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
-1/3 28 volt: na gue 400 Hz power supply Transmitting magnesyn Indicating magnesyn FIGURE 12-26. Magnesyn system.
the magnesyn stator, depending on the position of the magnet. The voltage is transmitted to a magnesyn indicator which indicates on a dial the values received from the transmitter. The indicator consists essentially of a magnesyn, a graduated dial, and a pointer. The pointer is attached to the shaft and the shaft is attached to the magnet; thus, movement of the magnet causes movement of the pointer.
REMOTE-INDICATING FUEL AND OIL PRESSURE GAGES Fuel and oil pressure indications can be conveniently obtained through use of the various synchro systems. The type of synchro system used may be the same for either fuel or oil pressure measurement; however, an oil system transmitter is usually not interchangeable with a fuel system transmitter.
A typical oil pressure indicating system is shown in figure 12-27. A change in oil pressure introduced into the synchro transmitter causes an electri Oil pressure indicator (instrument. panel) cal signal to be transmitted through the interconnecting wiring to the synchro receiver. This signal causes the receiver rotor and the indicator pointer to move a distance proportional to the amount of pressure exerted by the oil.
Most oil pressure transmitters are composed of two main parts, a bellows mechanism for measuring pressure and a synchro assembly. The pressure of the oil causes linear displacement of the synchro rotor. The amount of displacement is proportional to the pressure, and varying voltages are set up in the synchro stator. These voltages are transmitted to the synchro indicator.
In some installations, dual indicators are used to obtain indications from two sources. On some aircraft, both oil and fuel pressure transmitters are joined through a junction and operate a synchro oil and fuel pressure indicator (dual side-by-side), thus combining both gages in one case.
CAPACITOR-TYPE FUEL QUANTITY SYSTEM The capacitor-type fuel quantity system is an electronic fuel measuring device that accurately determines the weight of the fuel in the tanks of an aircraft. The basic components of the system are an indicator, a tank probe, a bridge unit, and an amplifer. In some systems the bridge unit and amplifier are one unit, mounted in the same box. More recent systems have been designed with the bridge unit and a transistorized amplifer built into the instrument case.
To engine oil pump Oil pressure transmitter (engine) Stator Stator Pressure connector Rotor Rotor Diaphragm Indicator pointer Vent - Open to atmosphere A.C. Power FICURE 12-27. Oil pressure synchro system.
The fuel quantity indicator shown in figure 12-28 is a sealed, self balancing, motor-driven instrument containing a motor, pointer assembly, transistorized amplifier, bridge circuit, and adjustment potentiometers. A change in the fuel quantity of a tank causes a change in the capacitance of the tank unit. The tank unit is one arm of a capacitance bridge circuit. The voltage signal resulting from the unbalanced bridge is amplified by a phase-sensitive amplifier in the power unit. This signal energizes one winding of a two-phase induction motor in the indicator. The induction motor drives the wiper or a rebalancing potentiometer in the proper direction to balance the bridge, and at the same time positions an indicator pointer to show the quantity of fuel remaining in the tank.
bluend fuel FIGURE 12-28. Indicator and probe of a capacitor type fuel quantity system.
FICURE 12-29. Simplified capacitance-tank circuit.
A simplified capacitance bridge circuit is shown in figure 12-30. The fuel tank capacitor and a fixed reference capacitor are connected in series across a transformer secondary winding. A voltmeter is connected from the center of the transformer winding to a point between the two capacitors. If the two capacitances are equal, the voltage drop across them will be equal, and the voltage between the center tap and point P will be zero. As the fuel quantity increases, the capacitance of the tank unit increases, causing more current to flow in the tank unit leg of the bridge circuit. This will cause a voltage to exist across the voltmeter that is in phase with the voltage applied to the transformer. If the quantity of fuel in the tank decreases, there will be a smaller flow of current in the tank unit leg of the bridge.
A simplified version of a tank unit is shown in figure 12-29. The capacitance of a capacitor depends on three factors: (1) The area of the plates, Fuel tank
(2) the distance between the plates, and (3) the capacitor dielectric constant of the material between the plates. The only variable factor in the tank unit is the dielectric of the material between the plates.
