portion between approximately 25% and 70% of the radius outward from the center. This portion is often referred to as the autorotative or driving region." Aerodynamic forces along this portion of the blade tend to speed up the blade rotation.
The inner 25% of the rotor blade, referred to as the "stall region," operates above its maximum angle of attack (stall angle), thereby contributing little lift but considerable drag which tends to slow the blade rotation.
The outer 30% of the rotor blade is known as the "propeller or driven region." Aerodynamic forces here result in a small drag force which tends to slow the tip portion of the blade.
The aerodynamic regions as described above are for vertical autorotations. During forward flight autorotations, these regions are displaced across the rotor disk to the left.
Rotor r.p.m. stabilizes when the autorotative forces (thrust) of the "driving region" and the autorotative forces (drag) of the "driven region" and the "stall region" are equal.
Forward speed during autorotative descent permits a pilot to incline the rotor disk rearward, thus causing a flare. The additional induced lift created by the greater volume of air momentarily checks forward speed as well as descent. The greater volume of air acting on the rotor disk will normally increase rotor r.p.m. during the flare. As the forward speed and descent rate near zero, the upward flow of air has practically ceased and rotor r.p.m.
again decreases; the helicopter settles at a slightly increased rate but with reduced forward speed. The fare enables the pilot to make an emergency landing on a definite spot with little or no landing roll or skid.
HELICOPTER AXES OF FLIGHT As a helicopter maneuvers through the air, its attitude in relation to the ground changes. These changes are described with reference to three axes (figure 2-43) of fight: (1) Vertical, (2) longitudinal, and (3) lateral.
Movement about the vertical axis produces yaw, a nose swing. (or change in direction) to the right or left. This is controlled by the directional-control pedals. The various methods of achieving directional control were discussed earlier in this section.
Movement about the longitudinal axis is called roll. This movement is effected by moving the cyclic pitch control to the right or left. The cyclic pitch control is similar to the control stick of a conventional aircraft. It acts through a mechanical linkage (figure 2-44) to change the pitch of each main-rotor blade during a cycle of rotation.
holl Longitudinal axis Vertical axis
• Lateral axis Pitch FIGURE 2-43. Axes of flight.
Swash plate FIGURE 2-41. Cyclic pitch control mechanism.
The rapidly rotating rotor blades create a disk area that can be tilted in any direction with respect to the supporting rotor mast. Horizontal movement 55
is controlled by changing the direction of tilt of the main rotor to produce a force in the desired direction.
Movement about the lateral axis produces a noseup or nosedown attitude. This movement is effected by moving the cyclic pitch control fore and aft.
The collective pitch control (figure 2-45) varies the lift of the main rotor by increasing or decreasing the pitch of all blades at the same time. Raising the collective pitch control increases the pitch of the blades, thereby increasing the lift. Lowering the control decreases the pitch of the blades, causing a loss of lift. Collective pitch control is also used in coordination with cyclic pitch control to regulate the airspeed of the helicopter.
FIGURE 2-45. Collective pitch control mechanism.
Many factors determine the amount of lift available in helicopter operation. Generally speaking, the pilot has control of two of these. One is the pitch angle of the rotor blades; the other is the power de.
livered to the rotor, represented by r.p.m. and manifold pressure. By controlling the pitch angle of the rotor blades, the pilot can establish the vertical flight of the helicopter. By manipulating the throttle control, a constant engine speed can be maintained regardless of the increase or decrease in blade pitch. The throttle is mounted on the collective pitch grip and is operated by rotating the grip. The throttle is synchronized with the main-rotor pitch control in such a manner that an increase of pitch increases power and a decrease in pitch decreases power. A complete control system of a conventional helicopter is shown in figure 2-46.
HIGH-SPEED AERODYNAMICS Developments in aircraft and powerplants are yielding high-performance transports with capabilities for very high speed flight. Many significant differences arise in the study of high-speed aerodynamics when compared with low-speed aerodynamics.
It is quite necessary, therefore, that persons associated with commercial aviation be familiar with the nature of high-speed airflow and the peculiarities of high-performance airplanes.
