Tachometer Tachometer Generator Indicator Generator Indicator OB B Engine No. 2 Engine No. 3 When 3-wire system is required connect"C" terminal of all units. The "C" terminal is connected by internal grounds.
Synchroscope Tachometer Tachometer Generator Indicator Indicator Generator › B Engine No. 1 Engine No. 4 FICURE 12-40. Four-engine synchroscope schematic, turbine engines be known. Many other temperatures must also be known, but these are some of the more important. Different types of thermometers are used to collect and present this information.
Electrical Resistance Thermometer 50 Electrical resistance thermometers are used 100 widely in many types of aircraft to measure carburetor air, oil, and free air temperatures. °c 150 The principal parts of the electrical resistance -50 thermometer are the indicating instrument, the temperature-sensitive element (or bulb), and the connecting wires and plug connectors.
Oil temperature thermometers of the electrical resistance type have typical ranges of from -10° to + 120° C., or from -70° to + 150° C. Carburetor air and mixture thermometers may have a range of from -50° to +50° C., as do many free air thermometers.
A typical electrical resistance thermometer is FIcURE 12-41. Typical electrical resistance temperature shown in figure 12-41. Indicators are also available indicator.
in dual form for use in multi-engine aircraft. Most indicators are self-compensated for changes in cockpit temperature.
The electrical resistance thermometer operates on the principle of the change in the electrical resistance of most metals with changes in temperature. In most cases, the electrical resistance of a metal increases as the temperature rises. The resistance of some metals increases more than the resistance of others with a given rise in temperature. If a metallic resistor with a high temperature-resistant coefficient (a high rate of resistance rise for a given incresae in temperature) is subjected to a temperature to be measured, and a resistance indicator is connected to it, all the requirements for an electrical thermometer system are present.
The heat-sensitive resistor is the main element in the bulb. It is manufactured so that it has a definite resistance for each temperature value within its working range. The temperature-sensitive resistor element is a winding made of various alloys, such as nickel/manganese wire, in suitable insulating material. The resistor is protected by a closed-end metal tube attached to a threaded plug with a hexagon head (figure 12-42). The two ends of the winding are brazed or welded to an electrical receptacle designed to receive the prongs of the connector plug.
FIGURE 12-42. Two types of resistance thermometer bulb assemblies.
The electrical resistance indicator is a resistancemeasuring instrument. Its dial is calibrated in degrees of temperature instead of ohms and measures temperature by using a modified form of the Wheatstone-bridge circuit.
The Wheatstone-bridge meter operates on the principle of balancing one unknown resistor against other known resistances. A simplified form of a Wheatstone-bridge circuit is shown in figure 12-43.
14 volts + Indicator Cal. I Heat-sensitive element or bulb FIGURE 12-43. Wheatstone-bridge meter circuit.
Three equal values of resistances (A, B, and C, figure 12-43) are connected to a diamond-shaped bridge circuit with a resistance of unknown value (D).
The unknown resistance represents the resistance of the temperature bulb of the electrical resistance thermometer system.. A galvanometer calibrated to read in degrees is attached across the circuit at point X and Y.
When the temperature causes the resistance of the bulb to equal that of the other resistances, no potential difference exists between points X and Y in the circuit, and no current flows in the galvanometer leg of the circuit. If the temperature of the bulb changes, its resistance will also change, and the bridge becomes unbalanced, causing current to flow through the galvanometer in one direction or the other.
The dial of the galvanometer is calibrated in degrees of temperature, converting it to a temperature-measuring instrument. Most indicators are provided with a zero adjustment screw on the face of the instrument to set the pointer at a balance point (the position of the pointer when the bridge is balanced and no current flows through the meter).
Thermocouple Thermometer Indicators The cylinder temperature of most air-cooled reciprocating aircraft engines is measured by a thermometer which has its heat-sensitive element attached to some point on one of the cylinders (normally the hottest cylinder). In the case of turbojet engines, the exhaust temperature is measured by attaching thermocouples to the tailcone.
A thermocouple is a circuit or connection of two 493
unlike metals; such a circuit has two junctions. If cylinder head. Here again, the same metal is used in one of the junctions is heated to a higher tempera- the lead as in the part of the thermocouple to which ture than the other, an electromotive force is pro- it is connected.
duced in the circuit. By including a galvanometer in the circuit, this force can be measured. The hotter the high-temperature junction (hot junction) be- Thermocouple. gasket comes, the greater the electromotive force produced.
