a pilot the trim may be hands on or hands off ; in the latter case in addition to zero net forces and moments on the helicopter the control forces are also zero.. In steady cruise the function of a tailplane is to provide a pitching moment to offset that produced by the fuselage and thereby reduce the net balancing moment which has to be generated by the rotor

140 The smaller this balancing moment can be, the less is the potential fatigue damage on the rotor. In transient conditions the tailplane pitching moment is stabilizing, as on a fixed-wing aircraft, and offsets the inherent static instability of the fuselage and to some extent that of the main rotor. A fixed tailplane setting is often used, although this is only optimum for one combination of flight condition and CG location. A central vertical fin is multi-functional: it generates a stabilizing yawing moment and also provides a structural mounting for the tail rotor. The central fin operates in a poor aerodynamic environment, as a consequence of turbulent wakes from the main and tail rotors and blanking by the fuselage, but fin effectiveness can be improved by providing additional fin area near the tips of the horizontal tailplane. When the flapping hinge axis is offset from the shaft axis (the normal condition for a rotor with three or more blades), the centrifugal force on a blade produces a pitching or rolling moment proportional to disc tilt. Known as direct rotor moment, the effect is large because although the moment arm is small the centrifugal force is large compared with the aerodynamic and inertial forces. A hingeless rotor produces a direct moment perhaps four times that of an articulated rotor for the same disc tilt. Analytically this would be expressed by according to the flexible element an effective offset four times the typical 3% to 4% span offset of the articulated hinge. Looking now at a number of trim situations, in hover with zero wind speed the rotor thrust is vertical in the longitudinal plane, with magnitude equal to the helicopter weight corrected for fuselage downwash. For accelerating away from hover the rotor disc must be inclined forward and the thrust magnitude adjusted so that it is equal to, and directly opposed to, the vector sum of the weight and the inertial force due to acceleration. In steady forward flight the disc is inclined forward and the thrust magnitude is adjusted so that it is equal to, and directly opposed to, the vector sum of the weight and aerodynamic drag. The pitch attitude adopted by the airframe in a given flight condition depends upon a balance of pitching moments about the CG. Illustrating firstly without direct rotor moment or tailplane-and airframe moment, the vector sum of aircraft drag (acting through the CG) and weight must lie in the same straight line as the rotor force. This direction being fixed in space, the attitude of the fuselage depends entirely upon the CG position. Like any aircraft, helicopters require a trim system that allows the pilot(s) to correct an undesirable attitude without the need to constantly hold the cockpit control to correct this condition. Helicopter trim systems are incorporated into the normal flight control system, and any input will move the control neutral datum from the cockpit control to the main rotor blades. This is normally accomplished by one of the following methods: Magnetic brakes and Electrical trim actuators. Some aircraft will use one or other method, whilst others will use a combination of the two, but in each case the feedback loop is via the pilot(s), where they will stop the trim input once the helicopter is in the desired flying attitude. Whichever of these systems is used, it is necessary to introduce the trim input, without affecting the feel force, provided by the gradient units, or in the case of some tail rotor control systems, yaw pedal dampers. If the trim system applied movement onto these feel producing devices, the amount of feel force would be different in each direction of movement

141 For example, if movement of the trim control produced a forward movement of the cyclic pitch stick, and compressed the cyclic gradient unit, any movement from this new neutral position would provide greater feel force for any additional forward movement, and less feel force for any rearwards movement than would normally be the case. This undesirable condition is prevented by arranging the trim system so that any trim input will not only move the control system to a new datum, but also the feel system, ensuring a proportional feel force whatever amount of trim was applied, this is further discussed below Magnetic brakes: The magnetic brake is an electrical device, which as its name suggests acts like a brake when electrical power is applied, and allows free movement when electrical power is switched off. The magnetic brake is securely mounted to the aircraft structure, and the body fixing of the gradient unit is connected to an arm fitted onto the brake rod. When electrical power is applied, the brake engages and provides a fixed point against which the gradient spring can operate, providing a feel force against the movement of the cockpit controls. Trim release button on cyclic pitch stick grip The magnetic brake is usually de-energised by a switch on the cyclic pitch stick, or in some cases the collective pitch lever, called a trim release switch that is usually marked TRIM REL. The illustration shows the button as fitted to a cyclic pitch stick handgrip

