Instruments for monitoring the operation of aircraft engines, classification. Aviation on-board instruments. Basic technical data

Engine types. Different types of aircraft use different types of engines. For example, light and medium-sized aircraft are equipped with gasoline internal combustion engines, which differ in the cooling method (air or water) and in the carburetion method (with a float or floatless carburetor); Heavy long-range aircraft use engines that run on heavy fuel, diesel engines, which provide greater fuel economy on long-distance flights.

For each of these engines there is a set of instruments that provide rational control of this engine and control of its operation (Fig. 11).

Due to the fact that stopping the engine in the air causes a forced landing of the aircraft, the most important role is played by instruments that monitor the operation of the engine as a whole and show the operating status of its individual units. Using these devices, the pilot also has the opportunity to correctly adjust the engine operating mode to maintain its strength and extend its service life.

In addition, the devices allow full use of engine power to achieve maximum flight speeds and maneuverability in air combat. Finally, with the help of instruments, you can set the most economical mode of engine operation, which saves fuel in flight.

Currently, due to the proliferation of jet engines, a new field of work has opened up for the designer of aircraft instruments. Built on completely different principles than internal combustion engines, jet engines require the use of new aircraft instrument designs.

Gasoline engine. The operation of this engine is based on the use of thermal energy released by gasoline during combustion in the engine cylinder. The energy of burned gasoline is converted into mechanical work in the air, creating a traction force that ensures the advancement of the aircraft.

For normal engine operation throughout the entire flight, an uninterrupted flow of fuel to the engine is necessary. Fuel is supplied to the engine cylinders by a group of units integrated into the engine power system. The fuel supply is located in gas tanks, usually placed inside the planes (aircraft wings).

Gasoline gauge indicates the amount of fuel in the tanks; The readings of this device are especially important for a pilot on a long flight.

Oxygen is required for the combustion of gasoline in the engine cylinders. Therefore, gasoline must enter the cylinders not in liquid form, but in an atomized state along with air, in the form of a so-called combustible mixture. The combustible mixture is prepared in the carburetor. A constant flow of gasoline to the carburetor is ensured by a gasoline pump, which continuously pumps gasoline from the tanks to the carburetor under a certain constant pressure, which is maintained by a pressure reducing valve. For gasoline engines with float carburetors, this pressure should be in the range of 0.2-0.35 atm, and if there is a floatless carburetor, 0.5-1 atm. With reduced pressure, the flow of fuel into the carburetor will be insufficient, which will cause interruptions in engine operation.

Fig. 11. Devices that control the operation of an aircraft engine.

The gasoline pressure gauge measures the pressure at which gasoline enters the carburetor. The readings of the gasoline meter and gasoline pressure gauge characterize the condition of the engine's gasoline supply system and the uninterrupted supply of fuel.

The composition of the combustible mixture prepared in the carburetor (i.e., the ratio of gasoline and air content) may be different. To determine the composition of the mixture, a gas analyzer is used, indicating the so-called excess air coefficient α. Small coefficient α. indicates that the amount of air in the mixture is not enough for complete combustion of gasoline; such a mixture is called “rich”. A high α coefficient indicates too much air, in which case the mixture is called "lean". Each engine operating mode requires its own mixture composition.

When moving, engine parts overcome frictional resistance, which entails wear of parts and loss of engine power. The engine lubrication system ensures a constant supply of oil to all rubbing parts to reduce friction and material wear. To ensure sufficient and uninterrupted lubrication, oil is supplied under pressure created by an oil pump. In modern aircraft engines, this pressure is maintained constant within 5-8 atm using a pressure reducing valve. The pressure in the lubrication system is shown by the oil pressure gauge.

Normal engine operation also largely depends on the temperature of the lubricating oil. At low temperatures (below 10-20° C), the viscosity of the oil increases greatly, its flow rate through the pipelines decreases, and it is especially difficult to supply oil through small cross-section channels to lubricate engine bearings.

Too high an oil temperature also has a bad effect on engine performance. At high temperatures, the viscosity of the oil decreases, it becomes fluid and is poorly retained in the gaps between the rubbing parts; at excessively high temperatures, the oil burns and its combustion products clog the rubbing surfaces. Thus, it is necessary to maintain the temperature of the lubricating oil within certain limits, for example, at the engine inlet 55-70 ° C, at the engine outlet 90-110 ° C. Short-term increases in oil temperature are acceptable within certain limits.

The oil temperature is measured oil thermometer. Changing the oil temperature in flight is achieved in two ways: either by changing the engine speed, or by changing the cooling conditions of the oil cooler. For example, when the oil temperature is too high, they either reduce the engine speed or open the oil cooler dampers, thereby increasing its airflow and, consequently, cooling.

When the combustible mixture burns, a large amount of heat is released, and the engine cylinders become very hot. At excessively high temperatures, the cylinders begin to deform, which can cause the engine pistons to seize. In order to maintain the temperature of the cylinders and pistons within acceptable limits, artificial cooling must be used. Depending on the method of heat removal, aircraft engines are divided into air-cooled and liquid-cooled engines.

With air cooling, the cylinders are blown by a stream of air. Cylinder temperatures on these engines are monitored by measuring the temperature of the cylinder heads with special thermometers. The permissible heating limit for engine cylinder heads is 240-250° C.

When the engine is liquid cooled, excess heat is removed by water or a special liquid that continuously washes the outer walls of the cylinders and transfers heat to the air in the radiator. In liquid-cooled engines, cylinder heating is judged indirectly - by the temperature of the liquid leaving the cylinder jackets. This temperature also has a permissible limit, which varies from engine to engine, depending on the design of the cooling system and the properties of the coolant.

