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Engine instruments are mounted on the instrument panel in the pilot's compartment. They are usually grouped together. They aid the pilot in monitoring the engine, rotor, or propeller RPM pressures temperatures and general engine performance. Refer to the engine's maintenance manual for maintenance, operational limitations, and color code markings. Instruments today read in vertical scales, digital readouts, and in some cases heads up display (HUD). The HUD system projects instrument images on the windshield in front of the pilots. In this way, pilots do not have to divert their eyes from the view in front of and around them. The vertical scale offers similar safety conditions. The pilots have the ability to quickly glance at the instrument for a reading. This chapter will discuss the most commonly used instruments and the newer vertical scale (VIDS).


The VIDS (Figure 9-1) consists of a vertical strip central display unit (CDU), two vertical strip pilot display units (PDU), and two signal data converters (SDC). These readings are shown by ascending and descending columns of multicolored lights (red yellow, and green) measured against vertical scales. The scales operate as segments and light in normal progression. They remain on as the received signal level increases. They go off in normal progression as the received signal level decreases. When scales with red-and amber-coded segments below green-coded receive a signal level of zero (bottom scale), the segments light in normal progression and remain on. When the first segment above the red or amber range goes on, all red- or amber-coded segments go off. These segments remain off until the received signal level indicates a reading at or within the red or amber range. At that time all red-or amber-coded segments on the scale display goon or off in normal progression. For an in-creasing indication when a scale with side arrows lights, the corresponding side arrow also lights. As the segments go on, the corresponding side arrows also go on, one at a time. Only the side arrow associated with the highest percent indication of the corresponding scale will be on. For a decreasing indication, scales with side arrows will operate in the same manner. Only the side arrow associated with the highest percent indication of the corresponding scale will be on. The CDU and PDUs contain photocells that automatically adjust indicator lighting to ambient light. If either photocell fails, the lights on the vertical scales or the PDUs and the CDU go off. The DIM knob on the CDU has an override capability which allows the pilot to manually set the display light level.


The tachometer provides the pilot with a continuous indication of engine RPM. A variety or combination of systems may be used on gas turbine engines. Gas producer to gas generator tachometers, turbine and rotor tachometers, and N-I and N-II tachometer are some of the tachometer systems used (Figure 9-2).

Engine rotor RPM can be sensed by a mechanically driven tachometer generator, mechanically driven permanent magnet, or a pulse pickup which senses passing compressor, fan blades, or gear teeth. The output or signal from any of the sensors is directed to an appropriate indicator in the cockpit. The indicator is calibrated to read directly in percent RPM. Dual axial-flow compressor engines are usually provided with two tachometers: one indicates low-pressure compressor speed (Nl), and the other indicates high-pressure compressor speed (N2). For the dual-axial and centrifugal compressors the second tachometer indicates power turbine speed (N2).

For most axial-flow compressor engines, the tachometer monitors RPM during engine start and indicates overspeed condition should one occur. Use of the tachometer for setting thrust is not recommended on axial-flow compressor engines. The low-pressure compressor (Nl) tachometer on engine pressure ratio (EPR)-controlled dual-compressor engines may be used as an approximate reference to set engine thrust in transient and certain other flight conditions. When the N1 tachometer sets engine thrust, the thrust setting should be more accurately adjusted by engine pressure ratio as soon as possible. On single-compressor, axial-flow engines, engine speed should not be used primarily to set or check engine thrust.

The General Electric T-700/701 system eliminates tach generators. Two Np sensors are located in the exhaust frame. One sensor extends through the 1:30 o'clock strut and the other through the 10:30 o'clock strut. The power turbine shaft is equipped with two pairs of teeth which induce electrical pulses in the Np sensors. These teeth permit measurement of the torsion or twist of the shaft, which is proportional to output torque. The sensors are identical and interchangeable but serve different functions.

The left-hand (10:30 o'clock) sensor provides an Np signal to the ECU. This signal is used in the Np governing circuitry and for generation of a cockpit tachometer signal. The right-hand (1:30 o'clock) sensor feeds the ECU torque computation circuit and the Np overspeed protection system. The sensors contain a permanent magnet and wire coil. They produce a pulse of current each time a shaft or reference tooth passes (Figure 9-3).