142 When the TRIM REL button is pressed the appropriate magnetic brake is released allowing free movement of the entire control system, including the gradient unit, to establish a new neutral datum for the system. Once any control movement has been made to trim the aircraft, the TRIM REL button is released and the magnetic brake is re-energised and holds the gradient unit and control system in this new position. Because the gradient unit was moved as a unit, rather than just an input occurring at the moveable rod end, it will provide a proportionate feel force from this new position. Any movement of the control system from this position will produce a feel or centring force of an equal amount for movements in either direction. Using the magnetic brake system the pilot(s) can apply trim changes to the aircraft, compensating for any out of trim tendencies, the trim movement being applied directly by manual input on the cockpit controls whilst temporarily disengaging the magnetic brake and gradient unit. Whilst the trim release button is pressed the pilot(s) retain full control of the aircraft, but as there is no feel force present, must be careful not to apply harsh or rapid control movements that would overstress the structure or transmissions Electrical trim actuators: An alternative to the magnetic brake is the electrical trim actuator, in this system an electrical actuator is placed between the gradient unit and its structural mounting. Operation of a spring loaded trim switch on the cyclic pitch stick normally mounted facing the pilot and operated by the thumb, will supply a current to the actuator causing it to extend or retract placing an input into the control system. When the desired flying attitude is obtained the switch is released and returns to its central position by the action of its spring. The illustration shows the position of the trim switch. When the actuator is operated, because of its position in relation to the gradient unit, as shown in the illustration, feel is unaffected and proportional feel force is provided from this new stick position

143 Electrical trim actuator arrangement In some aircraft the electrical trim actuator and gradient spring unit are combined into one unit, as shown below

144 The trim actuators used, are usually of the dual type, which incorporate two electrical motor and gearbox assemblies, in normal operation both function. If one of the motors should fail a clutch mechanism attached to its drive gear will disengage, and the remaining motor will continue to provide trim movement. The electrical trim actuators will normally provide some indication of position, so that during rigging of flight controls the neutral position of the actuator can be established. The combined actuator and gradient unit, shown above has a small transparent window on the top of the motor casing, as the actuator extends or retracts a coloured indicator will move up or down the tube to provide this indication

145 Other forms of trim actuator position indicators vary from a simple annular groove cut into the actuator rod, usually filled with paint of a contrasting colour that indicates the neutral or central position, to an analogue gauge on the instrument panel which indicates units of trim applied. 2. FIXED AND ADJUSTABLE STABILIZERS: Fixed wing aeroplanes use a tail plane or horizontal stabiliser to provide stability about the lateral axis (Pitching) in forward flight, in a similar way the helicopter also requires such a stabiliser. Some helicopters have a fixed stabiliser mounted on the rear of the tail boom, but even in the early days of helicopter development it was realised that moving the stabiliser in response to aircraft speed or control input provided advantages. It is generally the case that helicopters with multi-bladed main rotors will only require fixed stabilisers, whilst those of the two-bladed, semi rigid teetering main rotor design require adjustable stabilisers, because of the greater attitude changes that these cause in the helicopter fuselage. Stabilisers will be found on the entire conventional, single main rotor, helicopter types, but are not required on tandem or side-by-side rotor helicopters. A horizontal stabilizer helps in longitudinal stability, with its efficiency greater the further it is from the centre of gravity. It is also more efficient at higher airspeeds because lift is proportional to the square of the airspeed. Since the speed of a gyroplane is not very high, manufacturers can achieve the desired stability by varying the size of the horizontal stabilizer, changing the distance it is from the centre of gravity, or by placing it in the propeller slipstream the synchronized elevator installation consists of two elevator assemblies, a horn assembly, and two support sets. The horn assembly is mounted horizontally through the sides of the tailboom. The horn is secured to the structure by supports which serve as bearings for the horns rotational movement. A control arm on the horn provides attachment for linkage from fore and aft cyclic control system at swashplate. Each elevator is a horizontal airfoil section built up on a spar tube, which is inserted into a projecting end of the horn assembly and secured by two bolts Fixed stabilizers: There are several configurations of fixed stabilisers, and they may consist of a single aerofoil shaped panel mounted horizontally on the rear of the tail boom, or a pair of such panels, which are fitted one on each side of the tail boom. On some helicopters vertical stabiliser panels are fitted to the outboard edge of the horizontal element, to provide additional directional stability about the normal or vertical axis

