The range of the gage is typically from zero p.s.i.
to 20 p.s.i., with the scale marked in 2-p.s.i. graduations as shown in figure 12-5.
When installed and connected into an aircraft's de-icing pressure system, the gage reading always remains at zero unless the de-icing system is operating. The gage pointer will fluctuate from zero p.s.i.
to approximately 8 p.s.i. under normal conditions, because the de icer boots are periodically inflated and deflated. This normal fluctuation should not be confused with oscillation.
Diaphragm-Type Pressure Gages This type of pressure gage uses a diaphragm for measuring pressure. The pressure or suction to be measured is admitted to the pressure-sensitive diaphragm through an opening in the back of the instrument case (figure 12-6).
Diaphragm Pressure entrance Pinion gear Rocker arm Pointer Sector gear Vent J FICURE 12-6. Diaphragm-type pressure gage.
An opposing pressure, such as that of the atmosphere, is admitted through a vent in the case (figure 12-6). Since the walls of the diaphragm are very thin, an increase of pressure will cause it to expand, and a decrease in pressure will cause it to contract. Any movement of the diaphragm is transferred to the pointer by means of the rocker shaft, sector, and pinion, which are connected to the front side of the diaphragm. This gage is also a differential-pressure measuring device since it indicates the difference between the pressure applied at the vent of the case and the pressure or suction inside the diaphragm.
Suction Gages Suction gages are used on aircraft to indicate the amount of suction that actuates the air-driven gyroscopic instruments. The spinning rotors of gyroscopic instruments are kept in motion by streams of air directed against the rotor vanes. These airstreams are produced by pumping air out of the instrument cases by the vacuum pump. Atmospheric pressure then forces air into the cases through filters, and it is this air that is directed against the rotor vanes to turn them.
The suction gage indicates whether the vacuum system is working properly. The suction gage case is vented to the atmosphere or to the line of the air filter, and contains a pressure-sensitive diaphragm Plus the usual multiplying mechanism which amplifies the movement of the diaphragm and transfers it to the pointer. The reading of a suction gage indicates the difference between atmospheric pressure and the reducd pressure in the vacuum system.
Manifold Pressure Gage The manifold pressure gage is an important instrument in an aircraft powered by a reciprocating engine. The gage is designed to measure absolute pressure. This pressure is the sum of the air pressure and the added pressure created by the supercharger. The dial of the instrument is calibrated in inches of mercury (Hg).
When the engine is not running, the manifold pressure gage records the existing atmospheric pressure. When the engine is running, the reading obtained on the manifold pressure gage depends on the engine's r.p.m. The manifold pressure gage indicates the manifold pressure immediately before the cylinder intake ports.
The schematic of one type of manifold pressure gage is shown in figure 12-7. The outer shell of the gage protects and contains the mechanism. An opening at the back of the case provides for the connection to the manifold of the engine.
The gage contains an aneroid diaphragm and a linkage for transmitting the motion of the diaphragm to the pointer. The linkage is completely external to the pressure chamber, and thus is not exposed to the corrosive vapors of the manifold.
The pressure existing in the manifold enters the sealed chamber through a damping tube, which is a short length of capillary tubing at the rear of the 473
Altitude compensating aneroid Manifold pressure diaphragm Capillary assembly Frame Rocking shaft FIGURE 12-7. Manifold pressure gage.
case. This damping tube acts as a safety valve to prevent damage to the instrument by engine backfire. The sudden surge of pressure caused by backfire is considerably reduced by the restricted capillary tubing.
When installing a manifold pressure gage, care should be taken to ensure that the pointer is vertical when registering 30 in. Hg.
When an engine is not running, the manifold pressure gage reading should be the same as the local barometric pressure. It can be checked against a barometer known to be in proper operating condition. In most cases the altimeter in the aircraft can be used, since it is a barometric instrument. With the aircraft on the ground, the altimeter hands should be set to zero and the instrument panel tapped lightly a few times to remove any possible frictional errors. The barometer scale on the altimeter face will indicate local atmospheric pressure when the altimeter hands are at zero. The manifold pressure gage should agree with this pressure reading. If it does not, the gage should be replaced with a gage that is operating properly.
If the pointer fails to respond entirely, the mechanism is, in all probability, defective. The gage should be removed and replaced. If the pointer responds but indicates incorrectly, there may be moisture in the system, obstruction in the lines, a leak in the system, or a defective mechanism.
