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.
Typical System Operation The schematic of a vacuum system for a twin-engine aircraft is shown in figure 12-64. This vacuum system consists of the following components; two engine-driven pumps, two vacuum relief valves, two flapper-type check valves, a vacuum manifold, a vacuum restrictor for each turn-and-bank indicator, an engine four-way selector valve, one vacuum gage, and a turn-and-bank selector valve.
The left and right engine-driven vacuum pumps and their associated lines and components are isolated from each other, and act as two independent vacuum systems. The vacuum lines are routed from each vacuum pump through a vacuum relief valve and through a check valve to the vacuum four-way selector valve.
Turn-and-bank selector valve Copilot's turn-and-bank Right T & B Needle valve Copilot's artificial horizon From the engine four-way selector valve, which permits operation of the left or right engine vacuum system, the lines are routed to a vacuum manifold.
From the manifold, flexible hose connects the vacuum-operated instruments into the system. From the instrument, lines routed to the vacuum gage pass through a turn-and-bank selector valve. This valve has three positions: main, left T&B, and right T&B. In the main position the vacuum gage indicates the vacuum in the lines of the artificial horizon and directional gyros. In the other positions, the lower value of vacuum for the turn-and-bank indicators can be read.
VACUUM-DRIVEN ATTITUDE GYROS In a typical vacuum-driven attitude gyro system, air is sucked through the filter, then through pas-
- Vacuum gage Left T & B Main Pilot's turn-and-bank Needle valve Pilot's artificial horizon Pilot's directional gyro Vacuum manifokl Check valve • Vacuum 4-way selector valve Relief valve Relief valve Check valve Right engine vacuum pump Left engine vacuum pump FICURE 12-64. Vacuum system for a multi-engine aircraft.
Precess Port Port A Applied force Gyro precesses, Precession Exhaust air equal increasing exhaust Precessing force at port A erects gyro, in all directions. from port A. exhaust air again equal at all ports.
Gyro erect.
FICURE 12-65. Erecting mechanism of a vacuum-driven attitude indicator.
sages in the rear pivot and inner gimbal ring, then into the housing where it is directed against the rotor vanes through two openings on opposite sides of the rotor. The air then passes through four equally spaced ports in the lower part of the rotor housing and is sucked out into the vacuum pump or venturi (figure 12-65).
The chamber containing the ports is the erecting device that returns the spin axis to its vertical alignment whenever a precessing force, such as bearing friction, displaces the rotor from its horizontal plane. Four exhaust ports are each half-covered by a pendulous vane, which allows discharge of equal volumes of air through each port when the rotor is properly erected. Any tilting of the rotor disturbs the total balance of the pendulous vanes, tending to close one vane of an opposite pair while the opposite vane opens a corresponding amount. The increase in air volume through the opening port exerts a precessing force on the rotor housing to erect the gyro, and the pendulous vanes return to a balanced condition (figure 12-66).
The limits of the attitude indicator specified in the manufacturer's instructions refer to the maximum rotation of the gimbals beyond which the gyro will tumble. The bank limits of a typical vacuumdriven attitude indicator are from approximately 100° to 110°, and the pitch limits vary from approximately 60° to 70°, depending on the design of a specific unit. If, for example, the pitch limits are 60° with the gyro normally erected, the rotor will tumble when the aircraft climb or dive angle exceeds 60°. As the rotor gimbal hits the stops, the FIGURE 12-66. Action of pendulous vanes.
rotor precesses abruptly, causing excessive friction and wear on the gimbals. The rotor will normally preceas back to the horizontal plane at a rate of approximately 8° per min.
Many gyros include a caging device, used to erect the rotor to its normal operating position prior to fight or after tumbling, and a flag to indicate that the gyro must be uncaged before use.
Turning the caging knob prevents rotation of the gimbals and locks the rotor spin axis in its vertical position.
PRESSURE-OPERATED GYROS The availability of pressure pumps on which no lubrication is necessary makes pressure-operated gyro systems feasible. In such installations, air is pushed under pressure through gyro instruments, rather than being sucked through the system. Positive-pressure pumps are more efficient than vacuum pumps, especially at higher altitudes.
VACUUM SYSTEM MAINTENANCE PRACTICES Errors in the indications presented on the attitude indicator will result from any factor that prevents the vacuum system from operating within the design suction limits, or from any force that disturbs the free rotation of the gyro at design speed.
These include poorly balanced components, clogged filters, improperly adjusted valves and pump malfunction. Such errors can be minimized by proper installation, inspection, and maintenance practices.
Other errors, inherent in the construction of the instrument, are caused by friction and worn parts.