With water cooling, the permissible water temperature at the outlet is approximately 85-90 ° C. To increase this limit, special liquids with a boiling point above 100 ° C are used, as well as cooling systems operating at elevated pressure. In these cases, the upper limit of the liquid temperature can be increased to 110-120 ° C. The temperature of the liquid leaving the cylinder jackets is measured water thermometer.

Not only overheating is dangerous for the engine, but also excessive cooling of the cylinders, since this reduces the combustion rate of the combustible mixture. The engine loses throttle response, i.e. the speed of transition to another operating mode. Loss of throttle response is especially dangerous during landing, when in some cases it is necessary to quickly increase the propeller speed in order not to lose speed.

The minimum permissible temperature of cylinder heads for air-cooled engines is about 120 ° C. The minimum temperature of the coolant at the engine outlet, as well as the temperature of the lubricating oil, must be regulated strictly within specified limits.

In flight, the temperature is controlled by changing the engine operating mode or opening the radiator shutters, which changes the cooling conditions. Some engines are equipped with automatic machines that maintain a given temperature of the cylinders or fluid by changing the cooling conditions. However, the use of automatic machines does not exclude the use of thermometers to monitor the serviceability of the automatic machines.

The thrust of the propeller, which propels the aircraft in the air, depends on the number of revolutions per minute of the propeller, and therefore on the number of revolutions per minute of the crankshaft. The motor shaft rotation speed shows tachometer. Most engines are equipped with an automatic machine that maintains a constant number of propeller revolutions by changing the angle of its blades (propeller pitch). In this case, the tachometer shows how well the propeller machine is working. During takeoff, to better utilize engine power, the propeller control is usually changed to increase the speed.

For complete combustion of gasoline, a certain amount of oxygen is needed. Oxygen is contained in the air sucked in by the engine. But at high altitudes, the air is very rarefied and when it is sucked into the cylinders, there is not enough oxygen to burn the fuel. Because of this, engine power decreases at altitude. It is necessary to equip high-altitude engines with a supercharger that compresses the air and supplies it at the required pressure to the cylinders.

This pressure is called boost pressure and is measured pressure and vacuum gauge. A number of engines have an automatic device that maintains constant boost pressure in the suction line of an aircraft engine. During takeoff, the boost pressure is increased by 100-200 mm Hg. Art., which is necessary to increase the power developed by the engine.

To maintain the required engine response, gasoline in the carburetor must evaporate at a sufficient speed. The rate of evaporation depends on the carburetor temperature, which is measured with a carburetor thermometer.

Heavy fuel engine. Recently, diesel engines have begun to be used on airplanes - engines powered by heavy fuel (kerosene, oil, gas oil). The main advantage of a diesel engine over a gasoline engine is lower fuel consumption.

The diesel power system is similar to the power system of a gasoline engine, which has a floatless carburetor with direct fuel injection. Fuel flows from the tank to the fuel pump, from where it is supplied under pressure of 2-4 atm to the fuel pump. The pump pumps fuel under a pressure of 500-1000 atm into the injectors, which inject fuel into the engine cylinders. The fuel is not ignited by an electric spark plug, as in gasoline engines, but ignites itself by heating the air. The air is heated to the required temperature due to its high degree of compression in the engine cylinders.

The amount of fuel in the tanks is measured by a fuel gauge, just like in a gasoline engine. To measure the pressure under which fuel is supplied by the pump to the fuel pump, a fuel pressure gauge is used, similar in design to a gasoline pressure gauge, but differing in the measurement range. Fuel pressure gauges used on diesel engines have a measurement range of up to 6 atm, and a pressure gauge for a gasoline engine with a float carburetor - up to 1 atm; on a gasoline engine with direct injection, a pressure gauge with a measurement range of 1.5-3 atm is used.

An instrument that measures instantaneous fuel consumption, the so-called fuel flow meter.

Diesel engine control is based on a different principle than gasoline engine control. In a carburetor engine, power is varied by changing the amount of combustible mixture supplied to the cylinders. To do this, open the throttle valve connected to the control handle (throttle sector). Changing the diesel power is achieved by changing the amount of fuel supplied through a special bypass device in the fuel pump. The pump control rack is connected to the handle of the fuel sector located in the pilot's cockpit.

In a diesel engine, the supplied fuel must be accurately dosed, and therefore, an accurate measurement of instantaneous fuel consumption is necessary. Naturally, a diesel engine does not need a gas analyzer and a carburetor thermometer. The lubrication and cooling systems of a diesel engine correspond to similar circuits of a gasoline engine. Accordingly, the same control and measuring instruments are used in diesel engines: oil pressure gauge, water and oil thermometers, cylinder head thermometer.

Diesel engines also use a supercharging system to maintain their power at a high level. Due to the absence of fuel detonation, a diesel engine allows higher boost pressure than a gasoline engine. Pressure and vacuum gauges used in diesel engines have a correspondingly higher measurement limit.

Contents of the article

AVIATION INSTRUMENTS, instrumentation that helps the pilot fly the aircraft. Depending on their purpose, aircraft on-board instruments are divided into flight and navigation devices, aircraft engine operation monitoring devices and signaling devices. Navigation systems and automatic machines free the pilot from the need to constantly monitor instrument readings. The group of flight and navigation instruments includes speed indicators, altimeters, variometers, attitude indicators, compasses and aircraft position indicators. Instruments that monitor the operation of aircraft engines include tachometers, pressure gauges, thermometers, fuel gauges, etc.

In modern on-board instruments, more and more information is displayed on a common indicator. A combined (multifunctional) indicator allows the pilot to cover all the indicators combined in it at a glance. Advances in electronics and computer technology have allowed for greater integration in cockpit instrument panel design and avionics. Fully integrated digital flight control systems and CRT displays give the pilot a better understanding of the aircraft's attitude and position than previously possible.

A new type of combined display - projection - gives the pilot the opportunity to project instrument readings onto the windshield of the aircraft, thereby combining them with the external panorama. This display system is used not only on military aircraft, but also on some civil aircraft.