When doubt exists about which of these items is the cause of the malfunction, the engine should be operated at idle speed and the drain valve (usually located near the gage) opened for a few minutes.
This will usually clear the system of moisture. To clear an obstruction, the lines may be disconnected and blown clear with compressed air. The gage mechanism may be checked for leaks by disconnecting the line at the engine end and applying air pressure until the gage indicates 50 in. Hg. Then the line should be quickly closed. A leak is present if the gage pointer returns to atmospheric pressure.
If a leak is evident but cannot be found, the gage should be replaced.
PITOT-STATIC SYSTEM Three of the most important flight instruments are connected into a pitot-static system. These instruments are the airspeed indicator, the altimeter, and the rate-of-climb indicator. Figure 12-8 shows these three instruments connected to a pitot-static tube head.
Impact (pitot) pressure Rate of climb Airspeed Altimeter FIGURE 12-8. Pitot-static system.
The pitot-static system head, or pitot-static tube as it is sometimes called, consists of two sections. As shown in figure 12-9, the forward section is open at the front end to receive the full force of the impact air pressure. At the back of this section is a baffle plate to protect the pitot tube from moisture and dirt that might otherwise be blown into it. Moisture can escape through a small drain hole at the bottom of the forward section.
The pitot, or pressure, tube leads back to a chamber in the "shark-fin" projection near the rear of the assembly. A riser, or upright tube, leads the air from this chamber through tubing to the airspeed indicator.
Pressure chamber Static tube Pitot tube riser riser Pitot tabe Static hole . Drain hole Baffle plate Drain Hote Static hole . Static Electrical chamber contact Trater Heater FIGURE 12-9. Pitot-static system head.
The rear, or static, section of the pitot-static tube head is pierced by small openings on the top and bottom surfaces. These openings are designed and located so that this part of the system will provide accurate measurements of atmospheric pressure in a static, or still, condition. The static section contains a riser tube which is connected to the airspeed indicator, the altimeter, and the rate-of-climb indicator.
Many pitot-static tubes are provided with heating elements to prevent icing during flight (figure 12-9). During ice-forming conditions, the electrical heating elements can be turned on by means of a switch in the cockpit. The electrical circuit for the heater element may be connected through the ignition switch. Thus, in case the heater switch is inadvertently left in the "on" position, there will be no drain on the battery when the engine is not operating.
The pitot-static tube head is mounted on the outside of the aircraft at a point where the air is least likely to be turbulent. It is pointed in a forward direction parallel to the aircraft's line of flight. One general type of tube head is designed for mounting on a streamlined mast extending below the nose of the aircraft fuselage. Another type is designed for installation on a boom extending forward of the leading edge of the wing. Both types are shown in figure 12-10. Although there is a slight difference in their construction, they operate identically.
Most pitot-static tubes are manufactured with a union connection in both lines from the head, near Wing leading edge mounting Fuselage nose mounting FIcURE 12-10. Pitot-static tube heads.
In many aircraft equipped with a pitot-static tube, an alternate source of static pressure is provided for emergency use. A schematic diagram of a typical system is shown in figure 12-11. As shown in the diagram, the alternate source of static pressure may be vented to the interior of the aircraft.
Another type of pitot-static system provides for the location of the pitot and static sources at sepa- Pitot-static tube Pitot line Static linethe point at which the tube head is attached to the mounting boom or mast (figure 12-10). These connections simplify removal and replacement, and Alternate static source are usually reached through an inspection door in (cockpit air) the wing or fuselage. When a pitot-static tube head is to be removed, these connections should be disconnected before any mounting screws and lock- FICURE 12-11. Pitot-static system with alternate source washers are removed. of static pressure.
Sources of Alt. V/S A/S static pressure Pitot line (Schematic diagram of pitot head) Static vent Static line Heating line Pitot-pressure chamber Drain opening Pitot-heat element FIcURE 12-12. Pitot-static system with separate sources of pressure.
rate positions on the aircraft. This type of system is illustrated in figure 12-12.
Impact pressure is taken from the pitot-head (figure 12-12) which is mounted parallel to the longitudinal axis of the aircraft and generally in line with the relative wind. The leading edge of the wing, nose section, or vertical stabilizer are the usual mounting positions, since at those points there is usually a minimum disturbance of air due to motion of the aircraft.
