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.
Tachometer indicator Tachometer generator Magnetic drag cup Pointer Synchronous Three-phase AC
•notor FICURE 12-35. Schematic of a tachometer system.
A. Pad type B. Screw type FicURE 12-36. Tachometer generators.
Dual tachometers are also placed in the same case with a synchroscope for various purposes. One of these is the helicopter tachometer with synchroscope. This instrument shows simultaneously the speed of rotation of the engine crankshaft, the speed of rotation of the rotor shaft, and the slippage of the rotor due to malfunctioning of the clutch or excessive speed of the rotor when the clutch is disengaged in flight. The speed of both the rotor shaft and the engine shaft is indicated by a regular dual tachometer, and the slippage is indicated on a synchroscope (figure 12-37).
Tachometer Maintenance Tachometer indicators should be checked for loose glass, chipped scale markings, or loose pointers. The difference in indications between readings taken before and after lightly tapping the instrument should not exceed approximately +15 r.p.m.
СИ7 This value may vary, depending on the tolerance established by the indicator manufacturer. Both tachometer generator and indicator should be inspected for tightness of mechanical and electrical connections, security of mounting, and general condition. For detailed maintenance procedures, the manufacturer's instructions should always be consulted.
When an engine equipped with an electric tachometer is running at idle r.p.m. the tachometer indicator pointers may fluctuate and read low. This is an indication that the synchronous motor is not synchronized with the generator output. As the engine speed is increased the motor should synchronize and register the r.p.m. correctly. The r.p.m. at which synchronization occurs will vary with the design of the tachometer system.
If the instrument pointers oscillate at speeds above the synchronizing value, determine that the total oscillation does not exceed the allowable tolerance. If the oscillation exceeds the tolerance, determine if it is the instrument or one of the other components that is at fault.
Pointer oscillation can occur with a mechanical indicating system if the flexible drive is permitted to whip. The drive shaft should be secured at frequent intervals to prevent it from whipping.
When installing mechanical type indicators, be sure that the flexible drive has adequate clearance 15 8PM 20 10 ENGINE 20 RPM 25 50- 35 FASTE ROT. • FIGURE 12-37. Helicopter tachometer with synchroscope.
behind the panel. Any bends necessary to route the drive should not 'cause strain on the instrument when it is secured to the panel. Avoid sharp bends in the drive; an improperly installed drive can cause the indicator to fail to read, or to read incorrectly.
SYNCHROSCOPE The synchroscope is an instrument that indicates whether two (or more) engines are synchronized;
that is, whether they are operating at the same
r.p.m. The instrument consists of a small electric motor which receives electrical current from the tachometer generators of both engines. The synchroscope is designed so that current from the faster-running engine controls the direction in which the synchroscope motor rotates.
If both engines are operating at exactly the same speed, the synchroscope motor does not operate. If, however, one engine is operating faster than the other, its generator signal will cause the synchroscope motor to turn in a given direction. If the speed of the other engine then becomes greater than that of the first engine, the signal from its generator will then cause the synchroscope motor to reverse itself and turn in the opposite direction.
The motor of the synchroscope is connected by means of a shaft to a double-ended pointer on the dial of the instrument (figure 12-38). It is necessary to designate one of the two engines a master engine if the synchroscope indicators are to be useful. The dial readings, with leftward rotation of the pointer indicating "slow" and rightward motion in- FA FIGURE 12-38. Synchroscope dial, dicating "fast," then refer to the operation of the second engine in relation to the speed of the master engine.
For aircraft with more than two engines, additional synchroscopes are used. One engine is designated the master engine, and synchroscopes are connected between its tachometer and those of each of the other individual engines. On a complete installation of this kind, there will, of course, be one less instrument than there are engines, since the master engine is common to all the pairs.
One type of four-engine synchroscope is a special instrument that is actually three individual synchroscopes in one case (figure 12-39).
Rotor assembly Contact Balance ring disc Brush Stator FIGURE 12-39. Four-engine synchroscope.
The rotor of each is electrically connected to the tachometer generator of the engine designated as the master, while each stator is connected to one of the other engine tachometers. There are three hands, each indicating the relative speed of the number two, three, or four engine, as shown in figure 12-40.
The separate hands revolve clockwise when their respective engine is running faster than the master and counterclockwise when it is running slower.
Rotation of the hand begins as the speed difference reaches about 350 r.p.m., and as the engines approach synchronization the hand revolves at a ratio proportional to the speed difference.
TEMPERATURE INDICATORS Various temperature indications must be known in order for an aircraft to be operated properly. It is important that the temperature of the engine oil, carburetor mixture, inlet air, free air, engine cylinders, heater ducts, and exhaust gas temperature of 491
Tachometer Tachometer Generator Indicator Generator Indicator OB B Engine No. 2 Engine No. 3 When 3-wire system is required connect"C" terminal of all units. The "C" terminal is connected by internal grounds.