146 In forward flight, when sufficient free stream airflow velocity is present, the stabilisers provide an aerodynamic force, in the same way as a conventional aerofoil. This force can provide the necessary downward effort on the tail boom to counteract the helicopters tendency to pitch forward (nose down) during forward flight. Such a pitching movement would expose more of the upper fuselage to the free stream airflow and would cause greater form drag, limiting the helicopters forward speed, as more power would be required to overcome this increased drag. Symmetrical fixed horizontal stabilizers

147 Asymmetrical fixed horizontal stabilizer (KAI Surion helicopter) 2.2. Adjustable stabilizers: The ability to move the stabiliser in forward flight, offers the advantage that it can be set to the best angle for any given speed or control input. This will ensure that the helicopter derives the greatest benefit from the horizontal stabiliser at all forward speeds above that at which it becomes effective. The simplest form of adjustable stabiliser is that which is connected directly to the cyclic control system. Whenever fore and aft (pitching), cyclic inputs are made the stabiliser will also be moved, ensuring that it provides the correct restoring moment, keeping the helicopter fuselage at the desired angle for that mode of flight. Connection may be either by cable or push/pull tube system, with the stabiliser being mounted on a torque tube that runs through the tail boom mounted in bearings. The system will require rigging (a process of ensuring that the correct angles are achieved for the amount of control input) as part of the main flight control system

148 A more complex system is one that can adapt the amount and direction of stabiliser movement, in response not only to control system input, but also aircraft speed, changes in power and movement of the Centre of Gravity. This system will produce a relatively flat pitch attitude for the helicopter fuselage, in virtually all modes of forward flight. To achieve this, a control system is required that can change its input to the stabiliser in response to many differing inputs. In the case of the Bell model 214ST this is achieved by use of a wholly fly-by-wire system. In this helicopter the stabiliser is capable of 10 degrees nose up and 15 degrees nose down pitch, and has a marked effect upon the aircraft attitude, it operates in a similar way to a fixed wing aircraft elevator, and is referred to as such rather than a stabiliser. Bell 214ST helicopter The elevator fly-by-wire systems are duplicated so that in the event of a failure of one system, its actuator or sensors the other will continue to operate

149 Sensors are used to transmit collective and cyclic pitch changes, airspeed and elevator position to the dual elevator control amplifiers. When operating, the amplifiers will produce an output that will cause their respective electrical elevator actuator to move the elevator to the desired position. As the elevator actuators move, the elevator position transducers provide a feedback loop, and the input signal will be nulled when the desired position is reached and all movement will stop. More information will be done at the next lesson

150 X. SYSTEM OPERATION: MANUAL, HYDRAULIC, ELECTRICAL AND FLY-BY-WIRE 1. MANUAL SYSTEM OPERATION: 1.1. Introduction: Very few, if any, of the helicopters within the current PART-66 category weight limit of 5700 kg MTOM, use purely manual input systems for flying controls. However, there were a number of aircraft that have in the past used such systems, and it is right that these should be included here for completeness. Even when considering power operated systems, as described in the next section, some part of the system is still operated manually, up to the point where the power actuator takes over Screw jacks (or jackscrews): Even in small helicopters the loads imposed upon the rotating rotors are high, therefore, only in the smallest aircraft would we expect to see simple systems that connect cables or rods directly to the operating rods at main and tail rotors. It is more normal to use screw jacks in both locations for a number of reasons: They provide a mechanical advantage, allowing greater loads to be applied than originally input to the system. Use generally ACME thread patterns; this means that loads imposed on the rotors cannot move the system, therefore providing irreversibility. Screw threads also mean that very small and accurate inputs can be transmitted to the rotors. Screw jacks can be used to transfer the mechanical movement through an angle. Typical screw jack: Screw jacks are relatively simple devices, consisting basically of a rod with an external thread running through a rotating sleeve that has a matching internal thread. The rod may be splined to the outer casing or keyed to an inner rod, so that it makes a linear movement only and cannot rotate. Operation of the sleeve by rotating will cause the rod to move in or out of the screw jack body because of the interaction of the threaded assemblies

151 Screw jacks can be assembled in a number of different configurations. The first major configuration divides the jacks into translators and rotators. A translating jack has a lifting shaft that moves through the gear box. A nut is integrated with the worm gear such that the worm gear and nut rotate together. Screw jack description When the lift shaft is held to prevent rotation, the lift shaft will move linearly through the gear box to move the load. A rotating jack has a lift shaft that turns moving a nut. The lift shaft is fixed to the worm gear. This causes the load, which is attached to the travel nut, to move along the lift shaft: Both rotators and translators have an upright and inverted configuration