General Concepts of Supersonic Flow Patterns At low flight speeds, air experiences small changes in pressure which cause negligible variations in density, greatly simplifying the study of low-speed aerodynamics. The flow is called incompressible since the air undergoes small changes in pressure without significant changes in density. At high flight speeds, however, the pressure changes that take place are quite large and significant changes in air density occur. The study of airflow at high speeds must account for these changes in air density and must consider that the air is compressible, or that there are compressibility effects.
The speed of sound is very important in the study of high-speed airflow. The speed of sound varies with the ambient temperature. At sea level, on a standard day, the speed of sound is about
661.7 knots (760 m.p.h.).
As a wing moves through the air, local velocity changes occur which create pressure disturbances in the airfow around the wing. These pressure disturbances are transmitted through the air at the speed of sound. If the wing is traveling at low speed, the pressure disturbances are transmitted and extend indefinitely in all directions. Evidence of these pressure disturbances is seen in the typical subsonic flow pattern illustrated in figure 2-47 where upwash and flow direction change well ahead of the wing leading edge.
If the wing is traveling above the speed of sound, the airhow ahead of the wing is not infuenced by the pressure field of the wing, since pressure disturbances cannot be propagated faster than the speed of sound. As the flight speed nears the speed of sound, a compression wave forms at the leading edge and all changes in velocity and pressure take place quite sharply and suddenly. The airflow ahead of the wing is not influenced until the air molecules are suddenly forced out of the way by the wing. Evidence of this phenomenon is seen in the typical supersonic flow pattern shown in figure
2-48.
Compressibility effects depend not on airspeed, but rather on the relationship of airspeed to the 56
Rotor blade pitch control horn Flapping axis - Blade thrust bearing .
Upper link (adjustable for tracking blades) Rotating rocker arm (upper links move up together when center of rocker arms is moved up by collective pitch sleeve, and move cyclically when swash plates are tilted) Upper swash plate Lower link (tilts and rotates) Lower swash plate (mounted on gimbals which allow -Pilot's collective pitch lever it to tilt but not to rotate or to move vertically) Throttle linkage - - Throttle twist grip Pilot's control stick Gear case (phantom) fixed to fuselage Co-pilot's collective Collective pitch sleeve vertical positioning mechanism and throttle control Shaft , Co-pilot's control Collective pitch sleeve (moves up and stick down while rotating with shaft) FICURE 2-46. Control system of conventional helicopter.
Flow direction changes well ahead of leading edge Subsonic accelerating flow Subsonic decelerating flow FIcURE 2-47. Typical subsonic flow pattern, subsonic wing.
Trailing edge oblique shock wave Initial expansion wave Leading edge oblique shock wave Final expansion wave Flow direction does not change ahead of leading edge FICURE 2-18. Typical supersonic flow pattern, supersonic wing.
speed of sound. This relationship is called Mach number, and is the ratio of true airspeed to the speed of sound at a particular altitude.
Compressibility effects are not limited to flight speeds at and above the speed of sound. Since any airplane is made up of aerodynamic shapes, air accelerates and decelerates around these shapes and attains local speeds above the flight speed. Thus, an aircraft can experience compressibility effects at flight speeds well below the speed of sound. Since it is possible to have both subsonic and supersonic flows on the airplane at the same time, it is best to define certain regimes of flight. These approximate regimes are defined as follows:
(1) Subsonic--flight Mach numbers below
0.75.
(2) Transonic--flight Mach numbers from
0.75 to 1.20.
(3) Supersonic--fight Mach numbers from
1.20 to 5.00.
(4) Hypersonic-flight Mach numbers above
5.00.
While the flight Mach numbers used to define these regimes are approximate, it is important to appreciate the types of flow existing in each area.
In the subsonic regime, subsonic airflow exists on all parts of the aircraft. In the transonic regime, the flow over the aircraft components is partly subsonic and partly supersonic. In the supersonic and hypersonic regimes, supersonic flow exists over all parts of the aircraft. Of course, in supersonic and hypersonic flight some portions of the boundary layer are 58
subsonic, but the predominating flow is still supersonic.