By calibrating the galvanometer dial in degrees it Back of becomes a thermometer. indicating instrument Engine A typical thermocouple thermometer system cylinder (figure 12-44) used to indicate engine temperature consists of a galvanometer indicator calibrated in Red Yellow degrees of centigrade, a thermocouple, and thermocouple leads. Red (A)
- Indicator Cold junction Thermocouple Yellow Engine cylinder wall Hot junction (B) Thermocouple leads FIGURE 12-45. Thermocouples: (A) Gasket type, (B) Bayonet type.
Black Connector Temperature to be measured FIcURE 12-44. Reciprocating engine cylinder head temperature thermocouple system.
Thermocouple leads are commonly made from iron and constantan, but copper and constantan or chromel and alumel are other combinations of dissimilar metals in use. Iron/constantan is used mostly in radial engines, and chromel/alumel is used in jet engines.
Thermocouple leads are designed to provide a definite amount of resistance in the thermocouple circuit. Thus, their length or cross-sectional size cannot be altered unless some compensation is made for the change in total resistance.
The hot junction of the thermocouple varies in shape depending on its application. Two common types are shown in figure 12-45; they are the gasket type and the bayonet type. In the gasket type, two rings of dissimilar metals are pressed together to form a spark plug gasket. Each lead that makes a connection back to the galvanometer must be made of the same metal as the part of the thermocouple to which it is connected. For example, a copper wire is connected to the copper ring and a constantan wire is connected to the constantan ring. The bayonet type thermocouple fits into a hole or well in the The cylinder chosen for installing the thermocouple is the one which runs the hottest under most operating conditions. The location of this cylinder varies with different engines.
The cold junction of the thermocouple circuit is inside the instrument case.
Since the electromotive force set up in the circuit varies with the difference in temperature between the "hot" and "cold" junctions, it is necessary to compensate the indicator mechanism for changes in cockpit temperature which affect the "cold" junction. This is accomplished by using a bimetallic spring connected to the indicator mechanism.
When the leads are disconnected from the indicator, the temperature of the cockpit area around the instrument panel can be read on the indicator dial.
This is because the bimetallic compensator spring continues to function as a thermometer.
Figure 12-46 shows the dials of two thermocouple temperature indicators.
Gas Temperature Indicating Systems EGT (exhaust gas temperature) is a critical variable of turbine engine operation. The EGT indicat.
ing system provides a visual temperature indication in the cockpit of the turbine exhaust gases as they leave the turbine unit. In certain turbine engines the temperature of the exhaust gases is measured at 494
• 2 150 250 3 2 50 NU, 350 °c X100
• FIGURE 12-46. Two types of thermocouple temperature indicators.
the entrance to the turbine unit. This is usually referred to as a TIT (turbine inlet temperature) indicating system. The principal disadvantages of this method are that the number of thermocouples required become greater and the environmental temperatures in which they must operate are increased.
A gas temperature thermocouple is mounted in a ceramic insulator and encased in a metal sheath;
the assembly forms a probe which projects into the exhaust gas stream. The thermocouple is made from chromel (a nickel/chromium alloy) and alumel (a nickel/aluminum alloy) . The hot junction protrudes into a space inside the sheath. The sheath has transfer holes in the end of it which allow the exhaust gases to flow across the hot junction.
Several thermocouples are used and are spaced at intervals around the perimeter of the engine turbine casing or exhaust duct. The thermocouples measure engine EGT in millivolts, and this voltage is applied to the amplifier in the cockpit indicator, where it is amplified and used to energize the servomotor which drives the indicator pointer.
A typical EGT thermocouple system is shown in figure 12-47.
The EGT indicator shown is a hermetically sealed unit and has provisions for a mating electrical connector plug. The instrument's scale ranges from O° C. to 1,200° C., with a vernier dial in the upper right-hand corner. A power "off" warning flag is located in the lower portion of the dial.
The TIT indicating system provides a visual indication at the instrument panel of the temperature of gases entering the turbine. On one type of turbine aircraft the temperature of each engine turbine inlet is measured by 18 dual-unit thermocouples installed in the turbine inlet casing. One set of these thermocouples is paralleled to transmit signals to the cockpit indicator. The other set of paralleled thermocouples provides temperature signals to the temperature datum control. Each circuit is electrically independent, providing dual system reliability.
