Manual control torque tube Landing gear motor Manual control gearbox • Drag strut Gearbox Trunnion support Retracting mechanism Shock strut
• Drag strut Universal joints FIGURE 9-13. Electrical retraction system.
Electrical Landing Gear Retraction System An electrical landing gear retraction system, such as that shown in figure 9-13, has the following features:
(1) A motor for converting electrical energy into rotary motion.
(2) A gear reduction system for decreasing the speed and increasing the force of rotation.
(3) Other gears for changing rotary motion (at a reduced speed) into push-pull move-
(4) Linkage for connecting the push-pull movement to the landing gear shock Basically, the system is an electrically driven jack for raising or lowering the gear. When a switch in the cockpit is moved to the "up" position, the electric motor operates. Through a system of shafts, gears, adapters, an actuator screw, and a torque tube, a force is transmitted to the drag strut linkages. Thus, the gear retracts and locks. If the switch is moved to the "down" position, the motor reverses and the gear moves down and locks. The sequence of operation of doors and gears is similar to that of a hydraulically operated landing gear system.
Hydraulic Landing Gear Retraction Systems Devices used in a typical hydraulically operated landing gear retraction system include actuating cylinders, selector valves, uplocks, downlocks, sequence valves, tubing, and other conventional hydraulic components. These units are so interconnected that they permit properly sequenced retraction and extension of the landing gear and the landing gear doors.
The operation of a hydraulic landing gear retraction system is of such importance that is must be covered in some detail, First, consider what happens when the landing gear is retracted. As the selector valve (figure 9-14) is moved to the "up" position, pressurized fluid is directed into the gear up line. The fluid flows to each of eight units; to sequence valves C and D, to the three gear downlocks, to the nose gear cylinder, and to the two main actuating cylinders.
Notice what happens to the fluid flowing to sequence valves C and D in figure 9-14. Since the sequence valves are closed, pressurized fluid cannot flow to the door cylinders at this time. Thus, the doors cannot close. But the fluid entering the three downlock cylinders is not delayed; therefore, the gear is unlocked. At the same time, fluid also enters the up side of each gear-actuating cylinder and the gears begin to retract. The nose gear completes 349
Selector valve From main hydraulic Down system pressure manifold
• To return manifold check valve Orifice Sequence valves Orifice check valve Gear cylinder Gear cylinder Up Up Up-lock Up-lock Close Down-lock Down-lock Door cylinde Door cylinder Sequence valves D Up-lock Nose gear Down-lock cylinder Code Pressure manifold Return manifold Up gear up, door close line Up Gear down, door open line FIGURE 9-14. Hydraulic landing gear retraction system schematic.
retraction and engages its uplock first, because of the small size of its actuating cylinder. Also, since the nose gear door is operated solely by linkage from the nose gear, this door closes. Meanwhile, the main landing gear is still retracting, forcing fluid to leave the downside of each main gear cylinder.
This fluid flows unrestricted through an orifice check valve, opens the sequence check valve A or B, and flows through the landing-gear selector valve into the hydraulic system return line. Then, as the main gear reaches the fully retracted position and engages the spring-loaded uplocks, gear linkage strikes the plungers of sequence valves C and D.
This opens the sequence check valves and allows pressurized fuid to flow into the door cylinders, closing the landing gear doors.
accomplished using compressed air which is directed to uplock release cylinders.
In some aircraft, design configurations make emergency extension of the landing gear by gravity and airloads alone impossible or impractical. In such aircraft, provisions are included for forceful gear extension in an emergency. Some installations are designed so that either hydraulic fluid, or compressed air provides the necessary pressure; while others use a manual system for extending the landing gears under emergency conditions.
Hydraulic pressure for emergency operation of the landing gear may be provided by an auxiliary hand pump, an accumulator, or an electrically powered hydraulic pump, depending upon the design of the aircraft.
LANDING GEAR SAFETY DEVICES Wing Landing Gear Operation A typical wing landing gear operating sequence is illustrated in figure 9-15. The wing landing gear retracts or extends when hydraulic pressure is applied to the up or down side of the gear actuator.