Only a small part of the propulsive force produced by a turboprop is caused by jet thrust. Neither turbine discharge pressure nor engine pressure ratio is used to indicate power produced by the engine. Instead a torquemeter measures the power level that the engine develops on the ground and in flight. Consult the engine maintenance manual for a description of torquemeter system functions. In most systems, however, torquemeter oil pressure is used to actuate a torque indicating instrument in the aircraft. The torquemeter instrument portrays torquemeter oil pressure (which is proportional to engine power) in pounds per square inch. Some torquemeter instruments are calibrated to read poundfeet of torque. Some may read shaft horsepower directly (Figure 9-4).

One torquemeter indicating system sometimes called a torque pressure indicating system, contains a pressure-indicator-type instrument. It provides continuous reading of engine output shaft torque supplied by an electrical transmitter mounted on the engine inlet section. The transmitter is connected by hoses to a specialized oil pressure tap on the engine inlet housing and to a vent connection on the accessory drive gearbox. The 28-volt AC system operates the electrical circuit.

The Lycoming (T-55-L-712) torquemeter, a refined torque measuring system, comprises five components and interconnecting wiring. Two of these, the power supply and indicator, are airframe-supplied components. They may be replaced anytime without recalibration. The remaining three (power output shaft, head assembly, and junction box) are only replaced as a precalibrated set.

When torque is imposed on the engine power output shaft, tension and compression stresses change the magnetic reluctance of the shaft. This system measures the change of magnetic reluctance because of the torsionally induced forces (Figure 9-5).

The output shaft has welded to it a ferromagnetic sleeve which serves as a transformer core in the head assembly. A nonrotating transformer, the head assembly contains the primary coils which receive alternating current required to power the transformer from the power supply. It also contains six sets of secondary coils oriented at 90o to each and 45o to the shaft axis. They pick up induced current and register the differential in current on the dial face of the indicator. In operation, the head assembly surrounds the ferromagnetic sleeve on the shaft which acts as the transformer core.

The junction box acts as the system conditioner and includes key calibration controls. The potentiometers standardize the signal level for indicator display and balance out effects of magnetic nonhomogeneities. The box also rectifies the secondary voltages from the head assembly.

The cockpit-mounted indicator reads the engine-applied torque in percent. A network located within the indicator computes the algebraic difference between the two secondary-induced voltages coming from the junction box.

The power supply is an airframe-mounted unit which converts the airframe 28VDC signal to a constant 70VAC 2K Hz signal for system operation.

The General Electric T-700/701 torque-sensing system is a reference shaft that is pinned to the front end of the drive shaft and extends to the aft end. It is free to rotate relative to the drive shaft. The relative rotation is due to output torque. The resultant phase angle between the drive shaft teeth and reference teeth is electrically sensed by a pickup sensing the two teeth on the drive shaft plus two reference teeth. The electrical signal is conditioned in the electrical control unit, which provides a DC voltage proportional to torque (Figure 9-5). An intermediate power (1690 SHP), the output torque is 410 pound-feet, and the twist of the shaft is 7.4.


Turbine engines may be instrumented for exhaust gas temperature indication at locations before, between, or behind the turbine stages. Exhaust gas temperature is an engine operating limit and is used to monitor the mechanical integrity of the turbines. It also checks engine operating conditions. Actually, the temperature at the turbine inlet is the important consideration. This is the most critical of all engine variables. However, it is impractical to measure turbine inlet temperature in most engines. Consequently, thermocouples are inserted at the turbine discharge. This provides a relative indication of temperature at the inlet. The temperature at this point is much lower than that at the inlet. It enables the pilot to maintain surveillance over engine internal operation conditions.

Several thermocouples are usually used. They are spaced at intervals around the perimeter of the engine exhaust duct near the turbine exit. The exhaust gas temperature indicator in the aircraft shows the average temperature measured by the individual thermocouples. The readings of several thermocouples are usually obtained individually during ground engine maintenance by a selective switch. The spread between the lowest and the highest thermocouple reading is useful in maintenance. It serves to indicate the presence of hot or cold spots in the engine exhaust pattern which may indicate a problem inside the engine. The importance of exhaust gas temperature cannot be overemphasized. Two systems that are in use on Army helicopters follow.

Thermoelectrical Systems (T-55-L-11E/T-55-L-712)

The chromel-alumel, thermoelectric (MGT) measuring system is independent of all other engine electrical wiring (Figure 9-6). The engine components are five dual-probe segments; each probe is connected externally so each segment can be operationally checked independently and replaced individually. The 10 thermocouple probes protrude into the gas stream at the power turbine entrance (station 45). The probes react to variations in temperature by developing a proportional electromotive force across the chromel-alumel junction.

