A. Impact break B. Liner breakdown C. Contamination A. Brake heat damage B. Chaff damage FIGURE 65--Cont.
ANTISKID SYSTEM The purpose of a wheel brake is to bring a rapidly moving aircraft to a stop during ground roll.
It does this by changing the energy of movement into heat energy through the friction developed in the brakes. A feature found in high performance aircraft braking systems is skid control or antiskid protection. This is an important system because if a wheel goes into a skid, its braking value is greatly reduced.
The skid control system performs four functions:
(1) normal skid control, (2) locked wheel skid control, (3) touchdown protection, and (4) failsafe protection. The main components of the system consist of two skid control generators, a skid control box, two skid control valves, a skid control switch, a warning lamp, and an electrical control harness with a connection to the squat switch.
Normal Skid Control Normal skid control comes into play when wheel rotation slows down but has not come to a stop.
When this slowing down happens, the wheel sliding action has just begun but has not yet reached a full scale slide. In this situation the skid control valve removes some of the hydraulic pressure to the wheel. This permits the wheel to rotate a little faster and stop its sliding. The more intense the skid is, the more braking pressure is removed.
The skid detection and control of each wheel is completely independent of the others. The wheel 399
skid intensity is measured by the amount of wheel Skid Control Generator The skid control generator is the unit that measures the wheel rotational speed. It also senses any changes in the speed. It is a small electrical generator, one for each wheel, mounted in the wheel axle. The generator armature is coupled to, and driven by, the main wheel through the drive cap in the wheel. As it rotates, the generator develops a voltage and current signal. The signal strength indicates the wheel rotational speed. This signal is fed to the skid control box through the Skid Control Box The box reads the signal from the generator and senses change in signal strength. It can interpret these as developing skids, locked wheels, brake applications, and brake releases. It analyses all it reads, then sends appropriate signals to solenoids in the skid control valves.
Skid Control Valves The two skid control valves mounted on the brake control valve are solenoid operated. Electric signals from the skid control box actuate the solenoids. If there is no signal (because there is no wheel skidding), the skid control valve will have no effect on brake operation. But, if a skid develops, either slight or serious, a signal is sent to the skid control valve solenoid. This solenoids' action lowers the metered pressure in the line between the metering valve and the brake cylinders.
It does so by dumping fuid into the reservoir return line whenever the solenoid is energized.
Naturally, this immediately relaxes the brake application. The pressure flow into the brake lines from the metering valves continues as long as the pilot depresses the brake pedals. But the flow and pressure is rerouted to the reservoir instead of to the wheel brakes.
The utility system pressure enters the brake control valve where it is metered to the wheel brakes in proportion to the force applied on the pilot's foot pedal. However, before it can go to the brakes, it must pass through a skid control valve.
There, if the solenoid is actuated, a port is opened in the line between the brake control valve and the brake. This port vents the brake application pressure to the utility system return line. This reduces the brake application, and the wheel rotates faster again. The system is designed to apply enough force to operate just below the skid point. This gives the most effective braking.
Pilot Control The pilot can turn off the operation of the antiskid system by a switch in the cockpit. A warning lamp lights when the system is turned off or if there is a system failure.
Locked Wheel Skid Control The locked wheel skid control causes the brake to be fully released when its wheel locks. A locked wheel casily occurs on a patch of ice due to lack of tire friction with the surface. It will occur if the normal skid control does not prevent the wheel from reaching a full skid. To relieve a locked wheel skid, the pressure is bled off longer than in normal skid function. This is to give the wheel time to regain speed. The locked wheel skid control is out of action during aircraft speeds of less than 15-20 mph.
Touchdown Protection The touchdown protection circuit prevents the brakes from being applied during the landing approach even if the brake pedals are depressed.
This prevents the wheels from heing locked when they contact the runway. The wheels have a chance to begin rotating before they carry the full weight of the aircraft. Two conditions must exist before the skid control valves permit brake application. Without them the skid control box will not send the proper signal to the valve solenoids.
The first is that the squat switch must signal that the weight of the aircraft is on the wheels. The second is that the wheel generators sense a wheel speed of over 15-20 mph.
Fail-Safe Protection The fail-safe protection circuit monitors operation of the skid control system. It automatically returns the brake system to full manual in case of system failure. It also turns on a warning light.