Static pressure in this type of pitot-static system is taken from the static line attached to a vent or vents mounted flush with the fuselage or nose section. On aircraft using a flush-mounted static source, there may be two vents, one on each side of the aircraft. This compensates for any possible variation in static pressure on the vents due to erratic changes in aircraft attitude. The two vents are usually connected by a Y-type fitting. In this type of system, clogging of the pitot opening by ice or dirt (or failure to remove the pitot cover) affects the airspeed indicator only.
A pitot-static system used on a pressurized, mülti-engine aircraft is shown in figure 12-13. Three additional units, the cabin pressure controller, the cabin differential pressure gage, and the autopilot system are integrated into the static system. Both heated and unheated flushmounted static ports are used.
Altimeters There are many kinds of altimeters in general use today. However, they are all constructed on the same basic principle as an aneroid barometer. They all have pressure-responsive elements (aneroids) which expand or contract with the pressure change of different flight levels. The heart of an altimeter is its aneroid mechanism (figure 12-14). The expansion or contraction of the aneroid with pressure changes actuates the linkage, and the indicating hands show altitude. Around the aneroid mechanism of most altimeters is a device called the bimetal yoke. As the name implies, this device is composed of two metals and performs the function of compensating for the effect that temperature has on the metals of the aneroid mechanism.
The presentation of altitude by altimeters in current use varies from the multi-pointer type to the drum and single pointer, and the digital counter and single pointer types.
The dial face of the typical altimeter is graduated with numerals from zero to 9 inclusive, as shown in figure 12-15. Movement of the aneroid element is transmitted through a gear train to the three hands on the instrument face. These hands sweep the calibrated dial to indicate the altitude of the aircraft.
The shortest hand indicates altitude in tens of thousands of feet; the intermediate hand, in thousands of feet; and the longest hand, in hundreds of feet in 20-ft. increments. A barometric scale, located at the right of the instrument face, can be set by a knob located at the lower left of the instrument case. The barometric scale indicates barometric pressure in inches of mercury.
Since atmospheric pressure continually changes, the barometric scale must be re-set to the local station altimeter setting before the altimeter will indicate the correct altitude of the aircraft above 476
Pilot's instrument Copilot's instrument pitot tube Flush-mounted pitot tube ' Flush-mounted unheated unheated static tube Static vent -Static vent drain static tube drain valve in valve in nose nose gear well gear well Copilot's equalizer manifold Pilot's equali manifold Sumps and drains Flush-mounted Flush-mounted heated static tubes heated static tubes Auto- Pilot's flight instrument panel (2) instrument panel Copilot's flight pilot static drain valve Static manifold Static drain tee
- Static manifold
• Static selector Autopilot Static valve amplifier Cabin pressure control panel valve selector
1. Altimeter indicator Airspeed indicator Static system Drain valves 3. 4. Rate-of-climb indicator Cabin pressure controller mum Pitot pressure system
5. Cabin differential pressure gage To alternate static source FicURE 12-13. Schematic of typical pitot-static system on pressurized multi-engine aircraft.
sea level. When the setting knob is turned, the barometric scale, the hands, and the aneroid element move to align the instrument mechanism with the new altimeter setting.
Two setting marks, inner and outer, indicate barometric pressure in feet of altitude. They operate in conjunction with the barometric scale, and indications are read on the altimeter dial. The outer mark indicates hundreds of feet, and the inner mark thousands of feet. Since there is a limit to the graduations which can be placed on the barometric scale, the setting marks are used when the barometric pressure to be read is outside the limits of the scale.
In addition to the errors in the altimeter mechanism, another error called installation error affects Barometric the accuracy of indications. The error is caused by dial the change of alignment of the static pressure port with the relative wind. The change of alignment is caused by changes in the speed of the aircraft and in the angle of attack, or by the location of the Handstatts static port in a disturbed pressure field. Improper Balance / 100-ft. scale • Adjustment knob ass y- 1,000 ft. scale Out: Sets barometrical dial installation or damage to the pitot-static tube will 10,000-ft. scale In: Sets dial hands also result in improper indications of altitude.