When the tank is full, the dielectric material is all 115 v.
fuel. Its dielectric constant is about 2.07 at 0° C., 400 Hz compared to a dielectric constant of 1 for air. When the tank is half full, there is air between the upper half of the plates and fuel between the lower half.
Thus, the capacitor has less capacitance than it had when the tank was full. When the tank is empty, there is only air between the plates; consequently, Reference capacitor the capacitance is still less. Any change in fuel quantity between full and empty will produce a corresponding change in capacitance. FIGURE 12-30. Simplified capacitance bridge circuit.
The voltage across the voltmeter will now be out of phase with the voltage applied to the transformer.
In an actual capacitor type fuel gage, the input to a two-stage amplifier is connected in place of the voltmeter. It amplifies the signal resulting from an unbalance in the bridge circuit. The output of the amplifer energizes a winding of the two-phase indicator motor. The other motor winding, called the line phase winding, is constantly energized by the same voltage that is applied to the transformer in the bridge circuit, but its phase is shifted 90° by a series capacitor. As a result, the indicator motor is phase sensitive; that is, it will operate in either direction, depending on whether the tank unit capacitance is increasing or decreasing.
As the tank unit capacitance increases or decreases because of a change in fuel quantity, it is necessary to readjust the bridge circuit to a balanced condition so the indicator motor will not continue to change the position of the indicating needle. This is accomplished by a balancing potentiometer connected across one-half of the transformer secondary, as shown in figure 12-31. The indicator motor drives this potentiometer wiper in the direction necessary to maintain continuous balance in the bridge.
The circuit shown in figure 12-31 is a self-balancing bridge circuit. An "empty" calibrating potentiometer and a "full" calibrating potentiometer are connected across portions of the transformer secondary winding at opposite ends of the winding.
These potentiometers may be adjusted to balance the bridge voltages over the entire empty-to-full capacitance range of a specific system.
In some installations where the indicator shows the contents of only one tank and where the tank is fairly symmetrical, one unit is sufficient. However, for increased accuracy in peculiarly shaped fuel tanks, two or more units are connected in parallel to minimize the effects of changes in aircraft attitude and sloshing of fuel in the tanks.
ANGLE-OF-ATTACK INDICATOR The angle-of-attack indicating system detects the local angle of attack of the aircraft from a point on the side of the fuselage and furnishes reference Empty calibrating potentiometer Unshielded lead Shielded lead Test SW 115 v.
400 Hz Amplifier Tank unit = Amplifier output Motor Gear train Line phase Empty Scale calibrated in Ibs.
Balancing Full potentiometer
- Full calibrating potentiometer FICURE 12-31. Self-balancing bridge circuit.
information for the control and actuation of other units and systems in the aircraft. Signals are provided to operate an angle-of-attack indicator (figure 12-32), located on the instrument panel, where a continuous visual indication of the local angel of attack is displayed.
ted through two separate air passages to separate compartments in a paddle chamber. Any differential pressure, caused by misalignment of the probe with respect to the direction of airflow, will cause the paddles to rotate. The moving paddles will rotate the probe, through mechanical linkage, until the pressure differential is zero. This occurs when the slots are symmetrical with the airstream direction.
Paddle Potentiometer Paddle chamber Separator Dial adjustment screw (B) Probe Orifice (A) ' Air passages Slots FIcURE 12-32. Angle-of-attack system. (A) Indicator (B) Transmitter. FIGURE 12-33. Airstream direction detector.
A typical angle-of-attack system provides electrical signals for the operation of a rudder pedal shaker, which warns the operator of an impending stall when the aircraft is approaching the critical stall angle of attack. Electrical switches are actuated at the angle-of-attack indicator at various preset angles of attack.
The angle-of-attack indicating system consists of an airstream direction detector (transmitter) (figure 12-32B), and an indicator located on the instrument panel. The airstream direction detector contains the sensing element which measures local airflow direction relative to the true angle of attack by determining the angular difference between local airflow and the fuselage reference plane. The sensing element operates in conjunction with a servodriven balanced bridge circuit which converts probe positions into electrical signals.