FLIGHT AND NAVIGATION INSTRUMENTS

The combination of flight and navigation instruments provides a description of the condition of the aircraft and the necessary influences on the control elements. Such instruments include altitude, horizontal position, airspeed, vertical speed and altimeter indicators. For greater ease of use, the devices are grouped in a T-shape. Below we will briefly discuss each of the main devices.

Attitude indicator.

The attitude indicator is a gyroscopic device that provides the pilot with a picture of the outside world as a reference coordinate system. The attitude indicator has an artificial horizon line. The airplane symbol changes position relative to this line depending on how the airplane itself changes position relative to the real horizon. In the command attitude indicator, a conventional attitude indicator is combined with a flight control instrument. The command attitude indicator shows the aircraft's spatial position, pitch and roll angles, ground speed, speed deviation (true from the "reference" air speed, which is set manually or calculated by the flight control computer) and provides some navigation information. In modern aircraft, the command attitude indicator is part of the flight navigation instrument system, which consists of two pairs of color cathode ray tubes - two CRTs for each pilot. One CRT is a command attitude indicator, and the other is a planning navigation device ( see below). CRT screens display information about the spatial position and position of the aircraft in all phases of flight.

Planned navigation device.

The planned navigation device (PND) shows the course, deviation from the given course, the bearing of the radio navigation station and the distance to this station. PNP is a combined indicator that combines the functions of four indicators - heading indicator, radiomagnetic indicator, bearing and range indicators. An electronic POP with a built-in map indicator provides a color map image indicating the aircraft's true location relative to airports and ground-based radio navigation aids. Flight direction displays, turn calculations and desired flight paths provide the ability to judge the relationship between the aircraft's true position and the desired one. This allows the pilot to quickly and accurately adjust the flight path. The pilot can also display prevailing weather conditions on the map.

Airspeed indicator.

When an aircraft moves in the atmosphere, the oncoming air flow creates a high-speed pressure in a pitot tube mounted on the fuselage or on the wing. Airspeed is measured by comparing the velocity (dynamic) pressure with the static pressure. Under the influence of the difference between dynamic and static pressures, an elastic membrane bends, to which an arrow is connected, indicating the air speed in kilometers per hour on a scale. The airspeed indicator also shows the evolutionary speed, Mach number and maximum operational speed. A backup airspeed indicator is located on the central panel.

Variometer.

A variometer is necessary to maintain a constant rate of ascent or descent. Like an altimeter, a variometer is essentially a barometer. It indicates the rate of change in altitude by measuring static pressure. Electronic variometers are also available. Vertical speed is indicated in meters per minute.

Altimeter.

The altimeter determines the altitude above sea level based on the relationship between atmospheric pressure and altitude. This is, in essence, a barometer, calibrated not in pressure units, but in meters. Altimeter data can be represented in a variety of ways - using arrows, combinations of counters, drums and arrows, or through electronic devices that receive signals from air pressure sensors. See also BAROMETER.

NAVIGATION SYSTEMS AND AUTOMATICS

Airplanes are equipped with various navigation machines and systems that help the pilot navigate the aircraft along a given route and perform pre-landing maneuvers. Some such systems are completely autonomous; others require radio communication with ground navigation aids.

Electronic navigation systems.

There are a number of different electronic air navigation systems. Omnidirectional radio beacons are ground-based radio transmitters with a range of up to 150 km. They typically define airways, provide approach guidance, and serve as reference points for instrument approaches. The direction to the omnidirectional beacon is determined by an automatic on-board direction finder, the output of which is displayed by a bearing indicator arrow.

The main international means of radio navigation are VOR omnidirectional azimuthal radio beacons; their range reaches 250 km. Such radio beacons are used to determine the air route and for pre-landing maneuvers. VOR information is displayed on the PNP and rotating arrow indicators.

Rangefinding equipment (DME) determines the line-of-sight range within about 370 km from a ground-based radio beacon. Information is presented in digital form.

To work together with VOR beacons, instead of a DME transponder, ground equipment of the TACAN system is usually installed. The composite VORTAC system provides the ability to determine azimuth using the VOR omnidirectional beacon and range using the TACAN ranging channel.

An instrument landing system is a beacon system that provides precise guidance to an aircraft during final approach to a landing strip. Localization landing radio beacons (range of about 2 km) guide the aircraft to the center line of the landing strip; glide path beacons produce a radio beam directed at an angle of about 3° to the landing strip. The landing course and glide path angle are presented on the command attitude indicator and POP. The indices located on the side and bottom of the command attitude indicator show deviations from the glide path angle and the center line of the landing strip. The flight control system presents instrument landing system information via a crosshair on the command attitude indicator.

Omega and Laurent are radio navigation systems that, using a network of ground-based radio beacons, provide a global operating area. Both systems allow flights along any route chosen by the pilot. "Loran" is also used when landing without the use of precision approach equipment. The command attitude indicator, POP and other instruments show the aircraft's position, route and ground speed, as well as course, distance and estimated time of arrival for selected waypoints.

Inertial systems.

Flight data processing and display system (FMS).

The FMS system provides a continuous view of the flight path. It calculates airspeeds, altitudes, ascent and descent points that are most fuel efficient. In this case, the system uses flight plans stored in its memory, but also allows the pilot to change them and enter new ones through the computer display (FMC/CDU). The FMS system generates and displays flight, navigation and operational data; it also issues commands to the autopilot and flight director. In addition, it provides continuous automatic navigation from the moment of take-off to the moment of landing. FMS data is presented on the control panel, the command attitude indicator and the FMC/CDU computer display.

AIRCRAFT ENGINE OPERATION CONTROL DEVICES

Aircraft engine performance indicators are grouped in the center of the instrument panel. With their help, the pilot controls the operation of the engines, and also (in manual flight control mode) changes their operating parameters.