These errors, resulting in erratic preccasion and failure of the instrument to maintain accurate indications, increase with the service life of the instru- For the aviation mechanic the prevention or correction of vacuum system malfunctions usually consists of cleaning or replacing filters, checking Possible Cause Isolation Procedure Correction
(1) No Vacuum Pressure or Insufficient Pressure Defective vacuum gage. On multi-engine aircraft check Replace faulty vacuum gage.
opposite engine system on the gage.
Vacuum relief valve incorrectly Change valve adjustment. Make final adjustment to proper adjusted. setting of valve.
Vacuum relief valve installed Visually inapect. Install properly.
backwards.
Broken line. Visually inspect. Replace line.
Lines crossed. Visually inspect. Install lines properly.
Obstruction in vacuum line. Check for collapsed line. Clean and test line. Replace defective part (s).
Remove and inspect. Replace faulty pump.
Make valve adjustment and note Adjust to proper pressure.
pressure.
Clean and adjust relief valve. Replace valve if adjustment fails.
sted. Adjust relief valve to proper setting.
Check calibration of gage. Replace faulty gage.
spond Visually inspect. Uncage instrument.
Check filter. Replace or clean as necessary.
Check vacuum setting. Adjust relief valve to proper setting.
Replace instrument.
Check lines and vacuum system. Clean or replace lines and replace components of vacuum system as necessary.
Instrument filter clogged. Visually inspect. Replace filter.
Defective instrument. Test with properly functioning Replace faulty instrument.
instrument.
(6) Turn-and-Bank Pointer Vibrates Defective instrument. Test with properly functioning Replace defective instrument.
instrument FICURE 12-67. Vacuum system troubleshooting.
and correcting for insufficient vacuum, or removing and replacing the instruments. A list of the most common malfunctions, together with their correction, is included in figure 12-67.
ELECTRIC ATTITUDE INDICATOR In the past, suction-driven gyros have been favored over the electric type for light aircraft because of their comparative simplicity and lower cost. However, the increasing importance of the attitude indicator has stimulated development of improved electric-driven gyros suited for light plane installation. Improvements relating to basic gyro design factors, easier readability, erection characteristics, reduction of induced errors, and instrument limitations are reflected in several available types. Depending upon the particular design improvements, the details for the instrument display and cockpit controls will vary among different instruments. All of them present, to a varying degree, the essential pitch-and-bank information for attitude reference.
The typical attitude indicator, or gyro horizon as it is sometimes referred to, has a vertical-seeking gyro, the axis of rotation tending to point toward the center of the earth. The gyro is linked with a horizon bar and stabilizes a kidney-shaped sphere having pitch attitude markings. The sphere, horizon bar, and bank index pointer move with changes of aircraft attitude. Combined readings of these presentations give a continuous pictorial presentation of the aircraft attitude in pitch and roll with respect to the earth's surface.
magnet attached to the top of the gyro shaft spins at approximately 21,000 r.p.m. Around this magnet, but not attached, is a sleeve that is rotated by magnetic attraction at approximately 44 to 48 r.p.m.
As illustrated in figure 12-69, the steel balls are free to move around the sleeve. If the pull of gravity is not aligned with the axis of the gyro, the balls will fall to the low side. The resulting precession re-aligns the axis of rotation vertically.
The gyro can be caged manually by a lever and cam mechanism to provide rapid erection. When the instrument is not getting sufficient power for normal operation, an "off" flag appears in the upper right-hand face of the instrument.
Gyro stabilized horizon bar Power failure warning flag Vertical index Bank index Bezel ' mask The gyroscope motor is driven by 115 v., 400 Hz alternating current. The gyro, turning at approximately 21,000 r.p.m., is supported by the yoke and pivot assembly (gimbals). Attached to the yoke and pivot assembly is the horizon bar, which moves up and down through an arc of approximately 27°.
The kidney-shaped sphere provides a background for the horizon bar and has the words climb and dive and a bull's-eye painted on it. Climb and dive represent about 60° of pitch. Attached to the yoke and pivot assembly is the bank index pointer, which is free to rotate 360°. The dial face of the attitude indicator is marked with 0°, 10°, 20°, 30°, and 60° of bank, and is used with the bank index pointer to indicate the degree of bank left or right.
The face of one type of gyro-horizon is shown in figure 12-68.
The function of the erection mechanism is to keep the gyro axis vertical to the surface of the earth. A Trim knob Caging knob FIcuRE 12-68. Gyro-horizon indicator.
Magnetic Compass The magnetic compass is a simple, self-contained instrument which operates on the principle of magnetic attraction.
If a bar magnet is mounted on a pivot to be free to rotate in a horizontal plane, it will assume a position with one of its ends pointing toward the earth's north magnetic pole. This end of the magnet is called the north-seeking end of the magnet.