# Flight Mechanics ## Airfoil Characteristics Having a firm understanding of airfoil design and the various results of those airfoil designs is necessary for the understanding of aerodynamics. Different airfoil designs and airfoil terminology are explained in this section to better understand the mechanics of flight. **Helicopters and conventional aircraft are able to fly due to aerodynamic forces produced when air passes around the airfoil. An airfoil is any surface such as a wing, propeller, rudder, or even a trim tab, which provides aerodynamic force when it interacts with a moving stream of air. (FAA-H-8083-25B)** Airfoils are most often associated with production of lift. Airfoils are also used for stability (fin), control (elevator), and thrust or propulsion (propeller or rotor). Certain airfoils, such as rotor blades, combine some of these functions. Airfoils are carefully structured to accommodate a specific set of flight characteristics. *** ## Airfoil Terminology Table 1-1, provides airfoil terms and their definitions common to all aircraft. The first four terms describe the shape of an airfoil. The remaining terms describe development of aerodynamic properties.
TermsDefinitions
Blade SpanThe length of the rotor blade from the point of rotation to the tip of the blade.
Wing SpanThe maximum distance from wingtip to wingtip. (FAA-H-8083-25B)
Chord LineAn imaginary straight line drawn through an airfoil from the leading edge to the trailing edge. (FAA-H-8083-25B)
ChordThe length of the chord line from the leading edge to the trailing edge; it is the characteristic longitudinal dimension of the airfoil section. (McCormick, 1994)
Mean Camber LineA line drawn halfway between the upper and lower surfaces. The chord line connects the ends of the mean camber line. Camber refers to the curvature of the airfoil and may be considered as the curvature of the mean camber line. The shape of the mean camber is important for determining the aerodynamic characteristics of an airfoil section. Maximum camber (displacement of the mean camber line from the chord line) and its location help to define the shape of the mean camber line. The location of maximum camber and its displacement from the chord line are expressed as fractions or percentages of the basic chord length. By varying the point of maximum camber, the manufacturer can tailor an airfoil for a specific purpose. The profile thickness and thickness distribution are important properties of an airfoil section.
Leading-Edge RadiusThe radius of curvature given the leading edge shape.
Flight-Path VelocityThe speed and direction of the airfoil passing through the air. For fixed-wing (FW) airfoils, flight-path velocity is equal to true airspeed (TAS). For helicopter rotor blades, flight-path velocity is equal to rotational velocity, plus or minus a component of directional airspeed. (FAA-H-8083-21B)
Angle of Incidence (FW Aircraft)The angle between the airfoil chord line and the longitudinal axis or other selected reference plane of the airplane. (FAA-H-8083-25B)
Angle of Incidence (Rotary-Wing Aircraft)The angle between the chord line of a main or tail-rotor blade and rotational relative wind (tip-path plane). It is usually referred to as blade pitch angle. For fixed airfoils, such as vertical fins or elevators, the angle of incidence is the angle between the chord line of the airfoil and a selected reference plane of the helicopter.
Center of PressureThe point along the chord line of an airfoil through which all aerodynamic forces are considered to act. Since pressures vary on the surface of an airfoil, an average location of pressure variation is needed. As the angle of attack (AOA) changes, these pressures change, and the center of pressure moves along the chord line. (FAA-H-8083-25B)
Aerodynamic CenterThe point along the chord line where all changes to lift effectively take place. If the center of pressure is located behind the aerodynamic center, the airfoil experiences a nose-down pitching moment. The use of this point by engineers eliminates the problem of center of pressure movement during AOA aerodynamic analysis. (McCormick, 1994)
*** ## Airfoil Types The two basic types of airfoils are symmetrical and nonsymmetrical. ### **Symmetrical** The symmetrical airfoil (Figure 1-10) is distinguished by having identical upper and lower surface designs, with the mean camber line and chord line being coincident, producing zero lift at zero angle of attack (AOA). Symmetrical designs have advantages and disadvantages. One advantage is that the center of pressure remains relatively constant under varying angles of attack, reducing the twisting force exerted on the airfoil. Additionally, symmetrical designs afford ease of construction and reduced cost. However, the disadvantages include less lift production at a given AOA compared to a nonsymmetrical design and undesirable stall characteristics.