Numerous indicators and controls are required to monitor and control the hydraulic, electrical, fuel and maintenance systems. Indicators and controls, located either on the flight engineer's panel or on the hinged panel, are often located on a mimic diagram corresponding to the location of the actuators. Mnemonic indicators show the position of the landing gear, flaps and slats. The position of ailerons, stabilizers and spoilers may also be indicated.

ALARM DEVICES

In the event of malfunctions in the operation of engines or systems, or incorrect configuration or operating mode of the aircraft, warning, notification or advisory messages are generated for the crew. For this purpose, visual, audible and tactile signaling means are provided. Modern on-board systems can reduce the number of annoying alarms. The priority of the latter is determined by the degree of urgency. Electronic displays display text messages in the order and emphasis appropriate to their importance. Warning messages require immediate corrective action. Notification - require only immediate familiarization, and corrective actions - subsequently. Advisory messages contain information important to the crew. Warning and notification messages are usually made in both visual and audio form.

Warning alarm systems warn the crew of violations of normal aircraft operating conditions. For example, the stall warning system alerts the crew to such a threat by vibration of both control columns. The Ground Proximity Warning System provides voice warning messages. The wind shear warning system provides a visual signal and a voice message when an aircraft's route encounters a change in wind speed or direction that could cause a sudden decrease in airspeed. In addition, a pitch scale is displayed on the command attitude indicator, which allows the pilot to quickly determine the optimal angle of climb to restore the trajectory.

KEY TRENDS

“Mode S,” the proposed data link for air traffic control, allows air traffic controllers to transmit messages to pilots displayed on the aircraft's windshield. The Traffic Collision Alert System (TCAS) is an on-board system that provides information to the crew about required maneuvers. The TCAS system informs the crew about other aircraft appearing nearby. It then issues a warning priority message indicating the maneuvers required to avoid a collision.

The Global Positioning System (GPS), a military satellite navigation system that covers the entire globe, is now available to civilian users. By the end of the millennium, the Laurent, Omega, VOR/DME and VORTAC systems were almost completely replaced by satellite systems.

The Flight Status Monitor (FSM), an advanced combination of existing notification and warning systems, assists the crew in abnormal flight situations and system failures. The FSM monitor collects data from all on-board systems and issues text instructions to the crew to follow in emergency situations. In addition, he monitors and evaluates the effectiveness of the corrective measures taken.

Mechanical pressure gauges. They use pressure measurement methods in which the measured pressure forces are directly compared with the weight of a liquid column, a reference weight, or the forces of elastic sensing elements. Mechanical pressure gauges, designed on the basis of the first two methods, are used in stationary conditions or are used as reference gauges when checking and calibrating others. When implementing the third method of measuring pressure, membranes, membrane boxes, bellows and tubular springs are used as elastic sensitive elements (ESEs). Their deformation depends on the value of the measured pressure.

Rice. 12. Device of pressure and vacuum gauge

In the pressure-vacuum gauge (Fig. 12), manometric and barometric bellows 9 and 6 are used as a pressure gauge. Pressure r k which is measured is fed into the bellows 9 . Bellows 6 pressure is measured r a, equal to atmospheric. Under the influence of the pressure difference, the rod moves 8 , lever deflection 7 , thrust movement 2 , sector rotation 1 , tube rotation 5 and arrows 4 relative to scale 3 .

When measuring pressure with mechanical pressure gauges, methodological, instrumental and dynamic errors arise.

The methodological error appears due to changes in the absolute pressure of the environment.

Instrumental errors arise due to the presence of friction, play in the supports and hinges of moving elements, imbalance of the moving system, as well as changes in ambient temperature. The latter causes changes in the elastic modulus of the material from which the UCE is made, and in the geometric dimensions of the parts of the transmission mechanism. Reducing this error is achieved with the help of bimetallic temperature compensators and the selection of materials from which the UCEs are made.

Dynamic errors are caused by measurement lags, which depend on the parameters of the pipeline connecting the test object to the mechanical pressure gauge.

Electromechanical pressure gauges. In these pressure gauges, the forces of the measured pressure are converted into movement of the electrical elements, which affect the parameters of the measuring electrical circuits (resistance R, inductance L or capacity WITH). The pressure transducer is installed directly at the control object, which eliminates the need for long connecting pipelines, eliminates a number of errors, and simplifies installation and maintenance.

EDMU type pressure gauges. Electric remote pressure gauges of the unified EDMU type (Fig. 13) have the same structure and elements for all ranges of measured pressures, with the exception of the UChE and scale graduation. The electrical circuit diagram is shown below.

Rice. 13. Diagram of an EDMU type pressure gauge

Measured pressure r and fed to the UCHE, which is connected to the brush E 3 potentiometers IN 1 through the transmission mechanism. Resistance values Rx And Ry pressure transducer potentiometer, varying depending on pressure r and, form two arms of the bridge circuit. The other arms of the bridge circuit are resistors R 1 and R 2. Ratiometer frames L 1, L 2 and resistor R D constitute the measuring diagonal of the bridge. The common connection point of the frames is connected to a semi-diagonal consisting of resistors R 3 and R 4. They are designed to compensate for temperature errors caused by changes in the resistance of the ratiometer frames when the ambient temperature fluctuates. The ratiometer frames have the same number of turns, but different design dimensions. As a result, the inner frame has less resistance. To ensure symmetry of the circuit, an additional resistance is included in the circuit of the internal frame R D. When connected to the supply voltage circuit in case R x = R y the bridge circuit is symmetrical. Current flowing semidiagonally through resistors R 3 and R 4, branches into two equal currents I 1 and I 2 frames L 1i L 2(Fig. 14). If there is a violation of equality between Rx And Ry the symmetry in the circuit is broken, as a result of which the equality of currents is also violated. Currents I 1 and I 2, flowing through the frames of the ratiometer, create magnetic fields characterized by intensity vectors:

H 1 = I 1 w H 2 = I 2 w,

Where, w– the number of turns of each frame.

