{[ promptMessage ]}

Bookmark it

{[ promptMessage ]}

36002209-8-Landing-Gear - ATA-32 LANDING GEAR The landing...

Info iconThis preview shows page 1. Sign up to view the full content.

View Full Document Right Arrow Icon
This is the end of the preview. Sign up to access the rest of the document.

Unformatted text preview: ATA-32 LANDING GEAR The landing gear can extended or retracted by hydraulically or pneumatically or electrically. The purpose is to reduce the drag on the aircraft during flight. Landing gear Position Indication: An electrical indicating system is used to provide a positive indication to the crew of the operation of the locks and of the position of the landing gear. The system usually consists of microswitches on the up-locks and down-locks, which make or break when the locks operate, and which are connected to a landing gear position indicator on the instrument panel. On British manufactured aircraft, the electrical undercarriage indicating system operates in such a manner that a green light is displayed when the undercarriage is locked down. A red light is displayed when the undercarriage is in transit. No lights are visible when the undercarriage is locked up; Bulbs are usually duplicated to avoid the possibility of false indications as a result of bulb failures. On other aircraft, similar indications may be obtained by the use of magnetic indicators or lights, but on some light aircraft a single green light indicates that all undercarriages are locked down, and an amber light indicates that all undercarriages are locked up. Whichever indicating system is used, it is important that the microswitches are adjusted so that operation of the lights coincides with the corresponding position of the landing gear. Warning Devices: To guard against landing with the landing gear retracted or unlocked, a warning horn is incorporated in the system and connected to a throttle-operated switch. Current for the green indicator light flows through the 5-amp circuit breaker, wire 6, the nose-gear down switch, wire 5. The left gear-down switch, wire 4, the right gear-down switch, and wire 3 to the green light, causing it to illuminate. When the pilot moves the landing-year selector handle into the GEAR-UP position, its switch changes position, and the circuit is completed from the 20-amp circuit breaker through wires I and 13 to the right-hand terminal of the landing-gear relay. Current flows from this connection on the relay through wire 10, through the upper contacts of the up-limit switch, through the gear-safety switch, and through wire 12, lo the coil of the landing-gear relay. This current produces a magnetic pull, which closes the relay so that current can flow through the relay contacts and the winding of the reversible DC motor to raise the landing gear. As soon as the landing gear is released from its downlocks, the three landing-gear-down switches open and the green light goes out. The landing gear has not reached its up-and-locked position, so the red light is off. When landing gear is retracted and locked current flows through the lower contacts of the up-limit switch, through wires 19 and 8, to the red light, showing that the landing gear is up and locked. If either throttle is closed when the landing gear is not down and locked, current will flow through the nose-gear-clown switch, the throttle switch, and the down-limit switch and will sound the gear-warning horn. This horn warns the pilot that the landing gear has not been lowered in preparation for landing. When the landing-gear-selector switch is moved into the GEAR-DOWN position, current flows through it and the down-limit switch to the gear-down side of the reversible DC motor that lowers the landing gear. As soon as the gear is down and locked, the down-limit switch opens and shuts off the landing-gear motor. The three gear-down switches close and the green light comes on. Squat switch (shock strut micro switch prevents the landing gear being operated when the aircraft is on the ground this switch operates anti-skid system when the weight is on the strut). The motor is of the series-wound split-field type which is mechanically coupled to the three "leg" units, usually by a gearbox, torque shafts, cables, and screw jacks. The 28 volts d.c. supply to the motor is controlled by a selector switch, relay, and switches in the "down-lock" and "uplock" circuits. A safety switch is also included in the circuit to prevent accidental retraction of the gear while the aircraft is on the ground. The switch is fitted to the shock-strut of one of the main wheel gear units, such that the compression of the strut keeps the switch contacts in the open position as shown in the diagram. After take-off, the weight of the aircraft comes off the landing gear shock-struts, and because they have a limited amount of telescopic movement, the strut controlling the safety switch causes it to close the switch contacts. Thus, when the pilot selects "gear up", a circuit is completed via the selector switch, and closed contacts of the up-lock switch, to the coil of the relay which then completes the supply circuit to the "up" winding of the motor. When the landing gear units commence retracting, the down-lock switch is automatically actuated such that its contacts will also close, and will remain so up to and in the fully retracted position. As soon as this position is reached, the up-lock switch is also actuated so as to open its contacts, thereby interrupting the supply to the motor, and the "down" winding circuit of the motor is held in readiness for extending the landing gear. As and when the appropriate selection is made, and the landing gear units commence extending, the up-lock switch contacts now close and when the landing gear is down and locked, and the aircraft has landed, the circuit is again restored to the condition. To prevent over-run of the motor, and hence over-travel of the landing gear units, some form of braking is necessary. This is accomplished in some cases, by incorporating a dynamic brake relay in the circuit. The relay operates in such a manner that during over-run, the motor is caused to function as a generator, the resulting electrical load on the armature stopping the motor and gear instantly. Landing gear