Difference Between Subsonic and Supersonic Flow In a subsonic flow every molecule is affected more or leas by the motion of every other molecule in the whole field of flow. At supersonic speeds, an air molecule can influence only that part of the flow contained in the Mach cone formed behind that molecule.
The peculiar differences between subsonic flow and supersonic flow can best be seen by considering airflow in a closed contracting/expanding tube, as depicted in figure 2-48.
Decreasing air density Increasing air density Contracting tube Expanding tube Increasing air velocity Decreasing air velocity
(a) Subsonic flow Increasing air density Decreasing air dersity Contracting tube Expanding tube Decreasing air velocity Increasing air velocity
(b) Supersonic flow FIGURE 2-49. Comparison of subsonic and supersonic airflow through a closed tube.
Unlike subsonic flow, a supersonic airstream accelerates along an expanding tube, causing the air density to decrease rapidly to compensate for the combined effects of increased speed and increased cross sectional area.
Unlike subsonic flow, a supersonic airstream decelerates along a contracting tube, causing the air density to increase rapidly to compensate for the combined effects of decreased speed and decreased cross sectional area.
In order to clarify these fundamenta. points, ngure 2-50 lists the nature of the two types of tubes. An understanding of figures 2-49 and 2-50 is essential if one is to grasp the fundamentals of supersonic flow.
Contracting tube Expanding tube Accelerates and Decelerates and Subsonic fow rarefies slightly compresses slightly Decelerates and Accelerates and Supersonic flow compresses greatly rarefies greatly FIGURE 2-50. High-speed flows.
TYPICAL SUPERSONIC FLOW PATTERNS With supersonic flow, all changes in velocity, pressure, temperature, density, and flow direction take place suddenly and over a short distance. The areas of flow change are distinct and the phenomena causing the flow change are called wave formations. All compression waves occur abruptly and are wasteful of energy. Compression waves are more familiarly known as shock waves. Expansion waves result in smoother flow transition and are not wasteful of energy like shock waves. Three types of waves can take place in supersonic flow: (1) The oblique (inclined angle) shock wave (compression), (2) the normal (right angle) shock wave (compression), and (3) the expansion wave (no shock). The nature of the wave depends on the Mach number, the shape of the object causing the flow change, and the direction of flow.
A supersonic airstream passing through the oblique shock wave experiences these changes:
(1) The airstream is slowed down. Both the velocity and the Mach number behind the wave are reduced, but the flow is still supersonic.
(2) The flow direction is changed so that the airstream runs parallel to the new surface.
(3) The static pressure behind the wave is increased.
(4) The static temperature behind the wave is increased (and hence the local speed of sound is increased).
(5) The density of the airstream behind the wave is increased.
(6) Some of the available energy of the airstream (indicated by the sum of dynamic and static pressure) is dissipated by conversion into unavailable heat energy.
Hence, the shock wave is wasteful of energy.
The Normal Shack Wave If a blunt-nosed object is placed in a supersonic airstream, the shock wave which is formed is detached from the leading edge. The detached wave 59
- Subsonic region ~ Oblique shock Normal shock Normal shock - Subsonic region- -Oblique shock
(a) Blunt-nosed object (b) Wedge or cone FIGURE 2-51. Normal shock-wave formation with a detached wave.
also occurs when a wedge or cone half angle exceeds some critical value. Figure 2-51 shows the formation of a normal shock wave in the above two cases. Whenever a shock wave forms perpendicular to the free stream flow, the shock wave is termed a normal (right angle) shock wave and the flow immediately behind the wave is subsonic. No matter how high the free stream Mach number may be, the flow directly behind a normal shock is always subsonic. In fact, the higher the supersonic free stream Mach number (M) is in front of the normal shock wave, the lower the subsonic Mach number is aft of the wave. For example, if M, is 1.5, M2 is 0.7;
while if My is 2.6, M2 is only 0.5. A normal shock wave forms immediately in front of any relatively blunt object in a supersonic airstream, slowing the airstream to subsonic so that the airstream may feel the presence of the blunt object and thus flow around it. Once past the blunt nose, the airstream may remain subsonic or it may accelerate back to supersonic, depending on the shape of the nose and the Mach number of the free stream.