LANDING GEAR SYSTEM MAINTENANCE Because of the stresses and pressures acting on the landing gear, inspection, servicing, and other maintenance becomes a continuous process. The most important job in the maintenance of the aircraft landing gear system is thorough, accurate inspections. To properly perform the inspections, all surfaces should be cleaned to ensure that no trouble spots go undetected.
Periodically, it will be necessary to inspect shock struts, shimmy dampers, wheels, wheel bearings, tires, and brakes. During this inspection, check for the presence of installed ground safety locks. Check landing gear position indicators, lights, and warn- 400
ing horns for operation. Check emergency control handles and systems for proper position and condition. Inspect landing gear wheels for cleanliness, corrosion, and cracks. Check wheel tie bolts for looseness. Examine anti-skid wiring for deterioration. Check tires for wear, cuts, deterioration, presence of grease or oil, alignment of slippage marks, and proper inflation. Inspect the landing gear mechanism for condition, operation, and proper adjustment. Lubricate the landing gear, including the nose wheel steering. Check steering system cables for wear, broken strands, alignment, and safetying. Inspect landing gear shock struts for such conditions as cracks, corrosion, breaks, and security. Where applicable, check the brake clearances.
Various types of lubricants are required to lubricate points of friction and wear on the landing gear. These lubricants are applied by hand, an oil can, or a pressure-type grease gun. Before using the pressure-type grease gun, wipe the lubrication fittings clean of old grease and dust accumulations, Up line Down line because dust and sand mixed with a lubricant produce a very destructive abrasive compound. As each fitting is lubricated, the excess lubricant on the fitting and any that is squeezed out of the assembly should be wiped off. Wipe the piston rods of all exposed actuating cylinders; clean them frequently, particularly prior to operation, to prevent damage to seals and polished surfaces.
Periodically, wheel bearings must be removed, cleaned, inspected, and lubricated. When cleaning a wheel bearing, use a suitable cleaning solvent.
(Leaded gasoline should not be used.) Dry the bearing by directing a blast of dry air between the rollers. Do not direct the air so that it will spin the bearing, as it may fly apart and injure nearby persons. When inspecting the bearing, check for defects that would render it unservicable, such as flaked, cracked, or broken bearing surfaces; roughness due to impact pressure or surface wear; corrosion or pitting of the bearing surfaces; discoloration from excessive heat; cracked or broken cages;
Uplock switch Forward latch mechanism Latch cylinder Latch cylinder Emergency release cable Aft latch Sector mechanism
- Sector Latch hook Door - Latch roller FreuRE 9-66. Main gear door latch mechanisms, 401
or scored or loose bearing cups or cones which would affect proper seating on the axle or wheel. If any of these defects exist replace the bearing with a servicable one. To prevent rust or corrosion, lubricate the bearing immediately after cleaning and inspecting it.
To apply lubricant to a tapered roller bearing, place a small amount of the proper lubricant on the palm of the hand. Grasp the cone of the bearing assembly with the thumb and first two fingers of the other hand, keeping the larger diameter of the bearing next to the palm. Move the bearing assembly acroas the hand toward the thumb, forcing the lubricant into the space between the cone and rollers.
Turn the assembly after each stroke until all openings between the rollers are filled with lubricant.
Remove the excess lubricant from the cone and the outside of the cage.
Landing Gear Rigging and Adjustment Occasionally it becomes necessary to adjust the landing gear switches, doors, linkages, latches, and locks to assure proper operation of the landing gears and doors. When landing gear actuating cyl- Latch cylinder inders are replaced and when length adjustments are made, overtravel must be checked. Overtravel is that action of the cylinder piston beyond the move.
ment necessary for landing gear extension and retraction. The additional action operates the landing gear latch mechanism.
Because of the wide variety of aircraft types and designs, procedures for rigging and adjusting landing gear will vary. Uplock and downlock clearances, linkage adjustments, limit switch adjustments, and other landing gear adjustments vary widely with landing gear design. For this reason, always consult the applicable manufacturers maintenance or service manual before performing any phase of landing gear rigging or installation.
Adjusting Landing Gear Latches The adjustment of latches are of prime concern to the airframe mechanic. A latch is used in landing gear systems to hold a unit in a certain position after the unit has traveled through a part of, or all ot, its cycle. For example, on some aircraft, when the landing gear is retracted, each gear is held in the up position by a latch. The same holds true Latch cylinder Latched roller support Latch hook Latch hook (serrated) (latched position) (latched position) Latch roller Latch roller (on door) (on door) •Elongated holes 1/8 ÷ 3/32 inch Plate (serrated) (A) Latch roller clearance (B) Latch roller support adjustment FIGUR: 9.67. Landing gear door latch installations.
when the landing gear is extended. Latches are also used to hold the landing gear doors in the open and closed positions.