### Nonsymmetrical (Cambered) The nonsymmetrical airfoil (Figure 1-11) has different upper and lower surface designs, with a greater curvature of the airfoil above the chord line than below. The mean camber line and chord line are not coincident. The nonsymmetrical airfoil design produces useful lift even at negative angles of attack. Nonsymmetrical designs have advantages and disadvantages. Advantages: * More lift production at a given angle of attack compared to a symmetrical design. * Improved lift-to-drag ratio. * Better stall characteristics. Disadvantages: * The center of pressure travel can move up to 20 percent of the chord line, creating undesirable torque on the airfoil structure. * Greater production costs.
## Blade Twist (Rotary-Wing Aircraft) Because of the lift differential along the blade, it should be designed with a twist to alleviate internal blade stress and distribute the lifting force more evenly along the blade. Blade twist provides higher pitch angles at the root where velocity is low and lower pitch angles nearer the tip where velocity is higher. This increases the induced air velocity and blade loading near the inboard section of the blade. ## Airflow and Reactions in the Rotor System The different interactions between airfoil mechanics moving through the atmosphere are described by various terms in this section. The movement of air relative to the rotor system and the resulting movements after the interaction are critical to understanding rotary-wing flight. ### Relative Wind Knowledge of relative wind (Figure 1-12) is essential for an understanding of aerodynamics and its practical flight application for the aviator. Relative wind is the direction of the airflow produced by an object moving through the air. (FAA-H-8083-25B) The relative wind for an airplane in forward flight flows in a direction parallel with and opposite to the direction of flight; therefore, the actual flight path of the airplane determines the direction of the relative wind.
### Rotational Relative Wind The rotation of rotor blades as they turn about the mast produces rotational relative wind (Figure 1-13). The term "rotational" refers to the method of producing relative wind. Rotational relative wind flows opposite the physical flight path of the airfoil, striking the blade at 90 degrees to the leading edge and parallel to the plane of rotation, and is constantly changing in direction during rotation. Rotational relative wind velocity is highest at blade tips, decreasing uniformly to zero at the axis of rotation (center of the mast).
### Induced Flow (Downwash) At flat pitch, air leaves the trailing edge of the rotor blade in the same direction it moved across the leading edge; no lift or induced flow is being produced. As blade pitch angle is increased, the rotor system induces a downward flow of air through the rotor blades, creating a downward component of air that is added to the rotational relative wind. Because the blades are moving horizontally, some of the air is displaced downward. The blades travel along the same path and pass a given point in rapid succession. Rotor blade action changes the still air to a column of descending air. This downward flow of air is called induced flow (downwash). Induced flow is the component of air flowing vertically through the rotor system resulting from the production of lift. (FAA-H-8083-21B) It is most pronounced at a hover under no-wind conditions (Figure 1-14).
#### Resultant Relative Wind Resultant relative wind (Figure 1-15) at a hover is airflow from rotation that is modified by induced flow. (FAA-H-8083-21B) This wind is inclined downward at some angle and opposite the effective flight path of the airfoil, rather than the physical flight path (rotational relative wind). The resultant relative wind also serves as the reference plane for the development of lift, drag, and total aerodynamic force (TAF) vectors on the airfoil. When the helicopter has horizontal motion, airspeed further modifies the resultant relative wind. The airspeed component of relative wind results from the helicopter moving through the air. This airspeed component is added to or subtracted from the rotational relative wind, depending on whether the blade is advancing or retreating in relation to helicopter movement. Introduction of airspeed relative wind also modifies induced flow. Generally, the downward velocity of induced flow is reduced. The pattern of air circulation through the disk changes when the aircraft has horizontal motion. As the helicopter gains airspeed, the addition of forward velocity results in decreased induced flow velocity. This change results in improved efficiency (additional lift) being produced from a given blade pitch setting. Section V further covers this process.
