- 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.
in a straight line through or between structural members of the aircraft. Fairleads should never de- straight line. flect the alignment of a cable more than 3° from a Checknut threaded rod end Tube Adjustable Rivets antifriction rod end Adjustable rod end clevis Control or push-pull rod Quadrant Rear spar Torque tube Cable Push-pull rod Horn Rudder aileron interconnect cables B Torque tube Centerline of airplane Rig pin hole Sector Bellcrank D Sector Shaft Control wheel Bearing Drum E Cable drum FIGURE 2-60. Flight control system mechanical linkages.
Pressure seals are installed where cables (or rods) move through pressure bulkheads. The seal grips tightly enough to prevent excess air pressure loss but not enough to hinder movement of the cable. Pressure seals should be inspected at regular intervals to determine that the retaining rings are in place. If a retaining ring comes off, it may slide along the cable and cause jamming of a pulley.
Pulleys are used to guide cables and also to so that the flange of the rod end is interposed between the ball race and the anchored end of the attaching pin or bolt as shown in figure 2-61.
Flange.
change the direction of cable movement. Pulley Anchored bearings are sealed, and need no lubrication other end than the lubrication done at the factory. Brackets fastened to the structure of the aircraft support the pulleys. Cables passing over pulleys are kept in place by guards. The guards are close-fitting to prevent jamming or to prevent the cables from slipping off when they slacken due to temperature vari- Peening ations.
MECHANICAL LINKAGE Various mechanical linkages connect the cockpit controls to control cables and surface controls.
These devices either transmit motion or change the direction of motion of the control system. The linkage consists primarily of control (push-pull) rods, torque tubes, quadrants, sectors, bellcranks, and cable drums.
Control rods are used as links in flight control systems to give a push-pull motion. They may be adjusted at one or both ends. View A of figure 2-60 shows the parts of a control rod, Notice that it consists of a tube having threaded rod ends. An adjustable antifriction rod end, or rod end clevis, attaches at each end of the tube. The rod end, or clevis, permits attachment of the tube to flight control system parts. The checknut, when tightened, prevents the rod end or clevis from loosening.
Control rods should be perfectly straight, unless designed to be otherwise, when they are installed.
The bellcrank to which they are attached should be checked for freedom of movement before and after attaching the control rods. The assembly as a whole should be checked for correct alignment, When the rod is fitted with self-aligning bearings, free rotational movement of the rods must be obtained in all It is possible for control rods fitted with bearings to become disconnected because of failure of the peening that retains the ball races in the rod end.
This can be avoided by installing the control rods FIGURE 2-61. Rod end flange interposed between the bearing race and the end of the attaching bolt.
Another alternative is to place a washer, having a larger diameter than the hole in the fange, under the retaining nut on the end of the attaching pin or bolt.
TORQUE TUBES Where an angular or twisting motion is needed in a control system, a torque tube is installed. View B of figure 2-60 shows how a torque tube is used to transmit motion in opposite directions.
Quadrants, bellcranks, sectors, and drums change direction of motion and transmit motion to parts such as control rods, cables, and torque tubes. The quadrant shown in figure 2-60B is typical of flight control system linkages used by various manufacturers. Figures 2-60C and 2-60D illustrate a bell.
crank and a sector. View E illustrates a cable drum. Cable drums are used primarily in trim tab systems. As the trim tab control wheel is moved clockwise or counterclockwise, the cable drum winds or unwinds to actuate the trim tab cables.
STOPS Adjustable and nonadjustable stops (whichever the case requires) are used to limit the throw-range or travel movement of the ailerons, elevator, and rudder. Usually there are two sets of stops for each 68
of the three main control surfaces, one set being located at the control surface, either in the snubber cylinders or as structural stops (figure 2-62), and the other at the cockpit control. Either of these may serve as the actual limit stop. However, those situated at the control surface usually perform this function. The other stops do not normally contact each other, but are adjusted to a definite clearance when the control surface is at the full extent of its travel. These work as over-ride stops to prevent stretching of cables and damage to the control system during violent maneuvers. When rigging control systems, refer to the applicable maintenance manual for the sequence of steps for adjusting these stops to limit the control surface travel.
through a cable system by a spring-loaded plunger (pin) that engages a hole in the control surface mechanical linkage to lock the surface. A spring connected to the pin forces it back to the unlock position when the cockpit control lever is placed in the "unlock" position. An over-center toggle linkage is used on some other type aircraft to lock the control surfaces.
Control surface locking systems are usually so designed that the throttles cannot be advanced until the control surfaces are unlocked. This prevents taking off with the control surfaces in the locked position.
A typical control lock for small aircraft consists of a metal tube that is installed to lock the control wheel and rudder pedals to an attachment in the cockpit. Such a system is illustrated in figure 2-63.
Stop Rudder horn FIcURE 2-62. Adjustable rudder stops. Lock assembly
CONTROL SURFACE SNUBBERS AND LOCKING
DEVICES Various types of devices are in use to lock the control surfaces when the aircraft is parked or moored. Locking devices prevent damage to the control surfaces and their linkages from gusts and high-velocity winds. Common devices that are in use are the internal locking brake (sector brake) spring-loaded plunger, and external control surface locks.
Internal Locking Devices The internal locking device is used to secure the ailerons, rudder, and elevator in their neutral positions. The locking device is usually operated FIGURE 2-63. Typical control lock assembl or small aircraft.