The moving magnet, on the axis of which the arrow is attached, is located in the direction of the vector

H = H 1 + H 2,

Where, H– vector of the resulting magnetic field strength.

Rice. 15. Kinematic diagram of the pressure transducer

Measured pressure r and supplied through a fitting 9 into the cavity of the pressure transducer. Under the influence r and the center of the membrane moves 8 , pusher 6 ,rocking chairs 5 , lever 3 , and brush holder 13. Spring 4 returns the lever to its original position when the pressure decreases r and.

Rice. 16. Design of the EDMU logometer

The design of the EDMU logometer (Fig. 16) consists of a moving magnet 2 and fixed frames 3 And 10 . Magnet 2 and arrow 5 attach to axle 9, the ends of which are inserted into the thrust bearings 6 . Copper body 1 A magnetic damper is used to dampen vibrations of the movable system of the ratiometer.

Fixed magnet 4 returns the instrument needle to the zero position when the supply voltage is turned off.

The errors introduced into the measurement circuit by a pressure sensor are similar to the errors of mechanical pressure gauges. Errors introduced by the electrical circuit and the indicator arise when the ambient temperature changes, when the moving system of the indicator is exposed to frictional forces, imbalance and backlash, as well as from magnetic hysteresis in the material of the screen and the moving magnet. The overall total error (± 4) and the presence of an unreliable sliding contract are disadvantages of this type of pressure gauges.

Pressure gauges type EM are differential type devices that measure the difference between two pressures (Fig. 17). Corrugated membranes are used as ECEs, the deformation of which is converted into an electrical value using a potentiometric transducer. The pointer is a four-frame logometer with a moving magnet.

Rice. 17. Diagram of an EM type pressure gauge

The ends of the potentiometer are short-circuited, so it is equivalent to a circular potentiometer. Each potentiometer section is connected to a corresponding tap of the ratiometer frame. A supply voltage of 27 V ± 10% is supplied to the brush of the potentiometric converter and the point connecting all the frames of the ratiometer. When the potentiometer brush moves under the influence of pressure forces, currents are redistributed within the ratiometer. Magnetic fields are created in them, characterized by intensity vectors. The moving magnet of a four-frame ratiometer is located in the direction of the tension vector N total magnetic field. Resistance R 1 and R 2 are used to adjust the width and uniformity of the scale. The use of such a scheme makes it possible to obtain, with small movements of the rigid center of the membrane and the potentiometer brush, large deflection angles of the pointer needle (the scale span reaches 270 0). This significantly increases the accuracy of pressure measurement, all other things being equal. Due to the symmetry of the device circuit, the indicator readings are not affected by changes in the supply voltage or frame resistance when the ambient temperature fluctuates. Total instrument error ± 3%. The main disadvantages of the EM type pressure gauge are the presence of a sliding contact and an increased number of connecting wires, which reduces the reliability of the device, increases its weight and complicates installation on board the aircraft.

Pressure gauges type DIM. The disadvantages of potentiometric transducers associated with wear of potentiometric transducers, associated with wear of the potentiometer, disruption of contacts during vibrations and fluctuations in the measured pressure, elevated temperatures, are eliminated in remote inductive pressure gauges of the DIM type (Fig. 18). This is ensured by the use of a differential inductive converter. Pressure gauges of this type are used for measuring pressure at elevated temperatures and significant high-frequency interference (up to 700 Hz). The electrical circuit diagram of the pressure gauge is shown below.

Rice. 18. Diagram of a DIM type pressure gauge

Either corrugated membranes or membrane boxes are used as UCE. The rigid moving center of the UCHE is connected to the armature of the inductive converter. Inductive converter coils L 1 and L 2 together with resistors R 1 and R 2 form a bridge circuit that operates on AC 36V 400Hz. The diagonal bridge circuit includes ratiometric indicator frames. When measuring pressure, the deformation of the UCE is transmitted to the armature, which changes the air gap in the magnetic circuits of the coils L 1i L 2. This causes changes in the inductance of the coils and leads to a redistribution of currents within the ratiometer. Since the logometer operates on direct current, diodes are introduced into the measuring circuit as rectifiers D 1 and D 2. The maximum errors of DIM type pressure gauges are ± 4%, the span of the indicator scale is 120 0.

Pressure alarms. They are designed to provide information about the presence of nominal or critical modes in power plant systems. ECU 1 of the pressure alarm controls the operation of contacts 4.5 that switch the electrical circuit (Fig. 19).

Rice. 19. Pressure alarm circuit

Pressure alarm 2 opens the electrical circuit using stops 3 and 6 when the pressure difference decreases Δр = р 2 – p 1 .

Pressure ratio meter type IOD. It is designed to control engine thrust in relation to pressure

π =р 2 / р 1

Where, p 1 – total pressure at the engine inlet;

p 2– pressure behind the engine turbine.

The device diagram (Fig. 20) consists of a pressure ratio sensor (PRS) and a pressure ratio indicator (PRI). It is a compensation-type measuring circuit, in contrast to direct conversion measuring circuits. DOD consists of: a working bellows 17, into the cavity of which pressure is applied r 2, aneroid 1, responsive to pressure changes r 1 supplied to the sensor housing; contact system 15, which serves to control the electric motor 13, through an amplifier 16, potentiometer 2, which fixes the deviation of the lever 18 .

Rice. 20. Diagram of a pressure ratio meter of the IOD type

The UOD consists of: amplifier 8; engine 10; a feedback mechanism, which includes a gearbox and potentiometer 12; indicator mechanism, including a running mechanism, scale 4, tape mechanism 3 and return spring 7. Lamps L1 And L2 illuminate the pointer scale.