FIGURE 12-14. Mechanism of a sensitive altimeter. Rate-of-Climb Indicators The rate-of-climb, or vertical velocity, indicator (figure 12-16) is a sensitive differential pressure gage that indicates the rate at which an aircraft is climbing or descending. The rate-of-climb indicator is connected to the static system and senses the 100 FEET rate of change of static pressure.
9 1000 FEET indicator dial, is positive in a climb and negative The rate of altitude change, as shown on the 10 000 FEET 8 29€ 4 Restricted passage FICURE 12-15. Sensitive altimeter.
Altimeter Errors rors. A common one is that the scale is not correctly Altimeters are subject to various mechanical er- 2 oriented to standard pressure conditions. Altimeters 4 should be checked periodically for scale errors in IMB altitude chambers where standard conditions exist.
Another mechanical error is the hysteresis error.
This error is induced by the aircraft maintaining a DOWN given altitude for an extended period of time, then suddenly making a large altitude change. The resulting lag or drift in the altimeter is caused by the elastic properties of the materials which comprise the instrument. This error will eliminate itself with slow climbs and descents or after maintaining a new altitude for a reasonable period of time. FIcURE 12-16. Typical rate-of-climb indicator.
when descending in altitude. The dial pointer moves in either direction from the zero point, depending on whether the aircraft is going up or down. In level flight the pointer remains at zero.
The operation of a climb indicator is illustrated in figure 12-16. The case of the instrument is airtight except for a small connection through a restricted passage to the static line of the pitot-static system.
Inside the sealed case of the rate-of-climb indicator is a diaphragm with connecting linkage and gearing to the indicator pointer. Both the diaphragm and the case receive air at atmospheric pressure from the static line. When the aircraft is on the ground or in level fight, the pressures inside the diaphragm and the instrument case remain the same and the pointer is at the zero indication.
When the aircraft climbs the pressure inside the diaphragm decreases but, due to the metering action of the restricted passage, the case pressure will remain higher and cause the diaphragm to contract.
The diaphragm movement actuates the mechanism, causing the pointer to indicate a rate of climb.
When the aircraft levels off, the pressure in the instrument case is equalized with the pressure in the diaphragm. The diaphragm returns to its neutral position and the pointer returns to zero.
In a decent, the pressure conditions are reversed. The diaphragm pressure immediately becomes greater than the pressure in the instrument case. The diaphragm expands and operates the pointer mechanism to indicate the rate of descent.
When the aircraft is climbing or descending at a constant rate, a definite ratio between the diaphragm pressure and the case pressure is maintained through the calibrated restricted passage, which requires approximately 6 to 9 sec. to equalize the two pressures, causing a lag in the proper reading. Any sudden or abrupt changes in the aircraft's attitude may cause erroneous indications due to the sudden change of airflow over the static ports.
The instantaneous rate-of-climb indicator is a more recent development which incorporates acceleration pumps to eliminate the limitations asso.
ciated with the calibrated leak. For example, during an abrupt climb, vertical acceleration causes the pumps to supply extra air into the diaphragm to stabilize the pressure differential without the usual lag time. During level fight and steady-rate climbs and descents, the instrument operates on the same principles as the conventional rate-of-climb indicator.
A zero-setting system, controlled by a setscrew or an adjusting knob permits adjustment of the pointer to zero. The pointer of an indicator should indicate zero when the aircraft is on the ground or maintaining a constant pressure level in flight.
Airspeed Indicator Airspeed indicators are sensitive pressure gages which measure the difference between the pitot and static pressures, and present such difference in terms of indicated airspeed. Airspeed indicators are made by various manufacturers and vary in their mechanical construction. However, the basic construction and operating principle is the same for all types.
The airspeed indicator (figure 12-17) is a sensitive, differential pressure gage which measures and indicates promptly the differential between the impact and the static air pressures surrounding an airplane at any moment of flight. The airspeed indicator consists primarily of a sensitive metallic diaphragm whose movements, resulting from the slightest difference in impact and static air pressures, are multiplied by means of a link, a rocking shaft, a sector with hairspring and pinion, and a tapered shaft to impart rotary motion to the pointer, which indicates the aircraft velocity on the dial face in terms of knots or m.p.h.
Sector Rocking shaft Long lever 50 Static connection ED 100 P'itot connection Diaphragm Restraining Hairspring spring Handstaff pinion FIGURE 12-17. Airspeed indicator.