### Up Flow (Inflow) Up flow (inflow) is airflow approaching the rotor disk from below as a result of some rate of descent. Up flow also occurs as a result of blades flapping down or an updraft, which alter the angle of attack. *** ## Rotor Blade Angles Mechanical and aerodynamic angles of airfoils have various effects on engine power demands and aerodynamic properties. Angle of incidence is one angle that is easier to control because it is a mechanical angle. Angle of attack is harder to control because of its relationship with the atmosphere. Correlating the relationship between these two angles expands the understanding of power management and permission planning. ### Angle of Incidence Angle of incidence is the angle between the chord line of a main or tail rotor blade and the rotational relative wind of the rotor system (tip-path plane) (Figure 1-16, page 1-11). It is a mechanical angle rather than an aerodynamic angle and is sometimes referred to as blade pitch angle. In the absence of induced flow, angle of attack (AOA) and angle of incidence are the same. However, whenever induced flow, up flow (inflow), or airspeed modifies relative wind, then AOA is different from angle of incidence. Collective input and cyclic feathering change angle of incidence. A change in angle of incidence changes AOA, which changes the coefficient of lift, thereby changing the lift produced by the airfoil. ### Angle of Attack AOA is the angle between the airfoil’s chord line and relative wind. (FAA-H-8083-21B) The relative wind associated with AOA is the resultant relative wind. AOA is an aerodynamic angle (Figure 1-6). It can change with no change in angle of incidence. Several factors may change the rotor blade AOA. Aviators control some of those factors; others occur automatically due to rotor system design. Aviators adjust AOA through normal control manipulation; even with no aviator input, however, AOA changes as an integral part of travel of the rotor blade through the rotor-disk arc. This continuous process of change accommodates rotary-wing flight. Aviators have little control over blade flapping and flexing, gusty wind, or turbulent air conditions. AOA is one of the primary factors determining amount of lift and drag produced by an airfoil.
### Effects of Airflow As angle of attack (AOA) is increased, there is a greater acceleration of air atop the airfoil. This results in a larger pressure differential between the top and bottom of the airfoil, producing a larger aerodynamic force. If AOA is increased beyond a critical angle, flow across the top of the airfoil will be disrupted, boundary layer separation will occur, and a stall results. When this occurs, lift rapidly decreases, drag rapidly increases, and the airfoil ceases to fly. *** ## Rotor Blade Actions Different types of rotor systems react in different ways while rotating around the central axis. Rotor system design and their specific design attributes aid in the understanding of aircraft limits and capabilities. Understanding the individual rotor blade actions of a specific rotor system in flight strengthens the overall understanding of how that system works as a whole. ### Rotation Rotation of rotor blades is the most basic movement of the rotor system and produces rotational relative wind. During hovering, rotation of the rotor system produces airflow over the rotor blades. Figure 1-17, illustrates a typical rotor system with an arbitrary rotor diameter of 40 feet and rotor speed of 320 revolutions per minute (RPM) used to demonstrate rotational velocities. In this example, blade tip velocity is 670 feet per second, or 397 knots. At the blade root—nearer the rotor shaft or blade attachment point—blade speed is much less as the distance traveled at the smaller radius is much less. Halfway between the root and tip (point A in Figure 1-17), blade speed is 198.5 knots, or one-half tip speed. Blade speed varies according to the distance or radius from the center of the main rotor shaft. While the airspeed differential between root and tip is extreme, the lift differential is more extreme because lift varies as the square of the velocity (see lift equation on page 1-26). As velocity doubles, lift increases four times. The lift at point A in Figure 1-17 would be only one-fourth as much as lift at the blade tip—assuming the airfoil shape and AOA are the same at both points.
### Feathering Feathering is the action that changes the pitch angle of the rotor blades by rotating them around their feathering (spanwise) axis. (FAA-H-8083-21B) (Figure 1-18).
### Collective Feathering Collective feathering changes the angle of incidence equally and in the same direction on all rotor blades simultaneously. This action changes angle of attack (AOA), which changes coefficient of lift ($$CLCL​$$) and affects the overall lift of the rotor system. ### Cyclic Feathering Cyclic feathering changes the angle of incidence differentially around the rotor system. Cyclic feathering creates a differential lift in the rotor system by changing the AOA differentially across the rotor system. Aviators use cyclic feathering to control the attitude of the rotor system. It is the means to control the rearward tilt of the rotor (blowback) caused by flapping action and (along with blade flapping) counteract dissymmetry of lift (section V). Cyclic feathering causes the attitude of the rotor disk to change but does not change the amount of lift the rotor system is producing. ### Flapping The vertical movement of a blade about a flapping hinge is called flapping. (FAA-H-8083-21B) It occurs in response to changes in lift due to changing velocity or cyclic feathering (Figure 1-19). No flapping occurs when the tip-path plane is perpendicular to the mast. The flapping action alone, or along with cyclic feathering, controls dissymmetry of lift (section V). Flapping is the primary means of compensating for dissymmetry of lift.