Synchroscope Tachometer Tachometer Generator Indicator Indicator Generator › B Engine No. 1 Engine No. 4 FICURE 12-40. Four-engine synchroscope schematic, turbine engines be known. Many other temperatures must also be known, but these are some of the more important. Different types of thermometers are used to collect and present this information.
Electrical Resistance Thermometer 50 Electrical resistance thermometers are used 100 widely in many types of aircraft to measure carburetor air, oil, and free air temperatures. °c 150 The principal parts of the electrical resistance -50 thermometer are the indicating instrument, the temperature-sensitive element (or bulb), and the connecting wires and plug connectors.
Oil temperature thermometers of the electrical resistance type have typical ranges of from -10° to + 120° C., or from -70° to + 150° C. Carburetor air and mixture thermometers may have a range of from -50° to +50° C., as do many free air thermometers.
A typical electrical resistance thermometer is FIcURE 12-41. Typical electrical resistance temperature shown in figure 12-41. Indicators are also available indicator.
in dual form for use in multi-engine aircraft. Most indicators are self-compensated for changes in cockpit temperature.
The electrical resistance thermometer operates on the principle of the change in the electrical resistance of most metals with changes in temperature. In most cases, the electrical resistance of a metal increases as the temperature rises. The resistance of some metals increases more than the resistance of others with a given rise in temperature. If a metallic resistor with a high temperature-resistant coefficient (a high rate of resistance rise for a given incresae in temperature) is subjected to a temperature to be measured, and a resistance indicator is connected to it, all the requirements for an electrical thermometer system are present.
The heat-sensitive resistor is the main element in the bulb. It is manufactured so that it has a definite resistance for each temperature value within its working range. The temperature-sensitive resistor element is a winding made of various alloys, such as nickel/manganese wire, in suitable insulating material. The resistor is protected by a closed-end metal tube attached to a threaded plug with a hexagon head (figure 12-42). The two ends of the winding are brazed or welded to an electrical receptacle designed to receive the prongs of the connector plug.
FIGURE 12-42. Two types of resistance thermometer bulb assemblies.
The electrical resistance indicator is a resistancemeasuring instrument. Its dial is calibrated in degrees of temperature instead of ohms and measures temperature by using a modified form of the Wheatstone-bridge circuit.
The Wheatstone-bridge meter operates on the principle of balancing one unknown resistor against other known resistances. A simplified form of a Wheatstone-bridge circuit is shown in figure 12-43.
14 volts + Indicator Cal. I Heat-sensitive element or bulb FIGURE 12-43. Wheatstone-bridge meter circuit.
Three equal values of resistances (A, B, and C, figure 12-43) are connected to a diamond-shaped bridge circuit with a resistance of unknown value (D).
The unknown resistance represents the resistance of the temperature bulb of the electrical resistance thermometer system.. A galvanometer calibrated to read in degrees is attached across the circuit at point X and Y.
When the temperature causes the resistance of the bulb to equal that of the other resistances, no potential difference exists between points X and Y in the circuit, and no current flows in the galvanometer leg of the circuit. If the temperature of the bulb changes, its resistance will also change, and the bridge becomes unbalanced, causing current to flow through the galvanometer in one direction or the other.
The dial of the galvanometer is calibrated in degrees of temperature, converting it to a temperature-measuring instrument. Most indicators are provided with a zero adjustment screw on the face of the instrument to set the pointer at a balance point (the position of the pointer when the bridge is balanced and no current flows through the meter).
Thermocouple Thermometer Indicators The cylinder temperature of most air-cooled reciprocating aircraft engines is measured by a thermometer which has its heat-sensitive element attached to some point on one of the cylinders (normally the hottest cylinder). In the case of turbojet engines, the exhaust temperature is measured by attaching thermocouples to the tailcone.
A thermocouple is a circuit or connection of two 493
unlike metals; such a circuit has two junctions. If cylinder head. Here again, the same metal is used in one of the junctions is heated to a higher tempera- the lead as in the part of the thermocouple to which ture than the other, an electromotive force is pro- it is connected.
duced in the circuit. By including a galvanometer in the circuit, this force can be measured. The hotter the high-temperature junction (hot junction) be- Thermocouple. gasket comes, the greater the electromotive force produced.
By calibrating the galvanometer dial in degrees it Back of becomes a thermometer. indicating instrument Engine A typical thermocouple thermometer system cylinder (figure 12-44) used to indicate engine temperature consists of a galvanometer indicator calibrated in Red Yellow degrees of centigrade, a thermocouple, and thermocouple leads. Red (A)
- Indicator Cold junction Thermocouple Yellow Engine cylinder wall Hot junction (B) Thermocouple leads FIGURE 12-45. Thermocouples: (A) Gasket type, (B) Bayonet type.