152 1.3. Cables, pulleys, chains and push rods: Mechanical flight control systems are the most basic designs. They were used in early aircraft and currently in small aircrafts where the aerodynamic forces are not excessive. The flight control systems use a collection of mechanical parts such as rods, cables, pulleys and sometimes chains to transmit the forces of the cockpit controls to the surfaces. Typical mechanical flight control system 2. HYDRAULIC SYSTEM OPERATION: 2.1. Hydro-mechanical: In this system the control signals are relayed through a series of cables and linkages to mechanically position the servo valve

153 Typical hydro-mechanical actuator

154 The complexity and weight of a mechanical flight control systems increases considerably with size and performance of the aircraft. Hydraulic power overcomes these limitations. With hydraulic flight control systems aircraft size and performance are limited by economics rather than a pilot's strength. A hydro-mechanical flight control systems has 2 parts: The mechanical circuit The hydraulic circuit The mechanical circuit links the cockpit controls with the hydraulic circuits. Like the mechanical flight control systems, it is made of rods, cables, pulleys, and sometimes chains. Hydro-mechanical detailed system

155 The hydraulic circuit has hydraulic pumps, 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 servo valves control the movement of the actuators. Hydro-mechanical operation systems: The power control unit is the main component in a power operated control system and provides all of the force necessary to move a control surface, with the pilot only having to supply a small force to operate a servo valve

156 It consists of a jack ram/piston arrangement, which is fixed to the aircraft structure, hydraulic fluid, inlet/outlet ports, and a jack body. These parts form a hydraulic actuator, which is controlled by a servo (control) valve and is connected via a control run to the flight deck controls. When the valve is displaced in either direction from its neutral position it allows hydraulic fluid under pressure to pass to one side of the piston, and opens a return path from the other side. For example a rearward movement of the control column will cause the servo valve to move to the left

157 Since the jack is fixed in position the resulting pressure differential across the piston will cause the jack body to move to the left, which in turn will deflect the control surface upwards via a mechanical linkage. The body continues to move until it centralises itself on the servo valve, i.e. returning it to its neutral position. At this point the hydraulic fluid will be trapped either side of the jack and will form a hydraulic lock. This in turn will maintain the control surface rigidly in its selected position, and it will continue to remain so, irrespective of the aerodynamic loads acting on it, until the servo valve is repositioned by further flight deck control inputs

158 This is alternatively known as an irreversible control system. Conversely if the control column is moved forward the sequence of operations will be reversed, i.e. if the servo valve moves to the right, the jack body will move to the right, and the control surface will deflect downwards. Some power control units also operate in response to electrical inputs from the Autopilot and Auto-stabilisation systems when they are engaged. On transport aircrafts, the control surfaces are usually hydraulically activated (although some control surfaces may be electrically actuated) and are powered from the aircraft s main hydraulic systems. Due to the importance of the flying control systems the surfaces are also normally powered by at least two independent hydraulic systems. 3. ELECTRO-MECANICALSYSTEM OPERATION (FLY-BY-WIRE): In this system the control signals are measured by electrical transducers whose output is amplified and then relayed to electrically position the servo valve. This is commonly known as a fly-by-wire (FBW) system. Classic flight control systems are heavy and require careful routing of flight control cables through the helicopter using systems of pulley and cranks. Both systems often require redundant backup, which further increases weight. Furthermore, both have limited ability to compensate for changing aerodynamic conditions. By using computers and electrical linkages, designers can save weight and improve reliability. Electronic fly-by-wire systems can respond more flexibly to changing aerodynamic conditions, by tailoring flight control surface movements so that aircraft response to control inputs is consistent for all 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 prevent a stall, or can stop the pilot from overstressing the aircraft. Detailed fly-by-wire system and operation: Electronic fly-by-wire systems can respond more flexibly to changing aerodynamic conditions, by tailoring flight control surface movements so that helicopter response to control inputs is consistent for all flight conditions