position indication: The system operates from a 28 volts d.c. power supply which is connected to lamps within the indicator case, and also to the up-lock and down-lock micro-switches of the main and nose landing gear units. Three of the lamps are positioned behind red screens, and three behind green screens; thus, when illuminated they indicate respectively, "gear up and locked" and "gear down and locked", hi the "gear up and locked" position all lights are extinguished. In the event of failure of a green lamp filament, provision is made for switching-in a standby set of lamps. The circuit as drawn represents the conditions when the aircraft is on the ground in a completely static condition. As soon as power goes onto the busbar, the three green tamps will illuminate because their circuits are completed to ground via the left-hand set of contacts of the corresponding down-lock micro-switches. The engine throttle is closed, and although its microswitch is also closed, the warning horn circuit is isolated since there is no path to ground for current from the busbar. Assume now that the aircraft has taken off and the pilot has selected "landing gear up"; the down-lock mechanisms of the gear units are disengaged and they cause their micros witches to change contact positions, thus interrupting the circuits to the green lamps. At the same time, the red lamps are illuminated to indicate that the gear units are unlocked, the power supply for the circuit passing to ground via the up-lock switches, and the right-hand contacts of the down-lock switches. When the landing gear units reach their retracted positions, the up-lock mechanisms are engaged and cause their microswitches to interrupt the circuits to the red lamps; thus, all lamps are extinguished. When the pilot selects "landing gear down", the up-lock mechanisms now disengage and the micro-switches again complete the circuit to the red lamps to indicate an unlocked condition. As soon as the gear units reach the fully extended position, the down-lock mechanisms engage and their microswitches revert to the original position shown in Fig. 10.32 i.e., red lamps extinguished, and green lamps illuminated to indicate "down and locked". As noted earlier, a warning horn is included in the system, the making and breaking of the horn circuit being controlled by a throttle-operated microswitch. In the static condition, the throttle microswitch is closed, but the warning horn will not sound since the circuit is interrupted by all three down-lock microswitches. Similarly, the circuit will be interrupted by the throttle microswitch which is opened when the throttle is set for take-off and normal cruise power. In the case of an approach 10 land, the engine power is reduced by closing the throttle to a particular approach power setting and this action closes the throttle microswitch. If, in this flight condition, the landing gear has not been selected down in readiness for landing, then the warning horn will sound since the circuit to ground is then completed via the right-hand contacts of the down-lock microswitches. After selecting "down", the horn continues to sound, but it may be silenced by operating a push switch which, as will be noted from the diagram, energizes a relay to interrupt the horn circuit. The relay incorporates a hold-in circuit so that it will remain energized until the d.c. power supply is finally switched off. Functional testing of the horn circuit on the ground, and under engine static conditions, may be carried out by closing the throttle and its microswitch, and then operating a test switch. Anti skid control system: Anti skid system is provided on large transport a/c to prevent the main L/G wheels from skidding on wet or icy surface & for ensuring an effective breaking system by modulating the hydraulic supply's pressure to the breaks. fundamentally the anti skid system senses the rate of change of wheel deceleration, decreases the hydraulic pressure applied to the brakes when there is impending skid condition, but restores, the pressure as the wheel accelerates. The system consists of wheels speed sensing transducer one each for main wheels. An antiskid control module consisting of individual circuitry & electro hydraulic antiskid valves as many as there are transducer & a control cct. The transducer is a speed sensing device & consists of a stator which is attached to the wheel axel & a rotor which is attached to & rotates with the wheel. The stator consists of a permanent magnet & when the wheel & rotor are rotated on AC voltage is generated which is directly proportional to rotational speed of the wheel axel & a rotor which is attached to & rotates with the wheel. The stator consists of a permanent magnet and when the wheel & rotor are rotated the magnetic coupling or magnetic reluctance between the rotor and stator is varied this variation generates AC voltage in stator. The AC voltage is generated which is directly proportional to rotational speed of the wheel. The signal is fed into the converter of the control unit & is converted into DC which serves as a measure of deceleration of the wheel. The signal is then applied to a skid control cct where it is compared with a reference signal which has been predetermined from a known deceleration rate of A/c. Any difference of between the two signals produces the error voltage whether or not a correction signal is to be applied to the electro hydraulic control value. If wheel deceleration rate are below the reference velocity no correction signal is produced. If however the rates are above the reference velocity they are then treated as skids or approaching skid & correction signals are equipped to control valve which reduces the hydraulic pressure applied to the brakes. When the wheel speed falls below the references deceleration rate the skid control unit transmits a release signal to the control valves subsequent wheel spin up causes the brakes to reapply. In locked wheel condition the cct will cause signals to be applied to relevant control valve such that they will fully release pressure. When the speed of the wheels less than 20 miles per hour the anti-skid system does not function. ...
View Full Document

{[ snackBarMessage ]}

Ask a homework question - tutors are online