A normal wave may also be formed when there is no object in the supersonic airstream. It so happens that whenever a supersonic airstream is slowed to subsonic without a change in direction, a normal shock wave forms as the boundary between the supersonic and subsonic regions. This is why airplanes encounter compressibility effects before the flight speed is sonic. Figure 2-52 illustrates the manner in which an airfoil at a high subsonic speed has local fow velocities which are supersonic. As the local supersonic flow moves aft, a normal shock wave forms so that the flow may return to subsonic and rejoin the subsonic free stream at the trailing edge without discontinuity. The transition of flow from subsonic to supersonic is smooth and is not accompanied by shock waves if the transition is made gradually with a smooth surface. The transition of flow from supersonic to subsonic without direction change always forms a normal shock wave.
A supersonic airstream passing through a normal shock wave experiences these changes:
(1) The airstream is slowed to subsonic. The local Mach number behind the wave is approximately equal to the reciprocal of the Mach number ahead of the wave. For example, if the Mach number ahead of the wave is 1.25, the Mach number of the flow behind the wave is about 0.8 (more exactly 0.81264).
(2) The airfow direction immediately behind the wave is unchanged.
(3) The static pressure behind the wave is greatly increased.
(4) The static temperature behind the wave is 60
Local Mach numbers Expansion wave
0.75 0.80 0.95 0.80 0,75
0.70-
0.70 (A) Flight Mach number 0.70 ITTI Supersonic region - 1.00
0.90 0.95 1.00 1.02 0.95 0.90 - Normal shock (a) Supersonic flow around a sharp corner
0.85. 0.85 Series of expansion waves (B) Flight Mach number 0.85 FIGURE 2-52. Normal shock-wave formation on an airfoil in a subsonic airstream.
greatly increased (and hence the local speed of sound is increased).
(5) The density of the airstream behind the wave is greatly increased.
(6) The available energy of the airstream (indicated by the sum of dynamic and static pressure) is greatly reduced. The normal shock wave is very wasteful of energy.
The Expansion Wave If a supersonic airstream is turned away from the preceding flow, an expansion wave is formed. The How around a corner shown in figure 2-53 does not cause sharp, sudden changes in the airflow except at the corner itself and thus is not actually a shock wave. A supersonic airstream passing through an expansion wave experiences these changes:
(1) The supersonic airstream is accelerated.
The velocity and Mach number behind the wave are greater.
(2) The flow direction is changed so that the airstream runs parallel to the new surface, provided separation does not occur.
(3) The static pressure behind the wave is decreased.
(4) The static temperature behind the wave is decreased (and hence the local speed of sound is decreased).
(5) The density of the airstream behind the wave is decreased.
(6) Since the flow changes in a rather gradual
(b) Supersonic flow around a rounded corner FIcURE 2-53. Expansion wave formation.
manner, there is no shock and no loss of energy in the airstream. The expansion wave does not dissipate airstream energy.
A summary of the characteristics of the three principal wave forms encountered with supersonic flow is shown in figure 2-54.
Figure 2-55 shows the wave pattern for a conventional blunt-nosed subsonic airfoil in a supersonic stream. When the nose is blunt, the wave must detach and become a normal shock wave immedi ately ahead of the leading edge. Since the flow just behind a normal shock wave is always subsonic, the airfoil's leading edge is in a subsonic region of very high static pressure, static temperature, and density.
In supersonic flight, the zero lift of an airfoil of some finite thickness includes a wave drag. Wave drag is separate and distinct from drag due to lift.