This potential difference results in meter deflections of the cockpit indicator. The cockpit indicator is calibrated to read temperature in degrees centigrade.

In addition, this system incorporates two buss-bars located on the aft fire shield. This allows for connection of the five MGT segments to the single MGT harness.

T-701 General Electric Thermocouple

The thermocouple harness is a one-piece assembly. It consists of seven single-immersion, equally spaced thermocouples for measuring power turbine inlet gas temperature (Figure 9-7). The thermoelements are made from special tolerance, oxidation-resistant, chromel-alumel wire. Each thermocouple junction is sealed within a Hastalloy X sheath. The thermoelement for the junction of each probe is continuous without joints or splices from the junction to the harness output connector junction box where all outputs are parallel. The harness output connector is hermetically sealed, has two alumel contacts, and has two chromel contacts. The average output of the seven probes provides the temperature signal to the ECU via the yellow cable. From the ECU, the signal is relayed to the cockpit TGT indicator and to the history recorder.


The indicator (Figure 9-8) in a typical exhaust gas temperature indicating system operates on electrical potential from an engine thermocouple harness. The probes are mounted in the aft section of the engine exhaust diffuser, The thermocouple converts heat into electricity. The exhaust gas temperature indicator (thermocouple-thermometer indicator) is actually a sensitive millivoltmeter calibrated in degrees centigrade. Its D'Arsonval movement is activated by an electrical force generated by its relative thermocouple. The indicator circuit is entirely independent of any other electrical power source. It includes a coil resistor which provides instrument calibration.


The fuel-flow-rate indicating system measures the rate of fuel flow consumed by the engine. A typical system consists of an indicator (dual indicator if a two engine system), a fuel-flow meter, and a fuel transmitter, which is an integral part of the fuel-flow meter located on the engine. The fuel-flow-rate indicator, driven by a differential autosyn, registers the rate of fuel flow to the engine in hundreds of pounds per hour. A four-digit subtracting counter on the indicator can be set to read the total pounds of fuel remaining in the aircraft. The fuel-flow-meter transmitter is installed on a mount on the engine. The flow meter consists of a metering chamber and a differential autosyn. As fuel travels from the main tank or valve manifold to the engine fuel control, it passes through the metering chamber, moving a pivoted vane. The pivot shaft of the vane is coupled to the rotor shaft of the differential autosyn. When fuel flows through the metering chambers, the differential autosyn rotates. This autosyn rotation, in turn, causes the indictor autosyn to rotate. This results in an indication of engine fuel consumption.


The fuel flow indicator shows fuel flow in pounds (or kilograms) per hour to the fuel nozzles (Figure 9-9). Fuel flow is monitored for in-flight fuel consumption, checking engine performance, and in-flight cruise control. The relationship of abnormal fuel flow to other instrument readings provides one of the best indicators for probable cause of engine malfunction.


A typical fuel pressure indicating system provides continuous reading of fuel pressure (psi) in the main fuel supply line from boost pumps in the tanks by an electrical transmitter. The transmitter is connected to a tap on the valve manifold. All fuel supply lines join the top electrical transmitter to deliver fuel to the engine through the fuel control inlet hose. Electricity is supplied to the transmitter by the 28-volt AC system.


Fuel system characteristics frequently make it advisable to monitor the fuel pump inlet pressure (Figure 9-10). In case of fuel flow stoppage in flight, the source of difficulty should be located quickly. This determines whether trouble developed in the engine or in the aircraft fuel system. Prompt corrective action may then be taken. In addition, the fuel pump inlet pressure indicates possible cavitation at the fuel pump inlet in flight. It will also show whether or not the fuel system is operating properly during engine ground checks.


A typical engine oil pressure indicating system provides continuous reading of engine oil pump pressure in psi to the indicator. This is provided by an electrical transmitter mounted on the engine inlet section. The transmitter is connected to the 28-volt AC electrical system and by a hose to a pressure tap on the engine oil falter housing.


To guard against engine failure resulting from inadequate lubrication and cooling of engine parts, oil supply to critical areas must be monitored (Figure 9-11). The oil pressure indicator shows the pressure relayed by the oil pressure transmitter. On most installations, the oil pressure transmitter takes breather pressure into consideration, relaying the true pressure drop across the oil jets in the oil system.