The thermocouple assemblies are installed on pads around the turbine inlet case. Each thermocouple incorporates two electrically independent junctions within a sampling probe. The average voltage of the thermocouples at the thermocouple terminal blocks represents the TIT.
A schematic for the turbine inlet temperature system for one engine of a four-engine turbine aircraft is shown in figure 12-48. Circuits for the other three engines are identical to this system. The indicator contains a bridge circuit, a chopper circuit, a two-phase motor to drive the pointer, and a feed.
back potentiometer. Also included are a voltage reference circuit, an amplifier, a power "off" flag, a power supply, and an overtemperature warning light. Output of the amplifier energizes the variable field of the two-phase motor which positions the indicator main pointer and a digital indicator. The motor also drives the feedback potentiometer to pro- 495
CXIDO 115 v.a.c. bus Chromel Indicator Alumel Turbine outlet circuit breaker Dual thermocouple FicURE 12-47. Typical exhaust gas temperature thermocouple system.
Eng. No. 1 Bridge 18 1200 Thermocouples Digital indicator ØA warning light Overtemp reference voltage Zener Chopper Amplifier test switch Power supply Overtemp Eng. 1 warning
• Power oft light warning flag Eng. 2 ng.
FICURE 12-48. Turbine inlet temperature indicating system.
vide a huming signal to stop the drive motor when the correct pointer position, relative to the temperature signal, has been reached. The voltage reference circuit provides a closely regulated reference voltage in the bridge circuit to preclude error from input voltage variation to the indicator power supply.
The overtemperature warning light in the indicator illuminates when the TIT reaches a predetermined limit. An external test switch is usually installed so that overtemperature warning lights for all the engines can be tested at the same time. When the test switch is operated, an overtemperature signal is simulated in each indicator temperature con.
trol bridge circuit.
RATIOMETER ELECTRICAL RESISTANCE THER-
MOMETER The basic Wheatstone-bridge, temperature-indicating system provides accurate indications when the pointer is at the balance point on the indicator dial. As the pointer moves away from the balance point, the Wheatstone-bridge indicator is increasingly affected by supply voltage variations. Greater accuracy can be obtained by inserting one of several types of automatic line voltage compensating circuits into the circuit. Some of these voltage regulators employ the filament resistance of lamps to achieve a more uniform supply voltage. The resistance of the lamp filaments helps regulate the voltage applied to the Wheatstone-bridge circuit since the filament resistance changes in step with supply voltage variation.
The ratiometer is a more sophisticated arrangement for obtaining greater accuracy in resistancebulb indicators. The ratiometer measures the ratio of currents, using an adaptation of the basic Wheatstone-bridge with ratio circuitry for increased sensitivity.
A schematic of a ratiometer temperature circuit is shown in figure 12-49. The circuit contains two parallel branches, one with a fixed resistance in series with coil A, and the other a built-in resistance in series with coil B. The two coils are wound on a rotor pivoted in the center of the magnet air gap.
The permanent magnet is arranged to provide a larger air gap betweer the magnet and the coils at the bottom than at the top. This produces a flux density that is progressively stronger from the bottom of the air gap to the top.
The direction of the current through each coil in respect to the polarity of the permanent magnet N S Sensitive element (bulb) FIGURE 12-49. Ratiometer temperature-measuring system schematic.
causes the coil with the greater current flow to react in the weaker magnetic field. If the resistance of the temperature bulb is equal to the value of the fixed resistance, and equal values of current are flowing through the coils, the torque on the coils will be the same and the indicator points will be in the vertical (zero) position.
If the bulb temperature increases, its resistance will also increase, causing the current through the coil B circuit branch to decrease. Consequently, the torque on coil B decreases and coil A pushes downward into a weaker magnetic field; coil A, with its weaker current flow, moves into a stronger magnetic field. The torques on the coils still balance since the product of current times flux remains the same for both coils, but the pointer has moved to a new position on the calibrated scale. Just the opposite of this action would take place if the temperature of the heat-sensitive bulb should decrease.
Ratiometer temperature-measuring systems are used to measure engine oil, outside air, and carburetor air temperatures in many types of aircraft.
They are especially in demand to measure temperature conditions where accuracy is important or large variations of supply voltages are encountered.
FUEL FLOWMETER SYSTEMS Fuel flowmeter systems are used to indicate fuel usage. They are most commonly installed on large multi-engine aircraft, but they may be found on any type of aircraft if fuel economy is an important factor.