When the operating mode of the engine changes, and therefore the pressure ratio changes, the movable contact of the contact system 15 located on the lever 18 will close with the upper or lower fixed contact, and the electric motor 13 will begin to rotate the aneroid, changing the angle of its inclination to the lever 18. When equilibrium is achieved, the given The forces of the bellows and aneroid open the contacts and the engine turns off. In this case, signals proportional to the pressure ratio are removed from potentiometer 2. It is included in the bridge measuring circuit of the pointer, containing a feedback potentiometer 12 and adjustable resistances 11. When the bridge is unbalanced, a voltage arises in the diagonal, which is amplified by the amplifier 8 and supplied to the electric motor 10 of the pointer, which balances the bridge circuit using potentiometric feedback 12 and moves the mechanism indicator with indicating tape 3. In this case, on scale 4 the value of the measured pressure ratio is indicated. In the event of a power failure or failure of the device elements, the tape returns to the lower mark of the scale by return spring 7. Adjustment resistors 11 allow you to adjust the span of the even-white border of the tape according to the pointer scale. By rotating the ratchet 6, the nut with arrow 5 moves along the scale to mark a preset value of the pressure ratio at the control point.

Thermal chip alarms. To promptly warn the crew about the occurrence of abnormalities in the operation of the bearing units of the middle and rear engine rotor supports, a housing with oil filters and thermal chip alarms (TCS) is installed in the lower part of the combustion chamber.

The system (Fig. 21) consists of the following main elements:

a) two thermal chip alarms 1, one of which is installed in the oil pumping line from the rear compressor rotor bearing, the other in the oil pumping line from the turbine rotor bearing;

b) a warning light located on the instrument panel in the cockpit.

There are two channels in the oil filter housing, one of which is connected to the cavity of the rear bearing of the compressor, the other to the cavity of the turbine bearing.

An oil filter 10 and a TCC 1 are installed in each channel, which with their flanges are jointly attached to the oil filter housing 11 with two bolts.

Rice. 21. Oil filter design

The oil filter housing 11, with its upper flange, is fastened with four bolts to the flange located on the lower stiffening rib of the combustion chamber housing. A paronite gasket is installed between the flanges.

In addition, two fittings are installed on the oil filter housing 11 to connect the housing channels with pipelines to the oil unit.

Each TSS consists of a sensor that signals the presence of steel shavings in the pumped-out oil, and a sensor for the maximum temperature of the air-oil mixture.

The steel chip presence sensor consists of a magnetic chip storage device, which consists of two permanent magnets 4 and 6, installed with an air gap opposite each other with different poles. The magnets are connected by wires 2 and 3 to the contacts of the plug connector of the thermal chip alarm. A plug connector is installed on the TCC body to connect it to the electrical circuits of the engine and aircraft.

The limit temperature sensor is located in the upper part of the housing 5 and consists of a housing 8, an insert 9 made of a low-melting alloy and contacts, one of which is the upper part of the magnet 6, and the other is the ring 7.

Insert 9 is placed inside cone 8 and supported by three equally spaced protrusions. Ring 7 is connected by wire 2 to magnet 4.

The operating principle of both the chip presence sensor and the temperature sensor is based on closing the negative circuit of the signal light of the thermal chip alarm system when chips appear or the temperature of the pumped-out air-oil mixture rises above the permissible value.

When metal shavings appear in one of the above-mentioned oil pumping lines, a closed network is formed between the magnets, since the gap between the magnets is filled with shavings.

As a result, the light on the instrument panel in the cockpit lights up for the presence of chips in the engine.

If the temperature of the air-oil mixture in the pumping line from the cavity of the rear bearing of the compressor rises above 180 0 C and the pump line from the cavity of the turbine bearing above 202 0 C, the low-melting inserts melt and connect the surface of the magnets 6 and rings 7 .A closed electrical circuit is formed, which turns on a light in the cockpit, signaling the presence of chips in the oil.

Conclusion: devices for monitoring the operation of aircraft power plants are designed to monitor the thrust and thermal conditions of aircraft engines, the state of lubrication, fuel reserve and consumption, and the operation of individual systems and units. These include instruments for measuring rotation speed, temperature, pressure, amount of fuel in tanks and fuel consumption. This group of devices also includes indicators for preset pressures in the fuel system and position indicators for the air intake cone, anti-surge flaps and fuel lever, which allow you to check the condition of the corresponding systems.

Aircraft engines, fuel and oil tanks, air system cylinders and other objects whose operation must be monitored during flight are located at a distance of several meters and even tens of meters from the cockpit, where aircraft control is concentrated. Therefore, all devices monitoring the operation of power plants must be remote.

Aircraft engines operate in intense thermal conditions close to the limit. Therefore, to thermometers used to monitor the thermal conditions of the engine and service systems. There is a requirement for increased measurement accuracy. Thus, at maximum values of measured temperatures, the error in measuring the temperature of turbojet gases should not exceed ± (0.5-1)%. The accuracy of temperature measurement in cooling systems of aircraft engines of all types is estimated at an acceptable error of ± (3-5)%.

Fuel pressure in gas turbine engines must be measured with an error of no more than ± 1.5% in the range of 0-10 kg/cm2 and ±4% in the range of 10-100 kg/cm2. The oil pressure measurement error should not exceed ± 4%.

Conclusion

Accurate measurement of the actual fuel supply on the aircraft and its instantaneous or total consumption is necessary to ensure flight safety and maintain optimal engine operating conditions. The error in measuring the amount of fuel when the aircraft is positioned in the flight line should not exceed 2-3% of the actual fuel supply and should not be more than ± 2.5%.

Preset pressure alarms must operate with an error not exceeding ± 5% of the nominal response pressure values.