Most airspeed indicators are marked to show speed limitations at a glance. The never-exceed velocity is designated by a red radial line. A yellow arc designates the cautionary range, and a white arc is used to indicate the range of permissible limits of flap operation.
The dial numbers used on different airspeed indicators are indicative of the type of aircraft in which the diaphragms are mechanically resolved to indicate true airspeed in knots. A typical true airspeed indicator is designed to indicate true airspeed from 1,000 ft. below sea level to 50,000 ft. above sea level under free air temperature conditions from +40° to -60° C.
they are used; for example, an airspeed indicator with a range of zero to 160 knots is commonly used in many light aircraft. Other types, such as a 430- knot indicator, are used on larger and faster air- 6 craft.
Another type of airspeed indicator in use is the TRUE 2 maximum allowable airspeed indicator shown in figure 12-18. This indicator includes a maximum allowable needle, which shows a decrease in max- -5 AIOSPEED imum allowable airspeed with an increase in altitude. It operates from an extra diaphragm in the airspeed indicator which senses changes in altitude 100 K PTS 3 and measures this change on the face of the instrument. Its purpose is to indicate maximum allowable indicated airspeed at any altitude.
FIGURE 12-19. True airspeed indicator.
=5 2- 100 KNOTS FICURE 12-18. Maximum allowable airspeed indicator.
The type of airspeed indicator known as a true airspeed indicator is shown in figure 12-19. It uses an aneroid, a differential pressure diaphragm, and a bulb temperature diaphragm, which respond respectively to changes in barometric pressure, impact pressure, and free air temperature. The actions of Mach Indicator Machmeters indicate the ratio of aircraft speed to the speed of sound at the particular altitude and temperature existing at any time during flight.
Construction of a Mach indicator is much the same as that of an airspeed indicator. It will usually contain a differential pressure diaphragm which senses pitot-static pressure, and an aneroid diaphragm which senses static pressure. By mechanical means, changes in pressures are then displayed on the instrument face in terms of Mach numbers.
The Machmeter shown in figure 12-20A is designed to operate in the range of 0.3 to 1.0 Mach and at altitudes from zero to 50,000 feet. The Machmeter shown in figure 12-20B is designed to operate in the range of 0.5 to 1.5 at altitudes up to 50,000 feet.
Combined Airspeed/ Mach Indicator Combined airspeed/Mach indicators are provided for aircraft where instrument space is at a premium and it is desirable to present information on a combined indicator. These instruments show indicated 480
a desired speed. A combined airspeed/Mach indicator is shown in figure 12-21.
I.O .3 Airspeed 5014 8 index 4 Mach limit index 8 MACH NUMBER
• Index adjusting knob FICURE 12-21. Combined airspeed/Mach indicator.
MAINTENANCE OF PITOT-STATIC SYSTEMS The specific maintenance instructions for any pitot-static system are usually detailed in the applicable aircraft manufacturer's maintenance manual.
However, there are certain inspections, procedures, and precautions to be observed that apply to all
• \ 1.4 1.5 56 systems. inspected for security of mounting and evidence of Pitot tubes and their supporting masts should be MACH damage. Checks should also be made to ensure that
1.3 NUMBER 7 electrical connections are secure. The pitot pressure entry hole, drain holes, and static holes or ports should be inspected to ensure that they are unob-
1.2 structed. The size of the drain holes and static holes is aerodynamically critical. They must never be cleared of obstruction with tools likely to cause .9 enlargement or burring.
1.0 Heating elements should be checked for functioning by ensuring that the pitot tube begins to warm up when the heater is switched "on." If an ammeter or loadmeter is installed in the circuit, a current reading should be taken.
B The inspections to be carried out on the individ FIGURE 12-20. Machmeters.
airspeed, Mach, and limiting Mach by use of impact and static pressures and an altitude aneroid.
These combined units utilize a dual-pointed neede which shows airspeed on a fixed scale and Mach indication on a rotating scale. A knurled knob located on the lower portion of the instrument is provided to set a movable index marker to reference ual instruments are primarily concerned with security, visual defects, and proper functioning. The zero setting of pointers must also be checked. At the time of inspecting the altimeter, the barometric pressure scale should be set to read field barometric pressure. With this pressure set, the instrument should read zero within the tolerances specified for the type installed. No adjustment of any kind can be made, if the reading is not within limits, the instrument must be replaced.