The thickness of the airfoil has an extremely powerful effect on the wave drag. The wave drag varies as the square of the thickness ratio (maximum thickness divided by the chord). For example, if the thickness is cut by one-half, the wave drag is cut by three-fourths. The leading edges of supersonic 61
Effect on static Flow direction Effect on velocity pressure, static Effect on available Type of wave change and Mach number temperature and energy density into a Flow Decreased, but still Increase Decrease Oblique shock wave corner supersonic No change Decreased to subsonic Great increase Great decrease Normal shock wave Flow around Increased to higher Decrease No change a corner supersonic (no shock) Expansion wave 22/A21 FIGURE 2-54. Supersonic wave characteristics.
shock Oblique - Subsonic region Expansion waves Supersonic region Oblique shock Supersonic Normal shock Supersonic free stream free stream Expansion waves Oblique shock Supersonic region Oblique shock Oblique shock Wave patter (not to scale) FICURE 2-55. The conventional subsonic airfoil in supersonic flow.
shapes must be sharp. If they are not, the wave formed near the leading edge is a strong detached normal shock wave.
Once the flow over the airfoil is supersonic, the aerodynamic center of the surface is located approximately at the 50% chord position. This contrasts with the subsonic location of the aerodynamic center, which is near the 25% chord position.
During supersonic flow all changes in velocity, Mach number, static pressure, static temperature, 62
density, and flow direction take place quite suddenly through the various wave forms. The shape of the object, the Mach number, and the required flow direction change dictate the type and strength of the wave formed.
Any object in subsonic flight which has some finite thickness or is producing lift must have local velocities on the surface which are greater than the free stream velocity. Hence, compressibility effects can be expected to occur at flight speeds which are less than the speed of sound. The transonic regime of flight provides the opportunity for mixed subsonic and supersonic local velocities and accounts for the first significant effects of compressibility.
As the flight speed approaches the speed of sound, the areas of supersonic flow enlarge and the shock waves move nearer the trailing edge. The boundary layer may remain separated or may re-attach, depending much upon the airfoil shape and angle of attack. When the flight speed exceeds the speed of sound, a bow wave suddenly appears in front of the leading edge with a subsonic region behind the wave. The normal shock waves move to the trailing edge. If the flight speed is increased to some higher supersonic value, the bow wave moves closer to the leading edge and inclines more downstream, and the trailing edge normal shock waves become oblique shock waves.
Of course, all components of the aircraft are affected by compressibility in a manner somewhat similar to that of the basic airfoil (the empennage, fuselage, nacelles, and so forth).
Since most of the difficulties of transonic flight are associated with shock-wave-induced flow separation, any means of delaying or lessening the shockinduced separation improves the aerodynamic characteristics. An aircraft configuration can make use of thin surfaces of low aspect ratio with sweepback to delay and reduce the magnitude of transonic force divergence. In addition, various methods of boundary layer control, high-lift devices, vortex generators, and so forth, may be applied to improve transonic characteristics. For example, the mounting of vortex generators on a surface can produce higher local surface velocities and increase the kinetic energy of the boundary layer, Thus, a more severe pressure gradient (stronger shock wave) would be necessary to produce the unwanted airflow separation.
A vortex generator is a complementary pair of amall, low aspect ratio (short span in relation to chord) airfoils mounted at opposite angles of attack to each other and perpendicular to the aerodynamic surface they serve. Figure 2-56 shows the airfoils and the airflow characteristics of a vortex generator. Like any airfoil, those of the generator develop lift. In addition, like any airfoil of especially low aspect ratio, the airfoils of the generator also develop very strong tip vortices. These tip vortices cause air to flow outward and inward in circular paths around the ends of the airfoils. The vortices generated have the effect of drawing high-energy air from outside the boundary layer into the slower moving air close to the skin. The strength of the vortices is proportional to the lift developed by the airfoils of the generator.
Airflow Lift "Airfoil Top View Upper edge of Tip vortex boundary layer
- Airfoil Rear view FIcURE 2-56. Wing vortex generator.