Self-study questions

1. Controlled parameters of power plants, assemblies and systems of the aircraft.

2. The operating principle of a TEU type thermometer.

3. Operating principle of the temperature sensor.

4. Operating principle of TNV.

5. Operating principle of thermoelectric thermometers.

6. Operating principle of a magnetoelectric galvanometer

7. Devices for monitoring the condition of engine oil systems.

Literature

1. V.D. Konstantinov, I.G. Ufimtsev, N.V. Kozlov "Aviation equipment of aircraft" pp. 119-148.

2. Yu. P. Dobrolensky "Aviation equipment" pp. 82-88.

3. A.S. Tyrtychko, N.N. Tochilov, M.M. Nogas, V.M. Bluvshtein "Aviation equipment for helicopters" pp. 254-282.

4. V.V. Glukhov, I.M. Sindeev, M.M. Shemakhanov "Aviation and radio-electronic equipment of aircraft." pp. 46-76.

5. Lecture notes.

Related information.

Pressure gauges used on aircraft to measure fuel pressure, oil pressure, boost pressure (in piston engines), etc.

Membrane boxes or manometric tubular springs are used as sensitive elements in pressure gauges. Membrane boxes are a connection of two or more corrugated metal membranes in such a way that a cavity is formed between them, communicating with the measured pressure. Rigid centers are soldered to the centers of the membranes, connected through a transmission mechanism to the pressure gauge pointer.

The pressure tube is a hollow tube of oval cross-section smoothly bent along a circular arc, one end of which is rigidly fixed and communicates with the medium being measured, and the other is free to move under the influence of pressure forces. The free end of the tubular spring is also connected through a transmission mechanism to the pressure gauge needle.

Pressure gauges with diaphragm boxes are used to measure low pressures, and with a pressure spring - high pressures. For fire safety purposes, in order not to supply fuel to the device located on the dashboard, pressure gauges for measuring fuel pressure are equipped with special receivers (separators). Pressure gauges that measure oil pressure also have receivers installed that increase the accuracy of the instrument's readings. If oil pressure was directly supplied to the pressure spring, the instrument readings would be somewhat delayed due to the high viscosity of the oil. The pressure gauge receiver is a chamber divided into two sealed cavities by an inelastic diaphragm. Oil (gasoline) is supplied into one cavity, the pressure of which must be measured, and the second cavity, connected to the indicator, is filled with a liquid (toluene) having a low viscosity.

In piston engines, it is important to know the air or mixture pressure in the suction pipes. This parameter is measured by a device called a pressure and vacuum gauge (Fig. 129). The sensitive element of the pressure-vacuum gauge is an aneroid box. The measured pressure from the supercharger is supplied through a fitting into the device body. The deformation of the aneroid box under the influence of pressure is transmitted through the rigid center to the transmission mechanism and then to the pointer. To reduce the instrument reading error due to the influence of temperature, it is equipped with bimetallic compensators.

Currently, electric pressure gauges are widely used, characterized by high accuracy, simplicity of design, low weight and dimensions. The schematic diagram of an electric remote pressure gauge is shown in Fig. 130.

The sensitive element of electric pressure gauges is the pressure box, which is deformed under pressure. The movement of the rigid center of the pressure box is transmitted through the rod to the rocker, which controls the movement of the rheostat lever. When the rheostat brushes are in the middle and the resistances R3 and R4 are equal (the bridge circuit is balanced), equal currents flow through frames I and II, creating magnetic fields of equal strength around them. The pointer arrow takes the middle position.

When the resistance pressure changes, R3 and R4 form two variable arms of the bridge circuit. The bridge will become unbalanced and the magnet with the pressure indicator arrow will deviate.

Thermometers designed to measure the temperature of gases in gas turbine engines, the temperature of cylinder heads of piston engines, etc.

According to the operating principle of the sensitive elements, thermometers are divided into the following groups:

expansion thermometers based on the principle of thermal expansion of liquids and solids at constant external pressure (mercury, alcohol, bimetallic, etc.);

manometric thermometers based on the principle of measuring the pressure of a liquid, steam or gas inside a closed vessel of constant volume when the temperature changes; electric thermometers; thermoelectric thermometers, etc.

The last two types of thermometers are most widespread, since they are easier to use remotely.

To measure the temperature of cylinder heads and exhaust gas temperature, thermoelectric thermometers are used, which are characterized by their simple design and high sensitivity.

The operating principle of thermoelectric thermometers is based on the use of the thermoelectric effect, which consists in the fact that in a closed circuit composed of two dissimilar conductors and having two junctions, currents arise at different junction temperatures. By the magnitude of the thermal currents arising in the circuit, one can judge the value of the body (environment) temperature. Thermal currents are measured using a galvanometer connected to the circuit, the scale of which is graduated in °C.

The principle of operation of electric thermometers is based on the property of conductors or semiconductors to change electrical resistance depending on temperature. Thermometers of this type are assembled according to a bridge design, one of the arms of which is a heat-sensitive element. The heat-sensitive element is placed in the environment whose temperature needs to be measured.

A galvanometer or logometer is used as a temperature meter in electric thermometers. The resistance value of the heat-sensitive element is usually selected such that the bridge circuit would be balanced at a temperature equal to the average value of the temperature range of the measured medium. As the temperature increases (decreases) the bridge becomes unbalanced and the instrument's pointer arrow deviates in one direction or another.

Tachometers serve to measure the number of revolutions of the engine shaft. According to the principle of operation of the sensitive part, tachometers can be: centrifugal, electric, magnetic, friction, etc. One of the simplest and most widely used in aviation is remote magnetic tachometers.

Their operating principle is based on the phenomenon of inducing eddy currents in a metal body under the influence of the magnetic field of a rotating permanent magnet. The diagram of a magnetic tachometer is shown in Fig. 131.