Leak Testing Pitot-Static Systems Aircraft pitot-static systems must be tested for leaks after the installation of any component parts, when system malfunction is suspected, and at the periods specified in the Federal Aviation Regulations.
The method of leak testing and the type of equipment to use depends on the type of aircraft and its pitot-static system. In all cases, pressure and suction must be applied and released slowly to avoid damage to the instruments. The method of testing consists basically of applying pressure and suction to pressure heads and static vents respectively, using a leak tester and coupling adapters. The rate of leakage should be within the permissible tolerances prescribed for the system. Leak tests also provide a means of checking that the instruments connected to a system are functioning properly. However, a leak test does not serve as a calibration test.
Upon completion of the leak test, be sure that the system is returned to the normal flight configuration. If it was necessary to blank off various portions of a system, check to be sure that all blanking plugs, adapters, or pieces of adhesive tape have been removed.
TURN-AND-BANK INDICATOR The turn-and-bank indicator, figure 12-22, also referred to as the turn-and-slip or needle-and-ball indicator, shows the correct execution of a bank and turn and indicates the lateral attitude of the aircraft in level flight.
The turn needle is operated by a gyro, driven either by a vacuum, air pressure, or electricity.
The turn needle indicates the rate, in numben of degrees per second, at which an aircraft is turning about its vertical axis. It also provides information on the amount of bank. The gyro axis is horizontally mounted so that the gyro rotates up and away from the pilot. The gimbal around the gyro is pivoted fore and aft.
Gyroscopic precession causes the rotor to tilt when the aircraft is turned. Due to the direction of rotation, the gyro assembly tilts in the opposite direction from which the aircraft is turning. This prevents the rotor axis from becoming vertical to the earth's surface. The linkage between the gyro assembly and the turn needle, called the reversing mechanism, causes the needle to indicate the proper direction of turn.
Power for the electric gyro may be supplied from either an a.c. or d.c. source.
The principal value of the electric gyro in light aircraft is its safety factor. In single-engine aircraft equipped with vacuum-driven attitude and heading indicators, the turn needle is commonly operated by an electric gyro. In the event of vacuum system failure and loss of two gyro instruments, the pilot still has a reliable standby instrument for emergency operation. Operated on current directly from the battery, the electric turn indicator is reliable as long as current is available, regardless of generator or vacuum system malfunction. In the electric instrument, the gyro is a small electric motor and flywheel. Otherwise both electric and vacuum-driven turn-needles are de- Two minute turn indicator Four minute turn indicator FIcURI: 12-22. Two types of turn-and-bank indicators.
signed to use the same gyroscopic principle of precession.
Power for the suction-driven turn needle is regulated by a restrictor installed between the main suction line and the instrument to produce a desired suction and rotor speed. Since the needle measures the force of precession, excessivly high or low vacuum results in unreliable turn-needle operation. For a specific rate of turn, low vacuum produces less than normal rotor speed and, therefore, less needle deflection for this specinc rate ot turn. The reverse is true for the condition of high vacuum.
Of the two types of turn needles shown in figure 12-22, the 2-min. turn indicator is the older. If the instrument is accurately calibrated; a single needle-width deflection on the 2-min. indicator means that the aircraft is turning at 3° per sec., or standard (2 min. for a 360° turn). On the 4-min. indicator, a single needle-width deflection shows when the aircraft is turning at 1½° per sec., or half standard rate (4 min. for a 360° turn). The 4-min, turn indicator was developed especially for high-speed aircraft.
The slip indicator (ball) part of the instrument is a simple inclinometer consisting of a sealed, curved glass tube containing kerosene and a black agate or a common steel ball bearing, which is free to move inside the tube. The fluid provides a damping action, ensuring smooth and easy movement of the ball. The tube is curved so that in a horizontal position the ball tends to seek the lowest point. A small projection on the left end of the tube contains a bubble of air which compensates for expansion of the fluid during changes in temperature. Two strands of wire wound around the glass tube fasten the tube to the instrument case and also serve as reference markers to indicate the correct position of the ball in the tube. During coordinated straight-and-level fight, the force of gravity causes the ball to rest in the lowest part of the tube, centered between the reference wires.
Maintenance Practices for Turn-and-Bank Indicators Errors in turn needle indications are usually due to insufficient or excessive rotor speed or inaccurate adjustment of the calibrating spring. There is no practical operational test or checkout of this instrument, other than visually noting that the indicator pointer and the ball are centered.