There are many variations in latch design. However, all latches are designed to accomplish the same thing. They must operate automatically, at the proper time, and hold the unit in the desired position. A typical landing gear door latch is described in the following paragraphs.
On this particular aircraft, the landing gear door is held closed by two door latches. As shown in figure 9-66, one is installed near the rear of the door. To have the door locked securely, both locks must grip and hold the door tightly against the aircraft structure. The principal components of each latch mechanism shown in figure 9-66 are a hydraulic latch cylinder, a latch hook, a spring loaded crank-and-lever linkage, and a sector. The latch cylinder is hydraulically connected with the landing gear control system and mechanically connected, through linkage, with the latch hook. When hydraulic pressure is applied, the cylinder operates the linkage to engage (or disengage) the hook with (or from) the latch roller on the door. In the gear-down sequence the hook is disengaged by the spring load on the linkage. In the gear-up sequence, spring action is reversed when the closing door is in contact with the latch hook and the cylinder operates the linkage to engage the hook with the latch roller.
Cables on the landing gear emergency extension system are connected to the sector to permit emergency release of the latch rollers. An uplock switch is installed on, and actuated by, each latch to provide a gear-up indication in the cockpit.
With the gear up and the door latched, inspect the latch roller for proper clearance as shown in figure 9-67, view A. On this installation the required clearance is 1/8 +3/32 in. If the roller is not within tolerance it may be adjusted by loosening its mounting bolts and raising or lowering the latch roller support. This may be done due to the elongated holes and serrated locking surfaces of the latch roller support and serrated plate (view B).
Landing Gear Door Clearances Landing gear doors have specific allowable clearances which must be maintained between doors and the aircraft structure or other landing gear doors.
These required clearances can be maintained by adjusting the door hinges and connecting links and trimming excess material from the door if necessary.
On some installations, door hinges are adjusted by placing the serrated hinge and serrated washers in the proper position and torquing the mounting bolts. Figure 9-68 illustrates this type of mounting.
which allows linear adjustments. The amount of linear adjustment is controlled by the length of the elongated bolthole in the door hinge.
Fairing door fitting Hinge support hinge Bushing.
Attach bolt FICURE 9-68. Adjustable door hinge installation.
The distance the landing gear doors open or close depends upon the length of door linkage and the adjustment of the door stops. The manufacturer's maintenance manuals specify the length of door linkages and adjustment of stops or other procedures whereby correct adjustments may be made.
Landing Gear Drag and Side Brace Adjustment The landing gear side brace illustrated in figure 9-69 consists of an upper and lower link, hinged at the center to permit the brace to jackknife during retraction of the landing gear. The upper end pivots on a trunnion attached to the wheel well overhead.
The lower end is connected to the shock strut.
On the side brace illustrated, a locking link is incorporated between the upper end of the shock strut and the lower drag link. Usually in this type installation, the locking mechanism is adjusted so that it is positioned slightly overcenter. This provides positive locking of the side brace and the locking mechanism, and as an added safety feature, prevents inadvertant gear collapse caused by the side brace folding.
To adjust the overcenter position of the side brace locking link illustrated in figure 9-69, place the landing gear in the down position and adjust the lock link end fitting so that the side brace lock link is held firmly overcenter. Manually break the 403
Push-pull Door arm stop Bellcrank U plock tube Door Gear actuator strut B Bellcrank Torque tube Push-pull tube Landing gear Uplock push-pull tube actuator Side link Bellcrank Side brace lock link B
- Spring scale .22 = .03 inches Overcenter Adjustable link FiccR: 9-69. Landing gear schematic showing overcenter adjustments.
lock link and move the landing gear to a position 5 or 6 in. inboard from the down locked position, then release the gear. The landing gear must freefall and lock down when released from this position.
In addition to adjusting overcenter travel, the down lock spring tension must be checked using a spring scale. The tension should be between 40 to 60 Ibs. for the lock link illustrated. The specific tension and procedure for checking will vary on other aircraft.