The gear actuator applies the force required to raise and lower the gear. The actuator works in conjunction with a walking beam to apply force to the wing gear shock strut, swinging it inboard and forward into the wheel well. Both the actuator and the walking beam are connected to lugs on the landing gear trunnion, The outboard ends of the actuator and walking beam pivot on a beam hanger which is attached to the aircraft structure. A wing landing gear locking mechanism located on the outboard side of the wheel well locks the gear in the "up" position. Locking of the gear in the "down" position is accomplished by a downlock bungee which positions an upper and lower jury strut so that the upper and lower side struts will not fold.
Accidental retraction of a landing gear may be prevented by such safety devices as mechanical downlocks, safety switches, and ground locks. Mechanical downlocks are built-in parts of a gear-retraction system and are operated automatically by the gear-retraction system. To prevent accidental operation of the downlocks, electrically operated safety switches are installed.
Safety Switch A landing gear safety switch (figure 9-16) in the landing gear safety circuit is usually mounted in a bracket on one of the main gear shock struts. This switch is actuated by a linkage through the landing gear torque links. The torque links spread apart or move together as the shock strut piston extends or retracts in its cylinder. When the strut is compressed (aircraft on the ground), the torque links are close together, causing the adjusting links to open the safety switch, During takeoff, as the weight of the aircraft leaves the struts, the struts and torque links extend, causing the adjusting links to close the safety switch. As shown in figure 9-16, a ground
EMERGENCY EXTENSION SYSTEMS is completed when the safety switch closes. The ing gears if the main power system fails. Some aircraft have an emergency release handle in the cockpit, which is connected through a mechanical solenoid then energizes and unlocks the selector valve so that the gear handle can be positioned to raise the gear.
Ground Locks linkage to the gear uplocks. When the handle is In addition to this safety device, most aircraft are operated, it releases the uplocks and allows the equipped with additional safety devices to prevent gears to free-tall, or extend, under their own collapse of the gear when the aircraft is on the weight. On other aircraft, release of the uplock is ground. These devices are called ground locks. One 351
Up and locked Inboard In transit Downlock bungee Gear actuator
- Jury strut (upper and lower) Upper side Downlock actuator strut Drag brace Lower (ommitted side strut for clarity) Down and locked FIcURE 9-15. Wing landing gear operating sequence, 352
Control handle - Warning light (off) - Landing Lock release solenoid selector gear UP UP UP 28V valve de bus Spring - bar Tab type indicators (A) Landing Gear "Up and Locked" Safety switch Lock-pin FICURE 9-16. Typical landing gear safety circuit. Warning light (on) - common type is a pin installed in aligned holes drilled in two or more units of the landing gear support structure. Another type is a spring-loaded clip designed to fit around and hold two or more units of the support structure together. All types of ground locks usually have red streamers permanently attached to them to readily indicate whether or not they are installed.
Goar Indicators To provide a visual indication of landing gear position, indicators are installed in the cockpit or flight compartment.
Gear warning devices are incorporated on all retractable gear aircraft and usually consist of a horn or some other aural device and a red warning light.
The horn blows and the light comes on when one or more throttles are retarded and the landing gear is in any position other than down and locked.
Several designs of gear position indicators are available. One type displays movable miniature landing gears which are electrically positioned by movement of the aircraft gear. Another type consists of two or three green lights which burn when the aircraft gear is down and locked. A third type (figure 9-17) consists of tab-type indicators with markings "up" to indicate that the gear is up and locked, a display of red and white diagonal stripes to show when the gear is unlocked, or a silhouette of each gear to indicate when it locks in the "down" position.
Nosewheel Centering Centering devices include such units as internal Diagonal red and fluorescent stripes (B) Gear unlocked and in an intermediate position Warning light (off) → Silhouette of wheels (C) Landing gear "Down and Locked" FIcURE 9-17. A typical gear position indicator and warning light.
centering cams (figure 9-18) to center the nose wheel as it retracts into the wheel well. If a centering unit were not included in the system, the fuselage wheel well and nearby units could be damaged.
During retraction of the nose gear, the weight of the aircraft is not supported by the strut. The strut is extended by means of gravity and air pressure within the strut. As the strut extends, the raised area of the piston strut contacts the slopping area of the fixed centering cam and slides along it. In so doing, it aligns itself with the centering cam and rotates the nose gear piston into a straight-ahead direction.