In a typical engine oil temperature indicating system, the indicator is electrically connected to the 28-volt DC system. A electrical resistance-type thermobulb installed in the engine oil pump housing measures the temperature of oil entering that unit. The temperature reading is transmitted to the indicator in degrees centigrade. Two dissembler metals heat to electricity.


The ability of the engine oil to lubricate and cool is a function of the oil temperature and the amount of oil supplied to the critical areas. An oil inlet temperature indicator is frequently provided to show the oil temperature as it enters the engine-bearing compartments. Oil inlet temperature also indicates proper operation of the engine oil cooler (Figure 9-12).


The air temperature indications (Figure 9-13) currently used in aircraft are free air temperature (FAT), outside air temperature (OAT), ram air temperature (RAT), total air temperature (TAT), and static air temperature (SAT). Regardless of which temperature is instrumented in a specific aircraft model, the flight manual shows how to use it, along with applicable charts or tables, to set the EPR values which provide rated thrust levels. The EPR setting varies with the thrust level desired and with the true TAT existing at the front of the engine (Tt2). Some aircraft have instrumentation which indicates Tt2 values that may be used without correction to determine EPR settings.


The way a pilot sets and monitors the thrust produced by the engine has been mentioned before (Figure 9-14). Thrust indication is discussed in detail below.

On some engines, engine RPM and exhaust gas temperature (EGT) are used to indicate and set thrust. On such engines, the pilot receives the full rated thrust of the engine for takeoff at 100 percent RPM and a specified EGT. The specified EGT at 100 percent RPM is established on a thrust-measuring ground test standby varying the exhaust nozzle area of the engine to achieve the desired EGT.

On centrifugal compressor engines, notably the military J48, thrust is indicated by RPM alone. Full rated thrust for takeoff is obtained when the tachometer reads 200 percent. The J48 has a fixed nozzle area which is established at manufacture. While there is an EGT limit for takeoff and for the other engine ratings, a J48 will normally operate at EGT below the applicable limit for the thrust rating used by the pilot. If the EGT reaches the allowable limit, the engine deteriorates or malfunctions.

Most afterburning and non-afterburning turbojets and turbofans with single or dual axial-flow compressors use engine ratio to measure engine thrust. EPR indicators compare total turbine discharge pressure to total air pressure entering the compressor. EPR then indicates the ratio of these pressures. Engines instrumented for EPR have a fixed exhaust nozzle area. Some military afterburning engines have two fixed areas. One is used for non-afterburning operation. A variable nozzle area is used for some afterburning engines, but it varies only while in afterburning. On both afterburning and nonafterburning engines, RPM and EGT may vary when the aircraft throttle is adjusted to obtain desired engine thrust.

Some military afterburning models have exhaust nozzles which are scheduled to vary the exhaust area when the engine is running. Consequently, these engines cannot be operated to EPR settings. They must be controlled by throttle position with various engine parameters checked to assure correct thrust output.

For engines with a fixed nozzle area, actual exhaust gas temperatures obtained during operation are usually below prescribed limits. It is permissible for an engine to operate at the temperature limit for any given thrust rating. However, an engine that does may have a problem which causes it to run abnormally hot.

With the exception noted in the use of the tachometer, engine RPM is considered a very inadequate parameter for setting and checking engine thrust on single and dual axial-flow compressor turbojet and turbofan engines having fixed exhaust nozzle areas. When RPM is used as the controlling variable on such engines, complications arise. The most important are--

  • RPM does not provide an accurate means of determining if the complete engine is functioning properly. High-pressure RPM on dual axial-flow compressor engines and the RPM of the whole compressor on single axial-flow compressor engines is governed by fuel control. For example, RPM alone will not enable an engine operator to detect a damaged or dirty compressor. RPM carefully used in conjunction with other engine variables such as fuel flow, exhaust gas temperature, and engine pressure ratio allow for detection.
  • RPM for any given thrust condition will vary slightly among individual engines, depending upon the engine trim speed. Engines are trimmed by a fuel control adjustment to produce full rated thrust at a fixed-throttle position on a standard day. The RPM variation must be taken into account when RPM is used to measure thrust being developed by the engine. This causes a complication which cannot be tolerated when precise thrust settings are necessary during flight.
  • On dual axial-flow compressor engines, RPM variation of one percent causes approximately four percent variation in thrust at the higher thrust settings for the low-pressure compressor rotor (N1) and five percent variation for the high-pressure compressor rotor (N2). One per-cent variation in turbine discharge pressure or engine pressure ratio results in only one and one-half percent variation in thrust. The five percent variation in thrust for one percent variation in RPM also holds true for single axial-flow compressor engines.
  • RPM does not vary in direct proportion to thrust produced by the engine over the entire thrust range.