Vortex generators serve two distinctly different purposes, depending on the aerodynamic surface upon which they are mounted. Rows of vortex generators located on the upper surface of the wing just upstream of the ailerons delay the onset of drag divergence at high speeds and also aid in maintaining aileron effectiveness at high speeds. In contrast, rows of vortex generators mounted on both sides of the vertical fin just upstream of the rudder prevent flow separation over the rudder dur- 63
ing extreme angles of yaw which are attained only when rudder application is delayed after an engine loss at very low speeds. In addition, rows of vortex generators placed on the underside (and occasionally on the upper surface) of the horizontal stabilizer just upstream of the elevators prevent flow separation over the elevators at very low speeds.
In summary, vortex generators on wing surfaces improve high-speed characteristics, while vortex generators on tail surfaces, in general, improve low-speed characteristics.
Control Surfaces The control surfaces used on aircraft operating at transonic and supersonic flight speeds involves some important considerations. Trailing edge control surfaces can be affected adversely by the shock waves formed in flight above the control surface critical Mach number. If the airflow is separated by the shock wave, the resulting buffet of the control surface can be very objectionable. Installation of vortex generators can reduce buffet caused by shock-induced flow separation. In addition to the buffet of the surface, the change in the pressure distribution due to separation and shock-wave location can create very large changes in control surface hinge moments. Such large changes in hinge moments produce undesirable control forces which may require the use of an irreversible control system. An irreversible control system employs powerful hydraulic or electric actuators to move the control surfaces, hence the airloads developed on the surfaces cannot be felt by the pilot. Suitable feedback must be synthesized by bungees, "q" springs, bobweights, and so forth.
AERODYNAMIC HEATING « he redu long When air flows over any aerodynamic surface, certain reductions in velocity take place which produce corresponding increases in temperature. The greatest reduction in velocity and increase in temperature occur at the various stagnation points on the aircraft. Of course, smaller changes occur at other points on the aircraft, but these lower temperatures can be related to the ram temperature rise at the stagnation point. While subsonic flight does not produce temperatures of any real concern, supersonic flight can create temperatures high enough to be of major importance to the airframe, fuel system, and powerplant.
Higher temperatures produce definite reductions in the strength of aluminum alloys and require the use of titanium alloys and stainless steels. Continued exposure at elevated temperatures further reduces strength and magnifies the problems of creep failure and structural stiffness.
The effect of aerodynamic heating on the fuel system must be considered in the design of a supersonic airplane. If the fuel temperature is raised to the spontaneous ignition temperature, the fuel vapors will burn in the presence of air without the need of an initial spark or flame.
Turbojet engine performance is adversely affected by high compressor inlet air temperature.
The thrust output of the turbojet, obviously, is some function of the fuel flow. But the maximum allowable fuel flow depends on the maximum permissible turbine operating temperature. If the air entering the engine is already hot, less fuel can be added in order to avoid exceeding turbine temperature limits.
FLIGHT CONTROL SYSTEMS Three types of control systems commonly used are: (1) The cable, (2) push-pull, and (3) the torque tube system. The cable system is the most widely used because deflections of the structure to which it is attached do not affect its operation.
Many aircraft incorporate control systems that are combinations of all three types.
Flight Control System Hardware, Mechanical Linkage, and Mechanisms The systems which operate the control surfaces, tabs, and flaps include flight control system hardware, linkage, and mechanisms. These items connect the control surfaces to the cockpit controls. Included in these systems are cable assemblies, cable guides, linkage, adjustable stops, control surface snubber or locking devices, surface control booster units, actuators operated by electric motors, and actuators operated by hydraulic motors.
Cable Assembly The conventional cable assembly consists of flexible cable, terminals (end fittings) for attaching to other units, and turnbuckles. Information concerning conventional cable construction and end fittings is contained in Chapter 6 of the Airframe and Powerplant Mechanics General Handbook, AC 65-9A.
At each regular inspection period, cables should be inspected for broken wires by passing a cloth along their length and observing points where the cloth snags. To thoroughly inspect the cable, move the surface control to its extreme travel limits. This 64
will reveal the cable in pulley, fairlead, and drum aréas. If the surface of the cable is corroded, relieve cable tension. Then carefully force the cable open by reverse twisting, and visually inspect the interior for corrosion. Corrosion on the interior strands of the cable indicates failure of the cable and requires replacement of the cable. If there is no internal corrosion, remove external corrosion with a coarseweave rag or fiber brush. Never use metallic wools or solvents to clean flexible cable. Metallic wools imbed dissimilar metal particles, which cause further corrosion. Solvents remove the internal cable lubricant, which also results in further corrosion.