159 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 prevent a stall, or can stop the pilot from overstressing the airframe. A fly-by-wire system literally replaces physical control of the aircraft with an electrical interface. The pilot's commands are converted to electronic signals, and flight control computers determine how best to move the actuators at each control surface to provide the desired response. Those actuators initially are usually hydraulic, but electric actuators have been investigated. The main concerns with fly-by-wire systems is reliability. While traditional mechanical or hydraulic control systems usually fail gradually, the loss of all flight control computers will immediately render the helicopter uncontrollable. For this reason, most fly-by-wire systems incorporated redundant computers and some kind of mechanical or hydraulic backup. This may seem to negate some advantages of fly-by-wire, but the redundant systems can be simpler, lighter, and offer only limited capability since they are for emergency use only. a) Actuator Control Electronic (ACE): Each electrical control system is equipped with one Actuator Control Electronic (ACE) which is composed of several analogue computers (PFCU) of different types. Each computer is a fail safe unit with similar hardware and software design. Each computer has: a control processing channel. The control processing channel operates the control surface. a monitor processing channel. The monitor channel receives feedback information from the sensors attached to the control surface. Each channel has separated power supply units. The horizontal stabilizer is the fixed horizontal wing section at the rear of an aircraft. Pitch trim (the nose up/down tendency of the aircraft) is achieved by rotating the entire horizontal stabilizer. The pitch tendency of the plane is in trim when the pilot can take his hands off the shifter, without any noticeable change in up or down movement. The horizontal stabilizer provides stability for the aircraft

160 b) Example of horizontal stabilizer system: In this example, the horizontal stabilizer subsystem consists of: 1 trim control panel, 2 motor actuator control electronic units (MACE), 1 MACE I, 1 MACE II, 1 horizontal stabilizer actuator / stabilizer setting mechanism (SSM), 2 stabilizer angular position sensor units (SAPS)

161 The horizontal stabilizer subsystem is equipped with two MACEs of two different types 1 MACE I 1 MACE II Both types of MACE are fail-safe units with dissimilar hardware and software design. Each MACE is a digital computer with a control and monitor processing channel. All functions of the system application are software controlled and fully automatic after power-up. The MACEs have separated power supply units for control and for monitor channel. The MACEs control the horizontal stabilizer actuator (SSM). 1. MACE function: The MACE receives signals from the trim control panel or the autopilot. The signal is demodulated within the MACE and in normal mode transferred to the PFCU. In the PFCU the initial signal is adapted to the current situation. The adapted signal is transferred back to the MACE. The MACE transmits the signal to the horizontal stabilizer actuator (SSM). in direct mode the MACE transmits the signal to the horizontal stabilizer actuator (SSM) directly. The signal operates the horizontal stabilizer configuration. 2. Control channel: The control channel in a MACE controls the configuration of the horizontal stabilizer. 3. Monitor Channel: The monitor channel in a MACE monitors the control channel. The monitor channel reacts in case of malfunction to a fail-safe mode function. The fail-safe mode prevents critical behavior of the system

162

163 XI. ARTIFICIAL FEEL 1. ARTIFICIAL FEEL SYSTEMS: In a manually operated flying control system the aerodynamic loads acting on a control surface are fed directly back through the control runs to provide stick force or feel on the flight deck controls. The loads thus vary depending on control surface deflection and airspeed. In the case of a power operated flying control system there is however no direct linkage between the control surface and the flight deck controls. In fact the only force felt is that associated with the movement of a servo valve, and the effort provided by the pilot therefore bears no direct relationship to the actual loads acting on the control surface. These loads are alternatively dissipated through the aircraft structure via the body of a dedicated power control unit, thereby relieving the pilot of all control loads. Consequently, to prevent over-controlling and overstressing of an aircraft, some form of artificial feel is incorporated in the control system, so that the forces experienced are representative of a manually controlled aircraft. A suitable feel unit must therefore be capable of producing an opposing force that varies with aired and control surface deflection. On transport category aircraft the requisite feel forces are provided by; spring or Pitot-static Q feel units, or in some cases a combination of both. Artificial feel systems normally also incorporate a self centring mechanism, so that if the flight deck controls are released they will automatically return to their neutral position, and will also centralise the control surface A simple spring feel units: This is the simplest form of artificial feel unit and is normally fitted in the operating linkage between the flight deck controls and the power control unit. It is designed so that any flight deck control movement is firstly made against string tension, so the larger the movement, the greater the opposite spring force. For example if the control column is moved rearward the left-hand side of the spring in the feel unit will be compressed in proportion to the control column movement and subsequent deflection of the control surface