Flapping also allows the rotor system to tilt in the desired direction in response to cyclic input. See Figure 1-20, Figure 1-21, Figure 1-22 (page 1-14), and Figure 1-23 (page 1-14) for depictions of flapping as it occurs throughout the rotor disk.
#### Semirigid Rotor System In the semirigid rotor system, a blade is not free to flap independently of the other blades because they are affixed through the hub. The blades form one continuous unit moving together on a common teetering hinge. This hinge allows one blade to flap up as the opposite blade flaps down, although blade flex limits the amount of blade flapping. In the fully articulated rotor system, blades flap individually about a horizontal hinge pin. Therefore, each blade is free to move up and down independently from all of the other blades. Aircraft design can reduce excessive flapping in several ways; for example, a forward tilt of the transmission and mast helps minimize flapping and installation of a synchronized elevator or stabilator (utility helicopter (UH)-60/attack helicopter (AH)-64) helps maintain the desired fuselage attitude to reduce flapping. #### Rigid Rotor System In a rigid rotor system, each blade flaps about flexible sections of the root, which are rigidly attached to the rotor hub. This rotor system is mechanically simple, but structurally complex because operating loads must be absorbed in bending rather than through hinges. Rigid rotor systems tend to behave like fully articulated systems through aerodynamics, but lack flapping or lead/lag hinges. ### Lead and Lag (Hunting) The fore (lead) and aft (lag) movement of the rotor blade in the plane of rotation is lead and lag. This movement responds to changes in angular velocity. This rotor blade action can only occur in a fully articulated rotor system, in which the system is equipped with a vertical hinge pin (drag hinge) or elastomeric bearing providing a pivot point for each blade to move independently. In directional flight, pitch angle and the AOA of the blades are constantly changing. These changes in AOA cause changes in blade drag. To prevent undue bending stress on the blades and blade root, the blade is free to move fore and aft in the plane of rotation. The need to lead and lag is due to the Coriolis force. It is governed by the law of conservation of angular momentum. This law states a body will continue to have the same rotational momentum unless acted on by an outside force. Two factors determine the rotational (angular) momentum—distance of the center of gravity (CG) from the center of rotation and rotational speed. If the CG moves closer to the center of rotation, the rotational speed must increase. If the CG moves farther away from the axis of rotation, rotational velocity will decrease.
### Lead As a blade flaps up, the CG of the blade (point C in figure 1-24) moves inboard toward the axis of rotation, producing a smaller radius of travel. The blade speeds up in reaction to this CG change, causing the blade to lead a few degrees ahead of its normal position in the tip-path plane (point D in figure 1-24). This motion relieves stress that would have been imposed on the blade structure. ### Lag As a blade flaps down, the CG of the blade (point A in figure 1-24) moves outboard away from the axis of rotation, producing a greater radius of travel. The blade slows down in reaction to this CG change, causing the blade to lag a few degrees behind its normal position in the tip-path plane (point B in figure 1-24). This motion relieves stress that would have been imposed on the blade structure. ### Semirigid Rotor System Because of the design (under slung) of the semirigid rotor system, no change occurs in the travel radius of the CG of the blade associated with blade flapping (figure 1-25, page 1-16). The angular velocity of the blade does not change. Drag does impose significant stresses on the blade roots; a drag brace is normally installed at the blade root to absorb some of these bending forces.