The internal centering cam is a feature common to most large aircraft. However, other centering 353
Heavy Aircraft Rod connected Large aircraft, with their larger mass and a need to cylinder strut for positive control, utilize a power source for nose- Cylinder wheel steering. Even though large aircraft nosewheel steering system units differ in their construction features, basically all of these systems work in approximately the same manner and require the same sort of units. For example, each steering sys.
tem (figure 9-19) usually contains:
(1) A cockpit control, such as a wheel, handle, lever, or switch to allow starting, stopping, and to control the action of the system.
(2) Mechanical, electrical, or hydraulic connections for transmitting cockpit control movements to a steering control unit.
Centering cam Lobed area Shock strut piston FIcURE 9-18. Cutaway view of a nose gear internal centering cam.
devices are commonly found on small aircraft.
Small aircraft characteristically incorporate an external roller or guide pin on the strut. As the strut is folded into the wheel well on retraction, the roller or guide pin will engage a ramp or track mounted to the wheel well structure. The ramp/track guides the roller or pin in such a manner that the nose wheel is straightened as it enters its well.
In either the internal cam or external track arrangement, once the gear is extended and the weight of the aircraft is on the strut, the nose wheel may be turned for steering.
NOSEWHEEL STEERING SYSTEM Light Aircraft Light aircraft are commonly provided nosewheel steering capabilities through a simple system of mechanical linkage hooked to the rudder pedals. Most common applications utilize push-pull rods to connect the pedals to horns located on the pivotal portion of the nosewheel strut.
354:
(3) A control unit, which is usually a metering or control valve.
(4) A source of power, which is, in most instances, the aircraft hydraulic system.
(5) Tubing for carrying fluid to and from various parts of the system.
(6) One or more steering cylinders, together with the required linkages, for using pressurized fluid to turn the nose gear.
(7) A pressurizing assembly to keep fluid in each steering cylinder always under pressure, thereby preventing shimmy.
(8) A followup mechanism, consisting of gears, cables, rods, drums, and/or bellcranks, for returning the steering control unit to NEUTRAL, and thus holding the nose gear at the correct angle of turn.
(9) Safety valves to allow the wheels to trail or swivel, in the event of hydraulic failure.
Steering emergency forces the piston to start extending. Since the rod of release switch this piston connects to the nose steering spindle, which pivots at point X, the extension of the piston Steering wheel turns the steering spindle gradually toward the Metering valve Steering right. This action turns the nose gear slowly toward drum Differential link the right, because the spindle is connected to the
• Pedestal nose gear shock strut. As the nose gear turns right, Differential arm fluid is forced out of cylinder B, through the leftsompensator turn alternating line, and into port B of the meter- Differential • Pulleys ing valve. The metering valve sends this return fluid assembly Steering cables into a compensator, which routes the fluid into the aircraft system return manifold.
Followup cables Thus, hydraulic pressure starts the nose gear Followup drum Pulleys turning. However, the gear should not be turned too far. The nose gear steering system contains Safety shutoff valve - Scissor followup links and hold it there. devices to stop the gear at the selected angle of turn Orifice rod Centering cams Steering " cylinder "A" Steering spindle Steering cylinder "B Point X Gear yoke FICURE 9-19. Nosewheel system mechanical and hydraulic units.
Nosewheel Steering Operation The nosewheel steering wheel connects through a shaft to a steering drum located inside the cockpit control pedestal. The rotation of this drum transmits the steering signal by means of cables and pulleys to the control drum of the differential assembly. Movement of the differential assembly is transmitted by the differential link to the metering valve assembly, where it moves the selector valve to the selected position. Then hydraulic pressure provides the power for turning the nose gear.