SYNCHRO-TYPE REMOTE INDICATING INSTRU- MENTS A synchro system is an electrical system used for transmitting information from one point to another.
Most position-indicating instruments are designed around a synchro system. The word "synchro" is a shortened form of synchronous, and refers to any one of a number of electrical devices capable of measuring and indicating angular deflection. Synchro systems are used as remote position indicators for landing gear and flap systems, in autopilot systems, in radar systems, and many other remoteindicating applications. There are different types of synchro systems. The three most cominon are: (1) Autosyn, (2) selsyn, and (3) magnesyn. These systems are similar in construction and all operate on identical electrical and mechanical principles.
D.C. Selsyn Systems The d.c. selsyn system is a widely used electrical method of indicating a remote mechanical condition. Specifically, d.c. selsyn systems can be used to show the movement and position of retractable landing gear, wing flaps, cowl flaps, oil cooler doors, or similar movable parts of the aircraft.
A selsyn system consists of a transmitter, an indi- 50° c Indicator Transmittter FIcURE 12-23. Schematic diagram of a d.c. selsyn system.
cator and connecting wires. The voltage required to operate the selsyn system is supplied from the aircraft's electrical system.
A selsyn system is shown schematically in figure
12-23. The transmitter consists of a circular resistance winding and a rotatable contact arm. The rotatable contact arm turns on a shaft in the center of the resistance winding. The two ends of the arm, or brushes, always touch the winding on opposite sides. The shaft to which the contact arm is fastened protrudes through the end of the transmitter housing and is attached to the unit (flaps, landing gear, etc.) whose position is to be transmitted. The transmitter is usually connected to the unit through a mechanical linkage. As the unit moves, it causes the transmitter shaft to turn. Thus, the arm can be turned so that voltage can be applied at any two points around the circumference of the winding.
As the voltage at the transmitter taps is varied, the distribution of currents in the indicator coils varies and the direction of the resultant magnetic field across the indicator is changed. The magnetic field across the indicating element corresponds in position to the moving arm in the transmitter.
Whenever the magnetic field changes direction, the polarized motor turns and aligns itself with the new position of the field. The rotor thus indicates the position of the transmitter arm.
When the d.c. selsyn system is used to indicate the position of landing gear, an additional circuit is connected to the transmitter winding, which acts as a lock-switch circuit. The purpose of this circuit is to show when the landing gear is up and locked, or down and locked. Lock switches are shown connected into a three-wire system in figure 12-24.
A resistor is connected between one of the taps of the transmitter at one end and to the individual lock switches at the other end. When either lock switch is closed, the resistance is added into the transmit- Rotatable contact Resisto Resistor with switches Lock taps cqually spaced Transmitter Indicating element FIcURE 12-24. A double-lock switch in a three-wire selsyn system.
ter circuit to cause an unbalance in one section of the transmitter winding. This unbalance causes the current flowing through one of the indicator coils to change. The resultant movement in the indicator pointer shows that the lock switch has been closed.
The lock switch is mechanically connected to the landing gear up- or down-locks, and when the landing gear locks either up or down, it closes the lock switch connected to the selsyn transmitter. This lock ing of the landing gear is repeated on the indicator.
Magnesyn System The magnesyn system is an electrical self-synchronous device used to transmit the direction of a magnetic field from one coil to another. The magne.
syn position system is essentially a method of measuring the extent of the movement of such elements as the wing and cowl flaps, trim tabs, landing gear, or other control surfaces. The two main units of the system are the transmitter and the indicator (figure 12-25).
A.C. power Junction box Connection to control surface through direct linkage Magnesyn Magnesyn position position transmitter indicator FIcURE 12-25. Magnesyn position-indicating system.
In a magnesyn transmitter a soft iron ring is placed around a permanent magnet so that most of the magnet's lines of force pass within the ring.
This circular core of magnetic material is provided with a single continuous electrical winding of fine wire. Figure 12-26 shows an electrical wiring schematic of a magnesyn system. The circular core of magnetic material and the winding are the essential components of the magnesyn stator. The rotor consists of the permanent magnet.
The movement of the control surface of the aircraft causes a proportional movement of the transmitter shaft. This in turn causes a rotary displacement of the magnet. Varying voltages are set up in 484