### Rigid Rotor System The rigid rotor system behaves more like a fully articulated rotor system, and its characteristics are based on how much flexibility the blade’s construction contains. The blades’ flexible root provides physical behavior similar to mechanical or elastomeric hinges, allowing the neutralizing forces in the same manner as lead or lag and flapping do with other rotor designs. The rigid rotor system, sometimes referred to as bearingless, 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. Rigid rotor systems tend to behave like fully articulated systems through aerodynamics but lack flapping or lead/lag hinges. Instead, the blades accommodate these motions by bending. They cannot flap or lead/lag, but they can be feathered. These rotor systems offer 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. Other advantages of the rigid rotor include 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. 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. *** ## Helicopter Design and Control Helicopter design continues to evolve due to advancements in materials and building processes, resulting in a more stable platform with inherent control features. Understanding the forces acting on the aircraft and the design features counteracting those forces is crucial for comprehending helicopter flight. ### Gyroscopic Precession The phenomenon of precession occurs in rotating bodies, where an applied force manifests 90 degrees after application in the direction of rotation. Although precession is not a dominant force in rotary-wing aerodynamics, aviators and designers must consider it, as turning rotor systems exhibit some characteristics of a gyro. In the context of a typical rotor disk, applying a downward force at a certain point results in a downward movement of the disk at a different point due to precession. This phenomenon is illustrated in Figure 1-26, showing the effects of precession on a rotor disk when force is applied at a given point. Understanding gyroscopic precession is essential for helicopter pilots and designers to anticipate and manage its effects during flight operations.
Table 1-2 depicts reactions to forces applied to a spinning rotor disk by control input or wind gusts. **Table 1-2: Aircraft reaction to forces** | Force Applied to Rotor Disk | Aircraft Reaction | | --------------------------- | ----------------- | | Up at nose | Roll right | | Up at tail | Roll left | | Up on right side | Nose up | | Up on left side | Nose down | This behavior explains some fundamental effects occurring during various helicopter maneuvers. For example, the helicopter behaves differently when rolling into a right turn than when rolling into a left turn. During roll into a right turn, the aviator must correct for a nose-down tendency to maintain altitude. This correction is required because precession causes a nose-down tendency. During a roll into a left turn, precession causes a nose-up tendency. Aviator input required to maintain altitude is different during a left versus right turn as gyroscopic precession acts in opposite directions. ### **Rotor Head Control** The control of the rotor head traditionally involves two main inputs such as a collective and cyclic pitch control. Different airframe designs incorporate more sophisticated systems that interact with these two main controls for enhanced controllability. ### **Cyclic and Collective Pitch** Aviator inputs to collective and cyclic pitch controls are transmitted to the rotor blades through a complex system. This system consists of levers, mixing units, input servos, stationary and rotating swashplates, and pitch-change arms (figure 1-27). In its simplest form, movement of collective pitch control causes stationary and rotating swashplates mounted centrally on the rotor shaft to rise and descend. The movement of cyclic pitch control causes the swashplates to tilt; the direction of tilt is controlled by the direction in which the aviator moves the cyclic (figure 1-28).
### **Tilted Swashplate Assembly** Figure 1-29, illustrates a swashplate tilted 2 degrees at two positions, points B and D. Points A and C form the axis about which the tilt occurs. At that axis, the swashplate remains at zero degrees. When the swashplate is moved, pitch-change arms transmit the resulting motion change to the rotor blade. As the pitch-change arms move up and down with each rotation of the swashplate, blade pitch constantly increases or decreases. If the aviator applies cyclic control to tilt the rotor, adding collective pitch does not change the tilt of the swashplate and rotor. It simply moves the swashplate upward so pitch is increased equally on all blades simultaneously, thereby increasing AOA and total lift.
### **Pitch-Change Arms** Figure 1-30 illustrates how pitch-change arms move up and down on the tilted swashplate. The rate of vertical change throughout the rotation is not uniform. Vertical movement is larger during the 30 degrees of rotation at point A than at points B and C. This variation repeats during each 90 degrees of rotation. The rate of vertical movement is lowest at the low and high points of the swashplate and highest when the pitch-change arms pass by the tilt axis of the swashplate.
### **Cyclic Pitch Change** Figure 1-31, shows a change in cyclic pitch (cyclic feathering). This causes rotor blades to climb from point A to point B, then dive or descend from point B to point A. In this way, the rotor is tilted in the direction of desired flight.