164 When the control column is centralised the spring unit off-loads itself, thereby centralising the linkage and returning the flying control surface to its neutral position. This type of feel unit by itself may be adequate at low airspeeds, but at higher airspeeds greater resistance to flight deck control movement is needed to prevent over-stressing the aircraft. This is because the amount of feel only varies in proportion to control surface deflection, and takes no account of airspeed Q feel units: These units like spring feel units are fitted in the operating linkage between the flight deck controls and the power control unit

165 A basic Q feel unit consists of a diaphragm with static pressure acting on one side and Pitot pressure on the other, with the difference between the two being dynamic pressure. The unit is also arranged so that movement of the flight deck controls in either direction deflects the diaphragm against pitot pressure. For example if the dynamic pressure increases due to an increase in forward airspeed (IAS), the forces required to move the right deck controls would similarly increase. Conversely, a reduction in forward airspeed will cause the load in the flight deck controls to decrease. This system therefore ensures that the stick forces vary during flight in proportion to varying loads acting on the control surfaces. These units however tend to be very large, so the

166 sensing pressures are alternatively used to operate a piston subjected to hydraulic pressure, thereby providing hydraulic Q feel In this system artificial feel is supplied hydraulically, enabling the unit itself to be much smaller. Like the spring feel unit the Q feel unit also incorporates a self-centring mechanism that operates when the flight deck controls are released. In practice this type of unit is typically used on most transport category helicopter in the elevator control systems, but is usually operated in conjunction with a spring feel unit. Two feel unit s normally act together to resist movement of the flight deck controls from their neutral position. Unless some form of artificial feel is incorporated into power operated flight control systems, pilots could inadvertently overstress the structure, rotors and transmissions. Unlike fixed wing aircraft, helicopters do not require complex feel systems that react to changes in speed, altitude and mach number, but rather a unit that will provide a feel force that is relative to the deflection applied to the controls. The feel force, which may also be referred to as a Centring Force because of its natural tendency to centre the control to which it is attached when pilot input force is removed, may most easily be produced by using a simple spring assembly, often referred to as a Gradient Unit

167 2. MAIN ROTOR CONTROL SYSTEM FEEL GRADIENT UNITS: In its simplest form the gradient unit may consists of two springs, preloaded against each other, or a single spring that has drive fittings at each end, fitted into a casing. As the flight controls are moved the springs are either compressed or stretched, depending upon what movement is applied to the unit. The gradient units are usually not an integral part of the flying control system, but an additional part upon which the control run, and any input force applied by the pilot, operates. The illustration shows this principle. Typical gradient unit arrangement The illustration below shows an exploded view of a gradient unit that uses a single spring; the spring is fitted over a rod and has retainers fitted at each end. This arrangement will ensure that any movement in either direction will be resisted by the spring, exerting a feel force into the manually operated parts of the flight control system, which will be proportional to the deflection of the spring

168 The more force that is applied, the more feel force will be felt. Gradient unit, exploded view Movement of the flight control system is transmitted onto the gradient unit by means of a fitting mounted centrally on the gradient unit body, as shown in the illustration below. One end of the gradient unit is attached to a fixed bracket, whilst the other is mounted onto an idler. The idler will provide support to the end of the unit, whilst allowing it to move in response to control system inputs. In this arrangement, an axial load will be applied to the idler, and an eccentric load, ie. a load that varies depending upon the amount of movement applied and therefore deflection of the spring, is applied to the fitting

169 Gradient unit mountings It is normal to find gradient units fitted into all flying control inputs; one for collective inputs, and two in the cyclic systems, one each for pitch (fore & aft) and roll (side to side lateral) inputs. By producing a control centring force (feel) the gradient units will increase the breakout force required to initially move the controls. They will also increase running friction, once the control is moving, and this running friction will continue to increase as the control system is moved further from the central position. If we now consider the situation of an aircraft which requires a constant control deflection, to correct some out of balance moment (ie. aircraft out of trim) then it will be seen that the pilot would be constantly subjected to the feel force, leading to fatigue