As shown in figure 9-20, pressure from the aircraft hydraulic system is directed through the opened safety shutoff valve and into a line leading to the metering valve. This metering valve then routes pressurized fluid out of port A, through the right-turn alternating line, and into steering cylinder A. This is a one-port cylinder, and pressure Followup Linkage As already pointed out, the nose gear is turned by the steering spindle as the piston of cylinder A (figure 9-20) extends. But the rear of the spindle has gear teeth which mesh with a gear on the bottom of the orifice rod. Thus, as the nose gear and its spindle turn, the orifice rod also turns, although in the opposite direction. This rotation is transmitted by the two sections of the orifice rod to the scissor followup links (figure 9-19), located at the top of the nose gear strut. As the followup links return, they rotate the connected followup drum, which transmits the movement by cables and pulleys to the differential assembly. Operation of the differential assembly causes the differential arm and links to move the metering valve back toward the neutral position.
The compensator unit (figure 9-21) which is a part of the nosewheel system keeps fluid in the steering cylinders pressurized at all times. This hydraulic unit consists of a three-port housing, which encloses a spring loaded piston and poppet. The left port is an air vent, which prevents trapped air at the rear of the piston from interfering with movements of the piston. The second port, located at the top of the compensator, connects through a line to the metering valve return port. The third port is located at the right side of the compensator. This port, which is connected to the hydraulic return 355
Orifice rod - From hydraulic system Safety shutoff valve pressure manifold Nose steering Steering cylinder "B" Poppet- - Spring spindle Emergency bypass valve Metering Left turn line valve Return port To hydraulic system return Point X manifold Right turn line Compensator Port "A Port 'B" Steering cylinder "A" za Pressurized fluid Return fluid FICURE 9-20. Nosewheel steering hydraulic flow diagram.
manifold, routes the steering system return fluid into the manifold when the poppet valve is open.
The compensator poppet opens when pressure acting on the piston becomes high enough to compress the spring. This requires 100 p.s.i.; therefore, fluid in the metering valve return line contains fluid trapped under that pressure. Now, since pressure in a confined fluid is transmitted equally and undiminished in all directions (Pascal's law), 100 p.s.i. also exists in metering valve passage H and in chambers From pressure E, D, G, and F (figure 9-21). This same pressure is also applied in the right- and left-turn alternating lines as well as in the steering cylinders themselves.
SHIMMY DAMPERS A shimmy damper controls vibration, or shimmy, through hydraulic damping. The damper is either attached to or built integrally with the nose gear and prevents shimmy of the nosewheel during taxiing, landing, or takeoff. There are three types of shimmy dampers commonly used on aircraft: (1) The piston type, (2) vane type, and (3) features incorporated in the nosewheel power steering system of some aircraft.
manifold Piston-Type Shimmy Damper Drilled passage "H" G Return port 9-22 consists of two major components: (1) The The piston-type shimmy damper shown in figure cylinder "A" From = cylinder From Metering valve der. lower end of the nose gear shock strut outer cylin- cam assembly and (2) the damper assembly. The shimmy damper is mounted on a bracket at the "B" The cam assembly is attached to the inner cylin- Spring Popper der of the shock strut and rotates with the nose- wheel. Actually, the cam consists of two cams that Air • are mirror images of each other. Lobes on the cams vent Compensator are so placed that the damping effect offers the Housing retum manifold To hydraulic centered. The cam follower crank is a U-shaped greatest resistance to rotation when the wheel is casting which incorporates the rollers that follow FICURE 9-21. Cutaway view of metering valve and the cam lobes to restrict rotation. The arm of the compensator. crank connects to the operating piston shaft.
Reservoir Reservoir orifice Check ball Nose wheel straight ahead Nose wheel castering Shimmy damper operational schematic Cam follower Shimmy Vent hole damper cam Shimmy Shimmy damper Red line on indicator rod Filler assembly Cam follower Cam follower crank crank FIcuRE 9-22. Typical piston-type shimmy damper.
The damper assembly consists of a spring-loaded reservoir piston to maintain the confined fluid under constant pressure, and an operating cylinder and piston. A ball check permits the flow of fluid from the reservoir to the operating cylinder to make up for any fluid loss in the operating cylinder.
Because of the rod on the operating piston, its stroke away from the filler end of the piston displaces more fluid than its stroke toward the filler end. This differeuce is taken care of by the reser- 357
voir orifice, which permits a small flow both ways between the reservoir and operating cylinder.
A red mark (ngure 9-22) on the reservoir indicator rod indicates the fluid level in the reservoir.