A typical flowmeter system for a reciprocating engine consists of a flowmeter transmitter and an indicator. The transmitter is usually connected into 497
the fuel line leading from the carburetor outlet to the fuel feed valve or discharge nozzle. The indicator is normally mounted in the instrument panel.
A cross sectional view of a typical transmitter fuel chamber is shown in figure 12-50. Fuel entering the inlet side of the fuel chamber is directed against the metering vane, causing the vane to swing on its shaft within the chamber. As the vane is moved from a closed position by the pressure of the fuel flow, the clearance between the vane and the fuel chamber wall becomes increasingly larger.
Figure 12-51 shows an exploded view of a fuel flowmeter system. Note that the metering vane moves against the opposing force of a hairspring.
When the force created by a given fuel flow is Relief valve Fuel inlet Fuel outlet Metering vane FIGURE 12-50. Flowmeter fuel chamber.
balanced by spring tension, the vane becomes stationary. The vane is connected magnetically to the rotor of a transmitter, which generates electrical signals to position the cockpit indicator. The distance the metering vane moves is proportional to, and a measure of, the rate of fuel flow.
The damper vane of the transmitter cushions fluctuations caused by air bubbles. The relief valve bypasses fuel to the chamber outlet when the flow of fuel is greater than chamber capacity.
A simplified schematic of a vane-type flowmeter system (figure 12-52) shows the metering vane connected to the flowmeter transmitter and the rotor and stator of the indicator connected to a common power source with the transmitter.
The dial of a fuel-flow indicator is shown in figure 12-53. Some fuel-flow indicators are calibrated in gallons per hour, but most of them indicate the measurement of fuel flow in pounds.
The fuel flowmeter system used with turbine engine aircraft is usually a more complex system than that used in reciprocating engine aircraft.
Hairspring
- Damper vane Metering vane (Indicator ) (Transmitter) FIcURE 12-51. Fuel flowmeter system.
Transmitter Indicator Fuel flow - Restraining spring FIGURE 12-52. Schematic of vane-type fowmeter system.
mounted in the main fuel line leading to the engine.
The impeller is driven at a constant speed by a 35 37' special three-phase motor. The impeller imparts an angular momentum to the fuel, causing the turbine 2 3° to rotate until the calibrated restraining spring force balances the force due to the angular momentum of the fuel. The deflection of the turbine posi- FUEL FLOW tions the permanent magnet in the position trans- LBS PER HR X 100 mitter to a position corresponding to the fuel flow in the line. This turbine position is transmitted electrically to the indicator in the cockpit.
FICURE 12-53. Typical fuel-flow indicator.
In the system shown schematically in figure 12-54, two cylinders, an impeller, and a turbine are
GYROSCOPIC INSTRUMENTS Three of the most common flight instruments, the attitude indicator, heading indicator, and the turn needle of the turn-and-bank indicator, are controlled by gyroscopes. To understand how these instruments operate requires a knowledge of gyroscopic principles, instrument power systems, and the operating principles of each instrument.
A gyroscope is a wheel or disk mounted to spin rapidly about an axis, and is also free to rotate about one or both of two axes perpendicular to Decoupling disk Fuel flow Second harmonic transmitter 1 Impeller motor Calibrated restraining Fluid passage Fluid passage springs 7J 115 v.a.c.
Timing contacts 28 v.d.c.
Indicator Power supply FIGURE 12-54. Schematic of a large turbine engine flowmeter system.
each other and to the axis of spin. A spinning gyroscope offers resistance to any force which tends to change the direction of the axis of spin.
A rotor and axle are the heart of a basic gyro (A of figure 12-55); a supporting ring with bearings on which the rotor and its axle can revolve are added to the basic unit (B of figure 12-55); and an outer ring with bearings at 90° to the rotor bearings has been added (C of figure 12-55). The inner ring with its rotor and axle can turn through 360° inside this outer ring.
A gyro at rest is shown in six different positions (figure 12-56) to demonstrate that unless the rotor is spinning a gyro has no unusual properties; it is simply a wheel universally mounted.
When the rotor is rotated at a high speed, the gyro exhibits one of its two gyroscopic characteristics. It acquires a high degree of rigidity, and its axle points in the same direction no matter how much its base is turned about (figure 12-57).
C FICURE 12-55. Basic gyroscope.
Gyrocsopic rigidity depends upon several design factors:
(1) Weight. For a given size, a heavy mass is more resistant to disturbing forces than a light mass.