Landing Gear Retraction Check There are several occasions when a retraction check should be performed, First, a retraction check should be performed during an annual inspection of the landing gear system. Second, when performing maintenance that might affect the landing gear linkage or adjustment, such as changing an actuator, make a retraction check to see whether everything is connected and adjusted properly. Third, it may be necessary to make a retraction check after a hard or overweight landing has been made which may have damaged the landing gear. Closely inspect the gear for obvious damage and then make the retraction check. And finally, one method of locating malfunctions in the. landing gear system is to perform a gear retraction check.
There are a number of specific inspections to perform when making a tetraction check of landing gear. Included are:
(1) Landing gears for proper retraction and extension.
(2) Switches, lights, and warning horn or buzzer for proper operation.
(3) Landing gear doors for clearance and freedom from binding.
(4) Landing gear linkage for proper operation, adjustment, and general condition.
(5) Latches and locks for proper operation and adjustment.
(6) Alternate extension or retraction systems for proper operation.
(7) Any unusual sounds such as those caused by rubbing, binding, chafing, or vibration.
The procedures and information presented herein were intended to provide familiarization with some of the details involved in landing gear rigging, adjustments, and retraction checks and do not have general application. For exact information regarding a specific aircraft landing gear system, consult the applicable manufacturer's instructions.
CHAPTER 10
FIRE PROTECTION SYSTEMS
GENERAL Because fire is one of the most dangerous threats to an aircraft, the potential fire zones of modern multi-engine aircraft are protected by a fixed fire protection system. A "fire zone" is an area or region of an aircraft designed by the manufacturer to require fire detection and/or fire extinguishing equipment and a high degree of inherent fire resistance. The term "fixed" describes a permanently CO. fire extinguisher.
A complete fire protection system on modern aircraft and on many older model aircraft includes both a fire 'detection and a fire extinguishing system.
To detect fres or overheat conditions, detectors are placed in the various zones to be monitored.
Fires are detected in reciprocating engine aircralt, using one or more of the following:
(1) Overheat detectors.
(2) Rate-of-temperature-rise detectors.
(3) Flame detectors.
(4) Observation by crewmembers.
In addition to these methods, other types of detectors arc used in aircraft fire protection systems, but are seldom used to detect engine fires; for example, smoke detectors are better suited to monitor areas such as baggage compartments, where materials burn slowly, or smolder. Other types of detectors in this category include carbon monoxide detectors and chemical sampling cquipment capable of detecting combustible mixtures that can lead to accumulations of explosive gases.
Detection Methods The following list of detection methods includes those most commonly used in turbine engine aircraft fire protection systems. The complete aircraft fire protection system of most large turbine engine aircraft will incorporate several of these different detection methods.
(1) Rate-of-temperature-rise detectors.
(2) Radiation sensing detectors.
(3) Smoke detectors.
(1) Overheat detectors.
(5) Carbon monoxide detectors.
(6) Combustible mixture detectors.
(7) Fiber-optic detectors.
(8) Observation of crew or passengers.
The three types of detectors most commonly used for fast detection of fires are the rate-of-rise.
radiation sensing, and overheat detectors.
Detection System Requirements Fire protection systems on modern aircraft do not rely on observation by crewmembers as a primary method of fire detection. An ideal fire detection system will includo as many as possible of the following teatures:
(1) A system which will not cause false warnings under any flight or ground operating conditions.
(2) Rapid indication of a fire and accurate location of the fire.
(3) Accurate indication that a fire is out.
(4) Indication that a fire has re-ignited.
(5) Continuous indication for duration of a fire.
(6) Means for electrically testing the detector system from the aircraft cockpit.
(7) Detectors which resist exposure to oil, water, vibration, extreme temperatures.
maintenance handling.
(8) Detectors which are light in weight and casily adaptable to any mounting position.
(9) Detector circuitry which operates directly from the aircraft power system without inverters.
(10) Minimum electrical current requirements when not indicating a fire.
(11) Each detection system should actuate a cockpit light indicating the location of the fire and an audible alarm system.
(12) A separate detection system for cach engine.
There are a number of detectors or sensing devices available. Many older model aircraft still operating are equipped with some type of thermal switch system or thermocouple system.
FIRE DETECTION SYSTEMS A fire detection system should signal the presence of a fire. Units of the system are installed in locations where there are greater possibilities of a fire.
Three detector systems in common use are the thermal switch system, thermocouple system, and the continuous-loop detector system.