To pass through points A and B, the blades must flap up and down on a hinge or teeter on a trunnion. At the lowest flapping point (point A), the blades would appear to be at their lowest pitch angle; at the highest flapping point (point B), they would be at their highest pitch angle. If only aerodynamic considerations were involved, this might be true. However, gyroscopic precession (figure 1-26, page 1-17) causes these points to be separated by 90 degrees of rotation. A cyclic movement decreases blade pitch at one point in the rotor disk while increasing blade pitch by the same amount 180 degrees of travel later. A decrease in lift resulting from a decrease in blade pitch angle and AOA causes the blade to flap down; the blade reaches its maximum downflapping displacement 90 degrees later in the direction of rotation. An increase in lift resulting from an increase in blade pitch angle and AOA causes the blade to flap up; the blade reaches its maximum upflapping displacement 90 degrees later in the direction of rotation. Figure 1-32, page 1-21, shows the resulting change to the rotor disk’s attitude. The cyclic pitch causing blade flap must be placed on the blades 90 degrees of rotation before the lowest and highest flap are desired. This 90 degrees of phase lag due to gyroscopic precession is accounted for when rotors are designed, and it ensures when the cyclic is pushed forward, the action tilts the swashplate assembly to place the cyclic pitch accordingly. To tilt the rotor disk forward, the lowest cyclic pitch on the blade needs to be over the right side of the helicopter and the highest cyclic pitch over the left side. The rotor always tilts in the direction in which the aviator moves the cyclic.
### Typical Design Features Figure 1-33, illustrates a typical design feature used in most four-bladed rotor systems offsetting cyclic control input 90 degrees from where the aviator desires rotor tilt. Rotor control input locations are the left lateral servo (point A), right lateral servo (point B), and fore and aft servo (point C). Each servo is offset 45 degrees from the position corresponding to its name. The fore and aft input servo, for example, is not located at the nose or tail position but at the right front about halfway between the nose and 3 o’clock position. Similarly, the left lateral servo is located halfway between the nose and 9 o’clock position. The right lateral servo is halfway between the tail and 3 o’clock position. Locations of the input servos account for part of the offset the aviator needs to correct for gyroscopic precession. In addition, the rotor blade has a pitch-change horn extending ahead of the blade in the plane of rotation about 45 degrees. A connecting rod, called a pitch-change rod, transmits aviator control inputs from the input servos to the pitch-change horn. The design of the pitch-change horn, coupled with placement of the servo and tilt of the swashplate, provides the total offset.
### Cyclic Pitch Variation Figure 1-34 illustrates typical cyclic pitch variation of a blade through one revolution with cyclic pitch control full forward. Degrees shown are for a typical aircraft rotor system; figures would vary with the type of helicopter. As described in the previous paragraph, the input servos and pitch-change horns are offset. With cyclic pitch control in the full forward position, the blade pitch angle is highest at the 9 o’clock position and lowest at the 3 o’clock position. The pitch angle begins decreasing as it passes the 9 o’clock position and continues to decrease until it reaches the 3 o’clock position. The pitch begins to increase and reaches the maximum pitch angle at the 9 o’clock position. Blade pitch angles over the nose and tail are about equal. Figure 1-34 shows that blades reach a point of lowest flapping over the nose 90 degrees in the direction of rotation from the point of lowest pitch angle. Highest flapping occurs over the tail 90 degrees in the direction of rotation from the point of the highest pitch angle. Simply stated, the force (pitch angle) causing blade flap must be applied to the blade 90 degrees of rotation before the point where the aviator desires maximum blade flap. A pattern similar to figure 1-34 could be constructed for other cyclic positions in the circle of cyclic travel. In each case, the same principles apply. Points of highest and lowest flapping are located 90 degrees in the direction of rotation from the points of highest and lowest blade pitch.
### **Fuselage Hovering Attitude** The hovering attitude varies between aircraft designs. There are a number of reasons why a fuselage has a specific hovering attitude. Aviators should be familiar with the hovering characteristics of their respective aircraft. ### **Single-Rotor Helicopter** The design of most fully articulated rotor systems includes an offset between the main rotor mast and blade attachment point. Centrifugal force acting on the offset tends to hold the mast perpendicular to the tip-path plane (figure 1-35). When the rotor disk is tilted left to counteract the translating tendency, the fuselage follows the main rotor mast and hangs slightly low on the left side. A fuselage suspended under a semirigid rotor system remains level laterally unless the load is unbalanced or the tail rotor gearbox is lower than the main rotor (figure 1-36). The fuselage remains level because there is no offset between the rotor mast and the point where the rotor system is attached to the mast (trunnion bearings). Because trunnion bearings are centered on the mast, the mast does not tend to follow the tilt of the rotor disk during hover. In addition, the mast does not tend to remain perpendicular to the tip-path plane as it does with a fully articulated rotor system. Instead, the mast tends to hang vertically under the trunnion bearings, even when the rotor disk is tilted left to compensate for translating tendency (figure 1-36, point B). Because the mast remains vertical, the fuselage hangs level laterally unless other forces affect it.