After thoroughly cleaning the flexible cable, apply corrosion-preventive compound. This compound preserves and lubricates the cable.
Breakage of wires occurs most frequently where cables pass over pulleys and through fairleads. Typical breakage points are shown in figure 2-57. Control cables and wires should be replaced if worn, distorted, corroded, or otherwise damaged.
rel is threaded with left-hand threads inside one end and right-hand threads inside the other. When adjusting cable tension, the cable terminals are screwed into either end of the barrel an equal distance by turning the barrel. After a turnbuckle is adjusted, it must be safetied. The methods of safetying turnbuckles is discussed in Chapter 6 of the Airframe and Powerplant Mechanics General Handbook, AC 65-9.
Cable Connectors In addition to turnbuckles, cable connectors are used in some systems. These connectors enable a cable length to be quickly connected or disconnected from a system. Figure 2-58 illustrates one type of cable connector in use. This type is connected or disconnected by compressing the spring.
Spring connector FIGURE 2-57. Typical breakage points.
Lockclad cable is used on some large aircraft for all long, straight runs. It consists of the conventional flexible steel cable with aluminum tubing swaged to it to lock the cable inside the tubing.
Lockclad cable construction has certain advantages.
Changes in tension due to temperature are less than with conventional cable. Furthermore, the amount of stretch at a given load is less than with conventional cable.
Lockclad cables should be replaced when the covering is worn through, exposing worn wire strands;
is broken; or shows worn spots which cause the cable to bump when passing over fairlead rollers.
Turnbuckles The turnbuckle is a device used in cable control systems to adjust cable tension. The turnbuckle bar- FICURE 2-58. Spring type of cable connector.
HYDRAULIC OPERATED CONTROL SYSTEMS As the airspeed of late model aircraft increased, actuation of controls in flight became more difficult.
It soon became apparent that the pilot needed assistance to overcome the airflow resistance to control movement. Spring tabs which were operated by the conventional control system were moved so that the airflow over them actually moved the primary control surface. This was sufficient for the aircraft operating in the lowest of the high speed ranges (250-300 mph).
For high speeds a power assist (hydraulic) control system was designed.
Conventional cable or push pull rod systems are installed and are tied into a power transmission quadrant. With the system activated, the pilot's effort is used to open valves thereby directing hydraulic fluid to actuators, which are connected to the control surfaces by control rods. The actuators move the control surface to the desired flight condition. Reversing the input effort moves the control surface in the opposite direction.
Manual Control The control system from the cockpit is connected by a rod across the power transmission quadrant to the control actuating system. During manual operation, the pilot's effort is transmitted from the control wheel through this direct linkage to the control surface. Those aircraft which do not have the manual reversion system may have as many as three sources of hydraulic power-primary, back-up and auxiliary. Any or all of the primary controls may be operated by these systems.
SPLIT FAIRLEAD Gust Lock A cam on the control quadrant shaft engages a spring-loaded roller for the purpose of centering and neutralizing the controls with the hydraulic system off (aircraft parked). Pressure is trapped in the actuators and since the controls are neutralized by the cam and roller, no movement of the control surfaces is permitted.
CABLE GUIDES Cable guides (figure 2-59) consist primarily of fairleads, pressure seals, and pulleys.
A fairlead may be made from a nonmetallic material, such as phenolic or a metallic material such as soft aluminum. The fairlead completely encircles the cable where it passes through holes in bulkheads. or other metal parts. Fairleads are used to guide cables
RUBSTRIP
- SOLID FAIRLEAD
RETAINING RINGS - • CONTROL CABLE
• BULKHEAD GROOVE UNPRESSURIZED -BRACKET
PRESSURIZED LPULLEY AIR SEAL . BULKHEAD- - GUARD PIN FIGURE 2-59. Cable guides.