The operation of the angle-of-attack indicating system is based on detection of differential pressure at a point where the airstream is flowing in a direction that is not parallel to the true angle of attack of the aircraft. This differential pressure is caused by changes in airflow around the probe unit. The probe extends through the skin of the aircraft into the airstream.
The exposed end of the probe contains two parallel slots which detect the differential airflow pressure (figure 12-33). Air from the slots is transmit- Two electrically separate potentiometer wipers, rotating with the probe, provide signals for remote indications. Probe position, or rotation, is converted into an electrical signal by one of the potentiometers which is the transmitter component of a self-balancing bridge circuit. When the angle of attack of the aircraft is changed and, subsequently, the position of the transmitter potentiometer is altered, an error voltage exists between the transmitter potentiometer and the receiver potentiometer the indicator. Current flows through a sensitive polarized relay to rotate a servomotor in the indicator.
The servomotor drives a receiver/potentiometer in the direction required to reduce the voltage and restore the circuit to an electrically balanced condition. The indicating pointer is attached to, and moves with, the receiver/potentiometer wiper arm to indicate on the dial the relative angle of attack.
TACHOMETERS The tachometer indicator is an instrument for indicating the speed of the crankshaft of a reciprocating engine and the speed of the main rotor assembly of a gas turbine engine.
The dials of tachometer indicators used with reciprocating engines are calibrated in r.p.m.; those used with turbine engines are calibrated in percentage of r.p.m. being used, based on the takeoff r.p.m.
Figure 12-34 shows a typical dial for each of the indicators just described.
10 20 ENT 30 100 40 90 80 60 (A) (B) FIctRE 12-34. Tachometer. (A) Reciprocating engine type. (B) Turbine engine type.
There are two types of tachometer systems in wide use today: (1) The mechanical indicating system, and (2) the electrical indicating system.
Mechanical Indicating Systems Mechanical indicating systems consist of an indicator connected to the engine by a flexible drive shaft. The indicator contains a flyweight assembly coupled to a gear mechanism that drives a pointer.
As the drive shaft rotates, centrifugal force acts on the flyweights assembly and moves them to an angular position. This angular position varies with the
r.p.m. of the engine. Movement of the flyweights is transmitted through the gear mechanism to the pointer. The pointer rotates to indicate the r.p.m. of the engine on the tachometer indicator.
Electric Indicating Systems A number of different types and sizes of tachometer generators and indicators are used in aircraft electrical tachometer systems. Generally, the varous types of tachometer indicators and generators operate on the same basic principle. Thus, the system described will be representative of most electrical tachometer systems; the manufacturer's instructions should always be consulted for details of a specific tachometer system.
The typical tachometer system (figure 12-35) is a three-phase a.c. generator coupled to the aircraft engine, and connected electrically to an indicator mounted on the instrument panel. These two units are connected by a current-carrying cable. The generator transmits three-phase power to the synchronous motor in the indicator. The frequency of the transmitted power is proportional to the engine speed. Through use of the magnetic drag principle, the indicator furnishes an accurate indication of engine speed.
Tachometer generators are small compact units, generally availble in three types: (1) The pad, (2) the swivel-nut, and (3) the screw type. These names are derived from the kind of mounting used in attaching the generator to the engine. The pad-type tachometer generator (figure 12-36A) is constructed with an end shield designed to permit attachment of the generator to a flat plate on the engine frame, or accessory reduction gearbox, with four bolts. The swivel-nut tachometer generator is constructed with a mounting nut which is free to turn in respect to the rest of the instrument. This type of generator can be held stationary while the mounting nut is screwed into place. The screw-type tachometer generator (figure 12-36B) is constructed with a mounting nut inserted in one of the generator end shields. The mounting nut is a rigid part of the instrument, and the whole generator must be turned to screw the nut onto its mating threads.
The dual tachometer consists of tvo tachometer indicator units housed in a single case. The indicator pointers show simultaneously on a single dial the r.p.m. of two engines. Some tachometer indicators are equipped with a flight-hour meter dial, usually located in the lower center area of the dial face, just below the pointer pivot.