When there is forward tilt of the mast, the tail rotor gearbox is probably lower than the main rotor. Main rotor thrust above tail rotor thrust to the right causes the fuselage to tilt laterally left (figure 1-37). Although main rotor thrust to the left is equal to tail rotor thrust to the right, it acts at a greater distance from the CG, creating a greater turning moment on the fuselage. This is more pronounced in helicopters with semirigid rotor systems than those with fully articulated rotor systems. Tail rotor thrust acting at the plane of rotation of the main rotor would not change the attitude of the fuselage. The main rotor mast in semirigid and fully articulated rotor systems may be designed with a forward tilt relative to the fuselage. During forward flight, forward tilt provides a level longitudinal fuselage attitude, resulting in reduced parasite drag; during hover, it results in a tail-low fuselage attitude.
### **Tandem-Rotor Helicopter** In tandem-rotor helicopters, the forward and aft rotor systems are tilted forward due to transmission mounting design. This tilt helps decrease excessive nose-low attitudes in forward flight and allows the aircraft to ground or water taxi forward. Most tandem-rotor helicopters hover at a nose-high attitude of about 5 degrees. Some models automatically compensate for this nose-high attitude through automatic programming of the rotor systems. ### **Pendular Action** The fuselage of the helicopter has considerable mass and is suspended from a single point (single-rotor helicopters). It is free to oscillate laterally or longitudinally like a pendulum. Normally, the fuselage follows rules governing pendulums, balance, and inertia. Rotor systems, however, follow rules governing aerodynamics, dynamics, and gyroscopes. These two unrelated systems have been designed to work well together, in spite of apparent conflict. Other factors, such as overcontrolling, cyclic-control response, and shift of attitude, affect the relationship of the rotor system and fuselage. ### **Overcontrolling** Overcontrolling occurs when the aviator moves the cyclic control stick, causing rotor tip-path changes not reflected in corresponding fuselage-attitude changes. Correct cyclic control movements (free of overcontrol) cause the rotor tip-path and fuselage to move in unison. ### **Cyclic Control Response** The rotor response to cyclic control input on a single-rotor helicopter has no lag. Rotor blades respond instantly to the slightest touch of cyclic control. The fuselage response to lateral cyclic is noticeably different from the response to fore and aft cyclic applications. Normally, considerably more fore and aft cyclic movement is required to achieve the same fuselage response as achieved from an equal amount of lateral cyclic. This is not a lag in rotor response; rather as figure 1-38, page 1-25, shows, it is due to more fuselage inertia around the lateral axis than around the longitudinal axis. For single-rotor helicopters, the normal corrective device for the lateral axis is the addition of a synchronized elevator or stabilator attached to the tail boom. This device produces lift forces keeping the fuselage of the helicopter in proper alignment with the rotor at normal flight airspeed. This alignment helps reduce blade flapping and extends the allowable CG range of the helicopter; however, it is ineffective at slow airspeeds.
### **Shift of Attitude** Fuel cells normally have a slight aft CG. As fuel is used, a slight shift to a more nose-low attitude occurs. Because of fuel expenditure and lighter fuselage, cruise attitudes tend to shift slightly lower. As fuel loads are reduced, drag affects the lighter fuselage more, resulting in a slight shift to a more nose-down attitude during flight. --- # Agent Instructions: Querying This Documentation If you need additional information that is not directly available in this page, you can query the documentation dynamically by asking a question. Perform an HTTP GET request on the current page URL with the `ask` query parameter: ``` GET https://229.gitbook.io/229-hub/knowledge/army-aviation-fundamentals/aerodynamics-of-flight/flight-mechanics.md?ask= ``` The question should be specific, self-contained, and written in natural language. The response will contain a direct answer to the question and relevant excerpts and sources from the documentation. Use this mechanism when the answer is not explicitly present in the current page, you need clarification or additional context, or you want to retrieve related documentation sections.