Black Connector Temperature to be measured FIcURE 12-44. Reciprocating engine cylinder head temperature thermocouple system.
Thermocouple leads are commonly made from iron and constantan, but copper and constantan or chromel and alumel are other combinations of dissimilar metals in use. Iron/constantan is used mostly in radial engines, and chromel/alumel is used in jet engines.
Thermocouple leads are designed to provide a definite amount of resistance in the thermocouple circuit. Thus, their length or cross-sectional size cannot be altered unless some compensation is made for the change in total resistance.
The hot junction of the thermocouple varies in shape depending on its application. Two common types are shown in figure 12-45; they are the gasket type and the bayonet type. In the gasket type, two rings of dissimilar metals are pressed together to form a spark plug gasket. Each lead that makes a connection back to the galvanometer must be made of the same metal as the part of the thermocouple to which it is connected. For example, a copper wire is connected to the copper ring and a constantan wire is connected to the constantan ring. The bayonet type thermocouple fits into a hole or well in the The cylinder chosen for installing the thermocouple is the one which runs the hottest under most operating conditions. The location of this cylinder varies with different engines.
The cold junction of the thermocouple circuit is inside the instrument case.
Since the electromotive force set up in the circuit varies with the difference in temperature between the "hot" and "cold" junctions, it is necessary to compensate the indicator mechanism for changes in cockpit temperature which affect the "cold" junction. This is accomplished by using a bimetallic spring connected to the indicator mechanism.
When the leads are disconnected from the indicator, the temperature of the cockpit area around the instrument panel can be read on the indicator dial.
This is because the bimetallic compensator spring continues to function as a thermometer.
Figure 12-46 shows the dials of two thermocouple temperature indicators.
Gas Temperature Indicating Systems EGT (exhaust gas temperature) is a critical variable of turbine engine operation. The EGT indicat.
ing system provides a visual temperature indication in the cockpit of the turbine exhaust gases as they leave the turbine unit. In certain turbine engines the temperature of the exhaust gases is measured at 494
• 2 150 250 3 2 50 NU, 350 °c X100
• FIGURE 12-46. Two types of thermocouple temperature indicators.
the entrance to the turbine unit. This is usually referred to as a TIT (turbine inlet temperature) indicating system. The principal disadvantages of this method are that the number of thermocouples required become greater and the environmental temperatures in which they must operate are increased.
A gas temperature thermocouple is mounted in a ceramic insulator and encased in a metal sheath;
the assembly forms a probe which projects into the exhaust gas stream. The thermocouple is made from chromel (a nickel/chromium alloy) and alumel (a nickel/aluminum alloy) . The hot junction protrudes into a space inside the sheath. The sheath has transfer holes in the end of it which allow the exhaust gases to flow across the hot junction.
Several thermocouples are used and are spaced at intervals around the perimeter of the engine turbine casing or exhaust duct. The thermocouples measure engine EGT in millivolts, and this voltage is applied to the amplifier in the cockpit indicator, where it is amplified and used to energize the servomotor which drives the indicator pointer.
A typical EGT thermocouple system is shown in figure 12-47.
The EGT indicator shown is a hermetically sealed unit and has provisions for a mating electrical connector plug. The instrument's scale ranges from O° C. to 1,200° C., with a vernier dial in the upper right-hand corner. A power "off" warning flag is located in the lower portion of the dial.
The TIT indicating system provides a visual indication at the instrument panel of the temperature of gases entering the turbine. On one type of turbine aircraft the temperature of each engine turbine inlet is measured by 18 dual-unit thermocouples installed in the turbine inlet casing. One set of these thermocouples is paralleled to transmit signals to the cockpit indicator. The other set of paralleled thermocouples provides temperature signals to the temperature datum control. Each circuit is electrically independent, providing dual system reliability.
The thermocouple assemblies are installed on pads around the turbine inlet case. Each thermocouple incorporates two electrically independent junctions within a sampling probe. The average voltage of the thermocouples at the thermocouple terminal blocks represents the TIT.
A schematic for the turbine inlet temperature system for one engine of a four-engine turbine aircraft is shown in figure 12-48. Circuits for the other three engines are identical to this system. The indicator contains a bridge circuit, a chopper circuit, a two-phase motor to drive the pointer, and a feed.
back potentiometer. Also included are a voltage reference circuit, an amplifier, a power "off" flag, a power supply, and an overtemperature warning light. Output of the amplifier energizes the variable field of the two-phase motor which positions the indicator main pointer and a digital indicator. The motor also drives the feedback potentiometer to pro- 495