When the piston goes into the reservoir so far that the mark is not visible, the reservoir must be serviced. The operating cylinder houses the operating piston. A small orifice in the piston head allows fluid to flow from one side of the piston to the other. The piston shaft is connected to the arm of the cam follower crank.
As the nosewheel fork rotates in either direction (figure 9-22), the shimmy damper cam displaces the cam follower rollers, causing the operating piston to move in its chamber. This movement forces fluid through the orifice in the piston. Since the orifice is very small, rapid movements of the piston, which commonly occur during landing and takeof, are restricted, and nosewheel shimmy is eliminated.
Gradual rotation of the nosewheel fork is not resisted by the damper. This enables the aircraft to be steered at slow speeds. If the fork rotates in either direction until the rollers are over the high spots on their portions of the cam, further movement of the nosewheel is practically unrestricted.
The piston-type shimmy damper generally requires a minimum of servicing and maintenance;
however, it should be checked periodically for evidence of hydraulic leaks around the damper asserbly, and the reservoir fluid level must be properly maintained at all times. The cam assembly should be checked for evidence of binding and for worn, loose, or broken parts.
Vane-Type Shimmy Damper The vane-type shimmy damper is located on the nosewheel shock strut just above the nosewheel fork and may be mounted either internally or externally.
If mounted internally, the housing of the shimmy damper is fitted and secured inside the shock strut, and the shaft is splined to the nosewheel fork. If mounted externally, the housing of the shimmy damper is bolted to the side of the shock strut, and the shaft is connected by mechanical linkage to the nosewheel fork.
The housing (figure 9-23) is divided into three
- Indicator rod connected to Abutment and replenishing piston valve assembly _Spring Replenishing check valve Replenishing chamber - Replenishing piston Valve orifice Rotating vane Filler cap Abutment flange Rotating vane. Key
- Fluid port Fluid port Dowel pin Hollow dowel pin Wing shaft
- Hydraulic seal Closing flange Spring.
retainer Packing spring Mounting Wing shaft packing flange Wing shaft FIcURE 9-23. Typical vane-type shimmy damper.
main parts: (1) The replenishing chamber, (2) the working chamber, and (3) the lower shaft packing chamber.
The replenishing chamber is in the top part of the housing and stores a supply of fluid under pressure, Pressure is applied to the fluid by the spring-loaded replenishing piston and piston shaft which extends through the upper housing and serves as a fluid level indicator. The area above the piston contains the piston spring and is open to atmosphere to prevent a hydraulic lock. Fluid is prevented from leaking past the piston by O-ring packings. A grease fitting provides the means for filling the replenishing chamber with fluid.
The working chamber is separated from the replenishing chamber by the abutment and valve assembly. The working chamber contains two one-way ball check valves, which will allow fluid to flow from the replenishing chamber to the working chamber only. This chamber is divided into four sections by two stationary vanes called abutment flanges, which are keyed to the inner wall of the housing, and two rotating vanes, which are an integral part of the wing shaft. The shaft contains the valve orifice through which the fluid must pass in going from one chamber to another.
Turning the nosewheel in either direction causes the rotating vanes to move in the housing. This results in two sections of the working chamber growing smaller, while the opposite two chambers grow larger. The rotating vane can move only as fast as the fluid can be displaced from one chamber to the other. All of the fluid being displaced must pass through the valve orifice in the shaft. Resist.
ance to the flow of fluid through the orifice is proportional to the velocity of flow. This means that the shimmy damper offers little resistance to slow motion, such as that encountered during normal steering of the nose gear or ground handling, but offers high resistance to shimmy on landing, takeoff, and high-speed taxiing. An automatic orifice adjustment compensates for temperature changes. A bimetallic thermostat in the shaft opens and closes the orifice as the temperature and viscosity change.
This results in a constant resistance over a wide temperature range, In case an exceptionally high pressure is suddenly built up in the working chamber by a severe twisting force on the nosewheel, the closing flange moves down, compressing the lower shaft packing spring, allowing fluid to pass around the lower ends of the vanes, preventing structure damage.
Maintaining the proper fluid level is necessary to the continued functioning of a vane-type shimmy damper. If a shimmy damper is not operating properly, the fluid level is the first item which should be checked by measuring the protrusion of the indicator rod from the center of the reservoir cover.