The tachometer consists of a permanent magnet, a lightweight copper or aluminum disk and a pointer. When a permanent magnet rotates, eddy currents are induced in the copper disk and interact with the magnetic field of the magnet. The copper disk begins to rotate. The moment of interaction between the copper disk and the permanent magnet is proportional to the rotation speed. The copper disk is connected to the pointer and is held from rotation by a coil spring, the degree of twist of which is proportional to the number of revolutions of the magnet. Do the angle of deflection of the arrow can be used to judge the value of the revolutions.

In electric tachometers, a tachometer sensor - an alternating current generator - is connected to the engine shaft through a gearbox. The frequency of the current generated by the generator is proportional to the number of revolutions of the engine shaft. The current flows through the connecting wires to the tachometer pointer, causing rotation of a synchronous electric motor, on the axis of which a multi-pole permanent magnet is attached. A permanent magnet is placed in a metal cap (sensing element). When a permanent magnet rotates, eddy currents are induced in the copper cap, tending to entrain it. But the rotation of the cap is counteracted by a spiral spring. Two arrows of the speed indicator are connected to the axis of the cap, one of which is connected directly to the axis of the cap and rotates at the same speed as the cap, and the other is connected to the axis through a gear transmission and rotates at a speed 10 times lower. Thanks to this connection, one pointer needle makes a full revolution when the engine speed changes by 1,000 rpm, and the other when the shaft speed changes by 10,000 rpm. This improves the accuracy of the instrument's readings.

Fuel meters are designed to measure the amount of fuel in aircraft tanks. The principles of constructing fuel meters are based on measuring the level (volume) of fuel using a floating float, the weight of the fuel column using a pressure gauge and the parameters of electrical circuits when exposed to signals related to the level or pressure of the fuel. This group of instruments also includes oil meters, i.e., instruments used to measure the amount of oil on an airplane.

On modern aircraft, fuel tanks are located at a great distance from the instrument panel, and therefore fuel gauges must be remote. Electric fuel meters fully satisfy this requirement. The most widely used at present are capacitive fuel meters, the operating principle of which is based on measuring the capacitance value of special capacitors (sensors) associated with a certain relationship with the amount of fuel in the tank.

The sensitive element of the capacitive fuel meter is a cylindrical capacitor sensor, which is a set of two to six pipes coaxially located in relation to each other. The consistency of the gaps between the pipes is ensured by installing special insulating gaskets. Depending on the liquid level in the tank, the capacitance of the condenser will be different.

If a capacitor sensor is included in a bridge circuit, then as its capacitance changes when the liquid level changes, the bridge will become unbalanced. Voltage from the diagonal of the bridge will be supplied to the actuator (electric motor), which will move the fuel gauge needle to a new position.

Flow meters are used to measure instantaneous or average flow of liquids and gases per unit of time. Flow meters are used, for example, to control the consumption of fuel, oil, and air.

Based on the operating principle of the sensitive part, flow meters are divided into several types. However, most instruments are based on Bernoulli's law. In this regard, measuring the flow of liquids and gases actually comes down to measuring the speed of their movement at a constant cross-sectional area of the pipeline or, conversely, to measuring a variable area at a constant speed. Flow meters are also widely used, the operating principle of which is based on measuring the rotation speed of an impeller placed in the liquid flow.

Literature used: "Fundamentals of Aviation" authors: G.A. Nikitin, E.A. Bakanov

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Devices for monitoring aircraft parameters (engine monitoring devices) are designed to monitor the engine and all moving parts of the aircraft.

Dashboard of a modern airliner

Flight safety largely depends on the reliability of the engines. Therefore, several propulsion systems are most often used, so that if one of them fails, it is possible to continue to fly safely. This naturally leads to an increase in the number of sensors, so that in many cases the devices that monitor engine operation are combined on a special instrument panel and controlled by a flight engineer. Instruments for monitoring aircraft parameters include speed counters, lubricant, coolant and jet thermometers, fuel reserve and consumption indicators, etc.

Revolution counters can be designed as direct reading meters or as remote revolution counters. In their simplest mechanical form there are centrifugal type meters in which the indicator is directly driven by an elastic shaft. Devices for remote speed reading, in most cases, consist of an AC sensor on the engine and an indicator in the cockpit. Induction revolution counters are sometimes also used, but they interfere with magnetic compasses and must therefore be mounted at a large distance from them.

Fuel reserve and consumption indicators. It is very important for the pilot to have complete information about the appropriate fuel supply, which allows him to determine the possible maximum flight range. Older aircraft were most often equipped with a fuel level float indicator, which, depending on the case, was even mounted as a direct indicator above the fuel tank - for example, at the wing fuel tank - and read by the pilot from his seat. The readings of these instruments depend on their location and could hardly be used to indicate the fuel content of all fuel tanks on the cockpit instrument panel.

There was a need to use electrical systems in which the sensor installed on the fuel tank consists of a float and a potentiometer. Floats can be rotating or pendulum. Indicating devices are controlled by potentiometers. Also, thanks to additional contacts, they can take on the functions of an indicator of the presence of fuel in the tank. Modern aircraft use electrical reserve measurement on a capacitive basis. This method has the significant advantage that the measurement is no longer limited to a specific mark in the fuel tank. Several pipes located next to each other are built into it, and their capacity changes depending on the degree of use and is displayed on a dial indicator using a simple amplifier.

But measuring the reserve alone is no longer sufficient, especially for turbine engines that consume large amounts of fuel. Therefore, special flow meters are needed that measure the amount of fuel consumed by each engine in the fuel line (the so-called instantaneous fuel consumption indicator). These measuring instruments, thanks to a counting mechanism, allow you to read data regarding the remaining fuel in the tank at any time. Recently, some interesting autonomous meters have been developed that show either the remaining flight time or the remaining maximum range. The basis for performing autonomous calculations is the corresponding fuel consumption and engine operating mode.