(2) Angular velocity. The higher the rotational speed, the greater the rigidity or resistance to deflection.
(3) Radius at which the weight is concentrated. Maximum effect is obtained from a mass when its principal weight is concentrated near the rim rotating at high speed.
(4) Bearing friction. Any friction applies a deflecting force to a gyro. Minimum bearing friction keeps deflecting forces at a minimum.
A second gyroscopic characteristic, precession, is illustrated in figure 12-58A by applying a force or pressure to the gyro about the horizontal axis. The applied force is resisted, and the gyro, instead of turning about its horizontal axis, turns or "precesses" about its vertical axis in the direction indicated by the letter P. In a similar manner, if pressure is applied to the vertical axis, the gyro will precess about its horizontal axis in the direction shown by the arrow P in figure 12-58B.
Force Force FIGURE 12-58. Gyrosocopic precession.
FIcURE 12-56. A gyro at rest. Two types of mountings are used, depending upon how the gyroscopic properties are to be used in the operation of an instrument. A freely or universally mounted gyro is set on three gimbals (rings), with the gyro free to rotate in any plane.
Regardless of the position of the gyro base, the gyro tends to remain rigid in space. In the attitude indicator the horizon bar is gyro-controlled to remain parallel to the natural horizon, and changes in position of the aircraft are shown pictorially on the FIGURE 12-57. Gyroscope inertia. indicator.
The semirigid, or restricted, mounting employs two gimbals, limiting the rotor to two planes of rotation. In the turn-and-bank indicator, the semirigid mounting provides controlled precession of the rotor, and the precessing force exerted on the gyro by the turning aircraft causes the turn needle to indicate a turn.
SOURCES OF POWER FOR GYRO OPERATION The gyroscopic instruments can be operated either by a vacuum system or an electrical power source. In some aircraft, all the gyros are either vacuum or electrically motivated; in others, vacuum (suction) systems provide the power for the attitude and heading indicators, while the electrical system drives the gyro for operation of the turn needle. Either alternating or direct current is used to power the gyroscopic instruments.
Vacuum System The vacuum system spins the gyro by sucking a stream of air against the rotor vanes to turn the rotor at high speed, essentially as a water wheel or turbine operates. Air at atmospheric pressure drawn through a filter or filters drives the rotor vanes, and is sucked from the instrument case through a line to the vauuum source and vented to the atmosphere.
Either a venturi or a vacuum pump can be used to provide the vacuum required to spin the rotors of the gyro instruments.
The vacuum value required for instrument operation is usually between 3½ in. to 4½ in. He and is usually adjusted by a vacuum relief valve located in the supply line. The turn-and-bank indicators used in some installations require a lower vacuum setting. This is obtained using an additional regulating valve in the individual instrument supply line.
Venturi-Tube Systems The advantages of the venturi as a suction source are its relatively low cost and simplicity of installation and operation. A light, single-engine aircraft can be equipped with a 2-in. venturi (2 in. 11g vacuum capacity) to operate the turn needle. With an additional 8-in. venturi, power is available for the attitude and heading indicators. A venturi vacuum system is shown in figure 12-59.
The line from the gyro (figure 12-59) is connected to the throat of the venturi mounted on the exterior of the aircraft fuselage. Throughout the normal operating airspeed range the velocity of the Turn-and-bank indicator Pressure gage Attitude indicator Heading indicator FIGURE 12-59. Venturi vacuum system.
air through the venturi creates sufficient suction to spin the gyro.
The limitations of the venturi system should be evident from the illustration in figure 12-59. The venturi is designed to produce the desired vacuum at approximately 100 m.p.h. under standard sealevel conditions. Wide variations in airspeed or air density, or restriction to airfow by ice accretion, will affect the pressure at the venturi throat and thus the vacuum driving the gyro rotor. And, since the rotor does not reach normal operating speed until after takeoff, preflight operational checks of venturi-powered gyro instruments cannot be made.
For this reason the system is adequate only for light-aircraft instrument training and limited flying under instrument weather conditions. Aircraft flown throughout a wider range of speed, altitude, and weather conditions require a more effective source of power independent of airspeed and less susceptible to adverse atmospheric conditions.
Engine-Driven Vacuum Pump The vane-type engine-driven pump is the most common source of vacuum for gyros installed in general aviation light aircraft. One type of enginedriven pump is mounted on the accessory drive shaft of the engine, and is connected to the engine lubrication system to seal, cool, and lubricate the pump.