170 Therefore a system is required, that allows trimming of the aircraft, without changing the feel force. This may be achieved by ensuring that whenever the controls, collective and cyclic, are set to a new neutral position to keep the aircraft flying in trim, the gradient unit is moved by the same amount to establish a new neutral position for the feel. 3. TAIL ROTOR CONTROL SYSTEM FEEL YAW PEDAL DAMPERS: Sometimes referred to as Pedal Dampers or Yaw Dampers, although in frequent use the latter term must be used with caution as yaw dampers are used on fixed wing aircraft for an entirely different purpose. The yaw damper is fitted to provide three functions, which are to: prevent too rapid an input being applied to the tail rotor control system, provide a feel force, provide a trim holding force. If there were no feel unit provided in the tail rotor flying control system, control inputs could be made rapidly, which would lead to over control, and damage to the aircraft structure. Basic yaw pedal damper

171 The purpose of the yaw pedal damper is to limit the rate at which a control input can be made, the unit also provides for trim adjustments to be applied to the tail rotor control system. Yaw pedal dampers usually operate by using hydraulic fluid as the damping medium, working on the dashpot principle, that is; the rate of movement is proportional to the rate of flow of fluid internally. The body comprises of a cylinder into which is fitted a balanced piston (ie. a piston ram is fitted at either side). The cylinder is filled with hydraulic fluid, and both ends are connected together by a small channel, drilled into the cylinder body, into which a restriction device is fitted. A small reservoir is usually fitted to the unit to provide fluid top-up in the event of small leaks occurring. Movement of the piston in one direction will displace fluid from one end of the cylinder and transfer it to the other. Restriction of the fluid flow will control the rate at which it transfers, and therefore the rate at which the control system may be moved. The illustration shows this principle. Yaw pedal damper operation Movement of the piston to the right, causes fluid to be transferred from cylinder A to cylinder B, via the transfer channel and restriction device. The restriction sets the rate of fluid transfer, and therefore the rate at which an input can be applied to the tail rotor control system. Additionally, because the flow of fluid is restricted, there is an increase in pressure in cylinder A, which is felt at the yaw pedals providing a feel force. However, this is not like the feel force felt on main rotor controls, as it has no centring function, removing the input force from the yaw pedals will allow an equalisation of fluid pressure across the yaw pedal damper piston, and it will remain in this new position. Some helicopters incorporate units that have this type of damper used in conjunction with a spring to provide an additional feel force

172 In many examples of these components the restriction device is adjustable, but it is normal for the flow rate to be adjusted whilst the unit is removed from the aircraft, as a special jig and timing equipment is required. In fact, this very aspect of its operation is what provides the trim holding function. When the pilot wishes to make an adjustment to aircraft trim by use of the tail rotor, it is only necessary to apply a force to the yaw pedals moving them to the to the desired position, and upon removing his input force they remain in this new selected position, until operated again. Safety features: There are a number of occasions when the operation or failure of the yaw pedal damper could endanger the helicopter and crew. For example even though it is fitted in parallel to the flight control push/pull rods, the damper could prevent control system operation if the small transfer channel or restriction device became blocked and fluid transfer could not occur. Additionally in an emergency the pilot(s) may need to apply a much faster rate of input to the tail rotor control system than would normally be required, but if a safety device is not fitted this would not be possible. To prevent either of these undesirable events occurring two pressure relief valves are fitted to the damper piston, one operating in each direction. Yaw pedal damper piston relief valves Should the normal transfer channel be blocked, or a much higher force than normal be applied there would be a rapid rise in fluid pressure. Once the pressure exceeded the setting of the relief valve, fluid transfer would take place through the relief valve and piston

173 Many yaw pedal dampers, especially those that do not have a transparent reservoir and are filled using pressure replenishment rather than a drip feed oil can, incorporate a third relief valve, which is fitted to relieve any excess pressure from the unit if it is overfilled, if it operates fluid is vented from it. Fluid transfer from the integral reservoir is usually by gravity fill, a small check valve at the base of the reservoir to cylinder channel closes off the supply from the reservoir whenever yaw pedal movement causes an increase in internal pressure. If this valve becomes jammed or blocked, fluid may flow into the reservoir, rather than the other side of the cylinder, causing a spongy operation of the damper, and allow a much faster operation time. Yaw pedal damper reservoir and check valve 4. FEEL INTEGRATION WITH AUTOFLIGHT SYSTEMS: When autoflight systems are incorporated into helicopter flying control systems, the feel function of these devices is not usually required, as the auto flight system requires no feel and is programmed to prevent control inputs being made too rapidly. However, should the pilot or co-pilot wish to make control corrections, or to take control from the auto flight system, then there is a need to immediately remove all auto flight inputs, and restore feel force feedback

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