Inspection of a shimmy damper should include a check for evidence of leakage and a complete examination of all fittings and connections between the moving parts of the shock strut and the damper shaft for loose connections.
Fluid should be added only when the indicator rod protrudes less than the required amount. The distance the rod should protrude varies with different models. A shimmy damper should not be overfilled. If the indicator rod is above the height speci fied on the nameplate, fluid should be bled out of the damper.
Steer Damper A steer damper is hydraulically operaied and accomplishes the two separate functions of steering and/or eliminating shimmying. The type discussed here is designed for installation on nose gear strute and is connected into the aircraft hydraulic system.
A typical steer damper is shown in figure 9-24.
FIGURE 9-24. Typical steer damper.
Basically, a steer damper consists of a closed cylinder containing rotary vane-type working chambers (similar to the vane-type shimmy damper) and a valving system.
Steer dampers may contain any even number of 359
working chambers. A steer damper with one vane on the wing shaft and one abutment leg on the abutment flange would have two chambers. Similarly, a unit with two vanes on the wingshaft and two abutment legs on the abutment flange would have four chambers. The single- or double-vaned units are the ones most commonly used.
A mechanical linkage is connected from the protruding splined portion of the wing shaft to the wheel fork and is used as a means for transmitting force. The linkage on the steer damper may be connected to a heavy coil spring on the outside of the reservoir for automatic wheel centering.
The steer damper accomplishes two separate functions: one is steering the nosewheel and the other is shimmy damping. Only the damping function of the steer damper is discussed in this section.
The steer damper automatically reverts to damping when, for any reason, the flow of high-pressure fluid is removed from the inlet of the steer damper.
This high pressure, which activates the steer damper valving system, is removed from the control passages by one of two methods, depending on the installation. When the inlet line is supplied with a three-way solenoid valve and the high-pressure supply is shut off, the fluid bleeds out of the unit through the outlet port of the valve to the discharge line. When a two-way solenoid valve is furnished, high-pressure fluid leaves the control passages through an orifice specially provided for this type of installation which is located in the center of the return line plunger.
Effective damping is assured by maintaining unaerated hydraulic fluid in the working chambers.
This is accomplished by allowing air and a very small amount of hydraulic fluid to leave the working chambers through strategically located vent grooves while unaerated fluid is allowed to enter through replenishing valves from the hydraulic return line. Excessive pressure in the unit due to temperature changes is prevented by the thermal relief valve in the inlet flange.
Daily inspection of a steer damper should include a check for leakage and a complete inspection of all hydraulic connections, steer damper mounting bolts for tightness, and all fittings and connections between the moving parts of the shock strut and the steer damper wing shaft.
BRAKE SYSTEMS Proper functioning of the brakes is of utmost importance on aircraft. The brakes are used for slowing, stopping, holding, or steering the aircraft.
They must develop sufficient force to stop the aircraft in a reasonable distance; brakes must hold the aircraft for normal engine turnup; and brakes must permit steering of the aircraft on the ground.
Brakes are installed in each main landing wheel and they may be actuated independently of each other.
The right-hand landing wheel is controlled by applying toe pressure to the right rudder pedal and the left-hand wheel is controlled by the left rudder pedal.
For the brakes to function efficiently, each component in the brake system must operate properly, and each brake assembly on the aircraft must operate with equal effectiveness. It is therefore important that the entire brake system be inspected frequently and an ample supply of hydraulic fluid be maintained in the system. Each brake assembly must be adjusted properly and friction surfaces kept free of grease and oil.
Three types of brake systems are in general use:
(1) Independent systems, (2) power control systems, and (3) power boost systems. In addition, there are several different types of brake assemblies in widespread use.
Independent Brake Systems In general, the independent brake system is used on small aircraft. This type of brake system is termed "independent" because it has its own reservoir and is entirely independent of the aircraft's main hydraulic system.
Independent brake systems are powered by master cylinders similar to those used in the conventional automobile brake system. The system is composed of a reservoir, one or two master cylinders, Brake pedals Reservoir Master cylinders Parking brake valves (operated by a single control handle) L.H. brake assembly R.H. brake assembly, FIGURE 9-25. Typical independent brake system.