Thermal Switch System A thermal switch system consists of one or more lights energized by the aircraft power system and thermal switches that control operation of the light (s). These thermal switches are heat-sensitive units that complete electrical circuits at a certain temperature. They are connected in parallel with each other but in series with the indicator lights (figure 10-1). If the temperature rises above a set value in any one section of the circuit, the thermal switch will close, completing the light circuit to indicate the presence of a fire or overheat condition.
auxiliary test circuit. The circuit in figure 10-1 includes a test relay. With the relay contact in the position shown, there are two possible paths for current flow from the switches to the light. This is an additional safety feature. Energizing the test relay completes a series circuit and checks all the wiring and the light bulb.
Also included in the circuit shown in figure 10-1 is a dimming relay. By energizing the dimming relay, the circuit is altered to include a resistor in series with the light. In some installations several circuits are wired through the dimming relay, and all the warning lights may be dimmed at the same time.
The thermal switch system uses a bimetallic thermostat switch or spot detector similar to that shown in figure 10-2. Each detector unit consists of a bimetallic thermoswitch. Most spot detectors are dual-terminal thermoswitches.
Indientor light FIGURE 10-1. Thermal switch fire circuit.
No set number of thermal switches is required;
the exact number is usually determined by the aircraft manufacturer. On some installations several thermal detectors are connected to one light; on others there may be only one thermal switch for an indicator light.
Some warning lights are the "push-to-test" type.
The bulb is tested by pushing it in to complete an FIGURE 10-2. Fenwal spot detector.
Fenwal Spot Detector Fenwal spot detectors are wired in parallel between two complete loops of wiring, as illustrated in figure 10-3. Thus, the system can withstand one fault, either an electrical open circuit or a short to ground, without sounding a false fire warning. A double fault must exist before a false fire warning can occur. In case of a fire or overheat condition, the spot-detector switch closes and completes a circuit to sound an alarm.
The Fenwal spot-detector system operates without a control unit. When an overheat condition or a fire causes the switch in a detector to close, the alarm bell sounds and a warning light for the affected area is lighted.
Thermocouple Systems The thermocouple fire warning system operates on an entirely different principle than the thermal switch system. A thermocouple depends upon the rate of temperature rise and will not give a warning when an engine slowly overheats or a short circuit 408
- Individual spot detectors - Alarm light and bell Wiring loop B- -Y 28v.d.c.
Loop A Test switch FIGURE 10-3. Fenwal spot-detector circuit.
develops. The system consists of a relay box, warning lights, and thermocouples. The wiring system of these units may be divided into the following circuits (figure 10-4) : (1) The detector circuit, (2) the alarm circuit, and (3) the test circuit.
The relay box contains two relays, the sensitive relay and the slave relay, and the thermal test unit.
Such a box may contain from one to eight identical circuits, depending on the number of potential fire zones. The relays control the warning lights. In turn, the thermocouples control the operation of the relays. The circuit consists of several thermocouples in series with each other and with the sensitive relay.
The thermocouple is constructed of two dissimilar metals such as chromel and constantan. The point where these metals are joined and will be exposed to the heat of a fire is called a hot junction. There is also a reference junction enclosed in a dead air space between two insulation blocks. A metal cage surrounds the thermocouple to give mechanical protection without hindering the free movement of air to the hot junction.
If the temperature rises rapidly, the thermocouple produces a voltage because of the temperature difference between the reference junction and the hot junction. If both junctions are heated at the same rate, no voltage will result and no warning signal is given.
If there is a fire, however, the hot junction will heat more rapidly than the reference junction. The ensuing voltage causes a current to flow within the detector circuit. Any time the current is greater than 4 milliamperes (0.004 ampere), the sensitive relay will close. This will complete a circuit from Thermocouples Sensitive relay Slave relay Detector circuit Test switch Alarm circuit Test circuit Heater FICURE 10-4. Thermocouple fire warning circuit.
the aircraft power system to the coil of the slave relay which closes and completes the circuit to the fire-warning light.
The total number of thermocouples used in individual detector circuits depends on the size of the fire zone and the total circuit resistance. The total resistance usually does not exceed 5 ohms. As shown in figure 10-4, the circuit has two resistors.
The resistor connected across the terminals of the slave relay absorbs the coil's self-induced voltage.
This is to prevent arcing across the points of the sensitive relay, since the contacts of the sensitive relay are so fragile they would burn or weld if arcing were permitted.
When the sensitive relay opens, the circuit to the slave relay is interrupted and the magnetic field around its coil collapses. When this happens, the coil gets a voltage through self-induction, but with the resistor across the coil terminals, there is a path for any current flow as a result of this voltage.
Thus, arcing at the sensitive relay contacts is eliminated.