Control Surface Snubbers Hydraulic booster units are used on some aircraft to move the control surfaces. The surfaces are usually protected from wind gusts by snubbers incorporated into the booster unit. On some aircraft an auxiliary snubber cylinder is connected directly to the surface to provide protection. The snubbers hydraulically check or cushion control surface movement when the aircraft is parked. This prevents wind gusts from slamming the control surfaces into their stops and possibly causing damage.
External Control Surface Locks External control surface locks are in the form of channeled wood blocks. The channeled wood blocks slide into the openings between the ends of the movable surfaces and the aircraft structure. This locke the surfaces in neutral. When not in use, these locks are stowed within the aircraft.
Tension Regulators Cable tension regulators are used in some flight control systems because there is considerable difference in temperature expansion of the aluminum aircraft structure and the steel control cables. Some large aircraft incorporate tension regulators in the control cable systems to automatically maintain a given cable tension. The unit consists of a compression spring and a locking mechanism which allows the spring to make correction in the system only when the cable system is in neutral.
type designed for different kinds of cable, cable sizes, and cable tensions.
One type of tensiometer is illustrated in figure
2-64. With the trigger lowered, place the cable to be tested under the two anvils. Then close the trigger (move it up). Movement of the trigger pushes up the riser, which pushes the cable at right angles to the two clamping points under the anvils. The force that is required to do this is indicated by the dial pointer. As the sample chart beneath the illustration shows, different numbered risers are used with different size cables. Each riser has an identifying number and is easily inserted into the tensiometer.
Pointer lock Riser
AIRCRAFT RIGGING Control surfaces should move a certain distance Anvils in either direction from the neutral position. These movements must be synchronized with the movement of the cockpit controls. The flight control system must be adjusted (rigged) to obtain these requirements.
Generally speaking, the rigging consists of the following: (1) Positioning the flight control system in neutral and temporarily locking it there with rig pins or blocks, and (2) adjusting surface travel, Trigger system cable tension, linkages, and adjustable stops to the aircraft manufacturer's specifications. Sample only Example When rigging flight control systems, certain items of rigging equipment are needed. Primarily, this No. 1 Riser No. 2 No. 3 equipment consists of tensiometers, cable rigging Dia. 1/16 3/32 11/8 Tension 5/32 3/16 7/32 1/4 tension charts, protractors, rigging fixtures, contour Lb.
templates, and rulers. 12 16 21 30 12 20 19 23 29 40 17 26 22 32 Measuring Cable Tension 26 37 30 To determine the amount of tension on a cable, a tensiometer is used. When properly maintained, a tensiometer is 98% accurate. Cable tension is determined by measuring the amount of force needed to 120 46 50 make an offeet in the cable between two hardened steel blocks, called anvils. A riser or plunger is pressed against the cable to form the offset. Several manufacturers make a variety of tensiometers, each FICURE 2-64. Tensiometer.
In addition, each tensiometer has a calibration table (figure 2-64) which is used to convert the dial reading to pounds. (The calibration table is very similar to the sample chart shown below the illustration.) The dial reading is converted to pounds of tension as follows. Using a No. 2 riser (figure 2-64) to measure the tension of a 5/32-in.
diameter cable a reading of "30" is obtained. The actual tension (see calibration table) of the cable is 70 Ibs. Observing the chart, also notice that a No. 1 riser is used with 1/16, 3/32-, and 1/8-in. cable.
Since the tensiometer is not designed for use in measuring 7/32- or 1/4-in. cable, no values are shown in the No. 3 riser column of the chart.
When taking a reading, it may be difficult to see the dial. Therefore, a pointer lock is present on the tensiometer. Push it in to lock the pointer. Then remove the tensiometer from the cable and observe the reading. After observing the reading, pull the lock out and the pointer will return to zero.
Cable rigging tension charts (figure 2-65) are graphic tools used to compensate for temperature variations. They are used when establishing cable tensions in flight control systems, landing gear systems, or any other cable-operated systems.
To use the chart, determine the size of the cable that is to be adjusted and the ambient air temperature. For example, assume that the cable size is 1/8-in. in diameter, that it is a 7 x 19 cable, and the ambient air temperature is 85° F. Follow the 85° F. line upward to where it intersects the curve for 1/8-in. cable. Extend a horizontal line from the point of intersection to the right edge of the chart.
The value at this point indicates the tension (rigging load in pounds) to establish on the cable. The tension for this example is 70 lbs.
Surface Travel Measurement The tools for measuring surface travel primarily include protractors, rigging fixtures, contour templates, and rulers. These tools are used when rig.
ging flight control systems to assure that the desired travel has been obtained.
Values include 10 percent structural deflection Design limit rig load 340
• 320
• 300 280 Rigging 260 load in
• 240 pounds Cable sizes - 220 1/4 7 x 19 • 200 3/16 7 × 19 • 180
- 160 5/32 1 × 19 140 1/8 7 × 19-
• 120 3/32 7 × 7. - 100 1/16 7×7 - 80
- 60
• 40
• 20
- 0 -65-60-50-40-30-20-10 0 10 20 30 40 50 60 70 80 90 100 110 120130140150 160 Temperature in degrees fahrenheit FIGURE 2-65, Typical cable rigging chart.