For these reasons, either turbine discharge pressure or engine pressure ratio must be used as the engine variable to indicate thrust on axial-flow compressor engines with fixed area exhaust nozzles. The use of either is simpler undermost conditions and is more accurate.


For engines other than those with fully variable exhaust nozzles, turbine discharge pressure, engine pressure, or engine pressure ratio can be used with good results to indicate or set engine thrust. They vary proportionally to the thrust the engine is developing. Most turbojet and turbofan aircraft today are instrumented for engine pressure ratio. This is the parameter generally used to set or measure engine thrust during takeoff, climb, and cruise. For very accurate thrust measurement, such as during ground trimming of an engine, turbine discharge pressure is often used to measure thrust. In such cases, it is common practice to temporarily connect a turbine discharge pressure indicator to the engine.


In a typical engine pressure ratio indicating system, the indicator is a dual-synchro instrument. The system shows a constant reading of engine performance. This is done by computing the ratio between the gas generator discharge pressure and the inlet pressure of the engine. These ratios are then transmitted to an indicator (both indicators if a two-engine aircraft). Sample pressures are taken from engine gas producer or gas generator ports and from the pilot pressure system. The system includes a transducer. The transducer includes a mounting bracket and a transmitter unit. The transmitter unit contains a multicell, diaphram-actuated computer; gear train; two-phase motor; and transmitting synchros. The indicators are graduated from 1.0 to 2.5 EPR units with 0.1 EPR markings and a vernier dial graduated in 0.01 EPR markings. An adjustable pointer on the face of the indicator is set to the maximum EPR for the ambient temperature to indicate possible engine overspeed. The maximum EPR for given temperatures is listed on the instrument panel. The AC Power system supplies electrical power through the EPR circuit breaker.


This instrument indicates the internal engine pressure upstream of the jet nozzle, immediately aft of the last stage of the turbine (Pt5 to Pt7). It indicates pressure available across the nozzle to generate thrust. Turbine discharge pressure must be used in conjunction with T12 and P12.


This instrument (Figure 9-15) indicates the engine pressure ratio as a measure of thrust developed by the engine. This is the ratio of the turbine discharge total pressure to the equivalent of the compressor inlet total pressure (Pt5/Pt2 or Pt7/Pt2). Values for Pt2 must be corrected for inlet duct loss on the engine pressure ratio curves or charts by the aircraft manufacturer. Therefore, both for static (takeoff) and flight use, the actual value for Pt2 will vary among different aircraft types and models because of installation effects. However, the relation of Pt2 at the engine face to Pam plus Pr is equivalent to total pressure at or near the compressor inlet when airborne. It is not advisable to instrument the compressor inlet when airborne. It is not advisable to instrument the compressor inlet directly for Pt2. The Pt2 sensor for the pressure ratio indicator may be placed at some other location on the aircraft, preferably as near the engine air inlet as possible. After appropriate corrections are made to the in-flight charts in the flight of operation manual, any rated thrust or percent of rated thrust may be set with the aircraft throttle (as a function of the TAT or Tt2).

Pressure ratio between the pressure at the engine air inlet and the discharge pressure at the jet nozzle indicates thrust developed by the engine. Turbine discharge pressure alone is not an accurate indication of engine output. Compressor inlet pressure (Pt2) must be taken into account on curves or charts when only turbine discharge pressure is instrumented.

For static engine operation, this is usually accomplished by showing barometric pressure (corrected for inlet duct loss) rather than Pt2 values on the curves or charts. In flight, curves or tables usually show airspeed and altitude. This eliminates the need for actually delineating Pt2 values in the operating data. Engine pressure ratio indicators have the Pt2 value introduced into the system. This factor is automatically taken into account on the observed instrument reading. Except for an indicator to measure engine thrust, the above represents the minimum adequate instrumentation for control of the engine. Some installations may have additional instruments.

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