Continuous-Loop Detector Systems A continuous-loop detector or sensing system permits more complete coverage of a fire hazard area than any type of spot-type temperature detectors.
Continuous-loop systems are versions of the thermal switch system. They are overheat systems, heat-sensitive units that complete electrical circuits at a certain temperature. There is no rate-of-heat-rise sensitivity in a continuous-loop system. Two widely used types of continuous-loop systems are the Kidde and the Fenwal systems.
In the Kidde continuous-loop system (figure 10-5), two wires are imbedded in a special ceramic core within an Inconel tube.
FIGURE 10-5. Kidde sensing element.
FICURE 10-6. Fenwal sensing element.
One of the two wires in the Kidde sensing system is welded to the case at each end and acts as an internal ground. The second wire is a hot lead (above ground potential) that provides a current signal when the ceramic core material changes its resistance with a change in temperature.
Another continuous-loop system, the Fenwal system (figure 10-6), uses a single wire surrounded by a continuous string of ceramic beads in an Inconel tube.
The beads in the Fenwal detector are wetted with a eutectic salt which possesses the characteristic of suddenly lowering its electrical resistance as the sensing element reaches its alarm temperature. In both the Kidde and the Fenwal systems, the resistance of the ceramic or eutectic salt core material prevents electrical current from flowing at normal temperatures. In case of a fire or overheat condition, the core resistance drops and current flows between the signal wire and ground, energizing the alarm system.
The Kidde sensing elements are connected to a relay control unit. This unit constantly measures the total resistance of the full sensing loop. The system senses the average temperature, as well as any single hot spot.
The Fenwal system uses a magnetic amplifier control unit. This system is non-averaging but will sound an alarm when any portion of its sensing element reaches the alarm temperature.
Both systems continuously monitor temperatures in the affected compartments, and both will automatically reset following a fire or overheat alarm after the overheat condition is removed or the fire extinguished.
• Averaging section Discrete element Operating power unit Responder 115v.a.c.
400~ Light Bell Teflon sheath Diaphragm switch 115v.a.c. c- Fire test switch FIcURE 10-7. Lindberg fire detection system schematic.
Continuous Element System The Lindberg fire detection system (figure 10-7) is a continuous-element type detector consisting of a stainless steel tube containing a discrete element.
This element has been processed to absorb gas in proportion to the operating temperature set point.
When the temperature rises (due to a fire or overheat condition) to the operating temperature set point, the heat generated causes the gas to be released from the element. Release of the gas causes the pressure in the stainless steel tube to increase.
This pressure rise mechanically actuates the diaphragm switch in the responder unit, activating the warning lights and an alarm bell.
A fire test switch is used to heat the sensors, expanding the trapped gas. The pressure generated closes the diaphragm switch, activating the warring system.
Overheat Warning Systems Overheat warning systems are used on some aircraft to indicate high area temperatures that may lead to a fire. The number of overheat warning systems varies with the aircraft. On some aircraft, they are provided for each engine turbine and each nacelle, on others they are provided for wheel well areas and for the pneumatic manifoid.
When an overheat condition occurs in the detector area, the system causes a light on the fire control pane's to flash.
in most systems the detector is a type of thermal switch. Each detector is operated when the heat rises to a specified temperature. This temperature depends upon the system and the type and model of the aircraft. The switch contacts of the detector are on spring struts, which close whenever the meter case is expanded by heat. One contact of each detector is grounded through the detector mounting bracket. The other contacts of all detectors connect in parallel to the closed loop of the warning light circuit. Thus, the closed contacts of any one detector can cause the warning lights to burn.
When the detector contacts close, a ground is provided for the warning light circuit. Current then flows from an electrical bus through the warning lights and a flasher or keyer to ground. Because of the flasher in the circuit, the lights flash on and off to indicate an overheat condition.
TYPES OF FIRES The National Fire Protection Association has classified fires in threc basic types:
a. Class A fires, defined as fires in ordinary combustible materials such as wood, cloth, paper, upholstery materials, etc.
b. Class B fires, defined as fires in flammable petroleum products or other flammable or combustible liquids, greases, solvents, paints,
c. Class C fires, defined as fires involving energized electrical equipment where the electrical non-conductivity of the extinguishing media is of importance. In most cases where electrical equipnent is de-energized, extinguishers suitable for use on Class A or B fires may be employed effectively.
Aircraft fires, in flight or on the ground, may encompass either or all of these type fires. There- 411