Another commonly used source of vacuum is the dry vacuum pump, also engine-driven. The pump operates without lubrication, and the installation requires no lines to the engine oil supply, and no air-oil separator or gate check valve. In other respects, the dry pump system and oil lubricated system are the same.
The principal disadvantage of the pump-driven vacuum system relates to erratic operation in highaltitude flying. Apart from routine maintenance of the filters and plumbing, which are absent in the electric gyro, the engine-driven pump is as effective a source of power for light aircraft as the electrical system.
Typical Pump-Driven Vacuum System Figure 12-61 shows the components of a vacuum system with a pump capacity of approximately 10" Hg at engine speeds above 1000 rpm. Pump capacity and pump size vary in different aircraft, depending on the number of gyros to be operated.
Inlet Shaft Vane
- Rotor Case FIcURE 12-60. Cutaway view of a vane-type enginedriven vacuum pump.
Heading Air indicator filter Air/oil separator Attitude Air indicator Vacuum Oil Selector pump Suction Pressure Gate valve Suction gage relief valve relief valve check valve Turn-and - bank indicator Restrictor valve FIcURE 12-61. Typical pump-driven vacuum system.
Air-Oil Separator.--Oil and air in the vacuum pump are exhausted through the separator, which separates the oil from the air; the air is vented outboard, and the oil is returned to the engine sump.
Suction Relief Valve.-Since the system capacity is more than is needed for operation of the instruments, the adjustable suction relief valve is set for the vacuum desired for the instruments. Excess suction in the instrument lines is reduced when the spring-loaded valve opens to atmospheric pressure.
(See fig. 12-62.)
Pressure Relief Valve.-Since a reverse flow of air from the pump would close both the gate check valve and the suction relief valve, the resulting pressure could rupture the lines. The pressure relief valve vents positive pressure into the atmosphere.
Gate Check Valve.-The gate check valve prevents possible damage to the instruments by engine back-fire, which would reverse the flow of air and oil from the pump. (See fig. 12-63.)
Selector Valve.--In twin-engine aircraft having vacuum pumps driven by both engines, the alternate pump can be selected to provide vacuum in the event of either engine or pump failure, with a check valve incorporated to seal off the failed pump.
Restrictor Valve.--Since the turn needle operates on less vacuum than that required for other gyro instruments, the vacuum in the main line must be reduced. This valve is either a needle valve adjusted to reduce the vacuum from the main line by approximately one-half, or a springloaded regulating valve that maintains a constant vacuum for the turn indicator, unless the main line vacuum falls below a minimum value.
Suction Gage.-The suction gage is a pressure gage, indicating the difference in inches of mercury, between the pressure inside the system and atmospheric or cockpit pressure. The desired vacuum, and the minimum and maximum limits, vary with gyro design. If the desired vacuum for the attitude and heading indicators is 5" and the minimum in 4.6", a reading below the latter value indicates that the airflow is not spinning the gyros fast enough for reliable operation. In many aircrait, the system provides a suction gage selector valve, permitting the pilot to check the vacuum at several points in the system:
Suction Suction pressures discussed in conjunction with the operation of vacuum systems are actually minus or negative pressures (below sea level). For example, if sea level equals 17.5 p.s.i, then 1" Hg (1 inch mercury) or 1 p.s.i. vacuum is equal to
- 1 p.si. negative pressure or 10.5 positive pressure. Likewise 3" Hg= -3 p.s.i. negative pressure or +14.5 positive pressure.
Of course, for every action there is an equal and opposite reaction. Therefore when the vacuum pump develops a vacuum (negative pressure) it must also create pressure (positive). This pressure (compressed air) is sometimes utilized to operate pressure instruments, deicer boots and inflatable seals.
• Spring To vacuum pump -
• Wire screen FICURE 12-62. Vacuum regulator valve.
Air filter.-The master air filter screens foreign matter from the air flowing through all the gyro Cate check valve open Gate check valve closed instruments, which are also provided with indi- - Spring vidual filters. Clogging of the master filter will reduce airflow and cause a lower reading on the suction gage. In aircraft having no master filter Air flee from installed, each instrument has its own filter. With Air flens to vacuum pump an individual filter system, clogging of a filter will not necessarily show on the suction gage. FIGURE 12-63. Gate check valve.