Lubrication systems

The lubrication system, the cylinder, the piston, the piston rings, the cams / camshaft and the rod bearing.


Lubrication system
The engine lubrication system is designed to deliver clean oil at the correct temperature and pressure to every part of the engine. The oil is sucked out the sump into the pump, being the heart of the system, than forced through an oil filter and pressure feeded to the main bearings and to the oil pressure gauge. From the main bearings, the oil passes through feed-holes into drilled passages in the crankshaft and on to the big-end bearings of the connecting rod. The cylinder walls and piston-pin bearings are lubricated by oil fling dispersed by the rotating crankshaft. The excess being scraped off by the lower ring in the piston. A bleed or tributary from the main supply passage feeds each camshaft bearing. Another bleed supplies the timing chain or gears on the camshaft drive. The excess oil then drains back to the sump, where the heat is dispersed to the surrounding air.

Journal Bearings
If the crankshaft journals become worn the engine will have low oil pressure and throw oil all over the inside of the engine. The excessive splash will probably overwhelm the rings and cause the engine to use oil. Worn bearings surfaces can be restored by simply replacing the bearings inserts. In good maintained engines bearing wear occurs immediately after a cold start, because there’s little or no oil film between the bearing and shaft. At the moment that sufficient oil is circulated through the system hydrodynamic lubrication manifests and stop the progress of bearing wear.

Piston rings – cylinder
Piston rings provide a sliding seal preventing leakage of the fuel/air mixture and exhaust from the combustion chamber into the oil sump during compression and combustion. Secondly they keep oil in the sump from leaking into the combustion area, where it would be burned and lost. Most cars that “burn oil” and have to have a quart added every 1,000 miles are burning it because the rings no longer seal properly.
Between the piston rings and the cylinder wall of a well maintained engine hydrodynamic lubrication prevails, essential for the lowest friction and wear. In the top and bottom dead centre where the piston stops to redirect, the film thickness becomes minimal and mixed lubrication may exist.
To realize a good head transfer from the piston to the cylinder, an optimal sealing and a minimum of oil burning, a minimal film thickness is desirable. The film thickness is kept minimal by a so called oil control ring. This ring is situated beyond the piston rings so that the surplus of oil is directly scraped downwards to the sump. The oil film left on the cylinder wall by the passage of this ring is available to lubricate the following ring. This process is repeated for successive rings. On the up stroke the first compression ring is lubricated by the oil left behind on the cylinder wall during the down stroke.
Leakage of the fuel/air mixture and exhaust from the combustion chamber into the oil sump result in oil degradation. This is the reason why, despite of frequent replenish of oil, oil change remain essential or even become more essential.



In a turbine dry-sump lubrication system, the oil supply is carried in a tank mounted externally on or near the engine. With this type of system, a larger oil supply can be carried and the oil temperature can be controlled An oil cooler usually is included in a dry-sump oil system (Figure 5-l). This cooler may be air-cooled or fuel-cooled. The dry-sump oil system allows the axial-flow engines to retain their comparatively small diameter. This is done by designing the oil tank and the oil cooler to conform to the design of the engine.


The following component descriptions include most of those found in the various turbine lubrication systems. However, not all of these components will be found in any one system.

The dry-sump systems use an oil tank which contains most of the oil supply. However, a small sump usually is included on the engine to hold a supply of oil for an emergency system. The dry-sump system usually contains–

  • Oil pump.
  • Scavenge and pressure inlet strainers.
  • Scavenge return connection.
  • Pressure outlet ports.
  • Oil filter.
  • Mounting bosses for the oil pressure transmitter.
  • Temperature bulb connections.

A typical oil tank is shown in Figure 5-2. It is designed to furnish a constant supply of oil to the engine. This is done by a swivel outlet assembly mounted inside -the tank a horizontal baffle mounted in the center of the tank, two flapper check valves mounted on the baffle, and a positive-vent system.


The swivel outlet fitting is controlled by a weighted end, which is free to swing below the baffle. The flapper valves in the baffle are normally open. They close only when the oil in the bottom of the tank rushes to the top of the tank during deceleration. This traps the oil in the bottom of the tank where it is picked up by the swivel fitting A sump drain is located in the bottom of the tank. The airspace is vented at all times.

All oil tanks have expansion space. This allows for oil expansion after heat is absorbed from the bearings and gears and after the oil foams after circulating through the system. Some tanks also incorporate a deaerator tray. The tray separates air from the oil returned to the top of the tank by the scavenger system. Usually these deaerators are the “can” type in which oil enters a tangent. The air released is carried out through the vent system in the top of the tank. Inmost oil tanks a pressure buildup is desired within the tank. This assures a positive flow of oil to the oil pump inlet. This pressure buildup is made possible by running the vent line through an adjustable check-relief valve. The check-relief valve normally is set to relieve at about 4 psi pressure on the oil pump inlet.

There is little need for an oil-dilution system. If the air temperature is abnorrnally low, the oil may be changed to a lighter grade. Some engines may provide for the installation of an immersion-type oil heater.


In some engines the lubrication system is the wet-sump type. Because only a few models of centrifugal-flow engines are in operation, there are few engines using a wet-sump type of oil system.

The components of a wet-sump system are similar to many of a dry-sump system. The oil reservoir location is the major difference.

The reservoir for the wet-sump oil system may be the accessory gear case, which consists of the accessory gear casing and the front compressor bearing support casing. Or it may be a sump mounted on the bottom of the accessory case. Regardless of configuration reservoirs for wet-sump systems are an integral part of the engine and contain the bulk of the engine oil supply.

The following components are included in the wet-sump reservoir:

  • A bayonet-type gage indicates the oil level in the sump.
  • Two or more finger strainers (filters) are inserted in the accessory case for straining pressure and scavenged oil before it leaves or enters the sump. These strainers aid the main oil strainer.
  • A vent or breather equalizes pressure within the accessory casing.
  • A magnetic drain plug may be provided to drain the oil and to trap any ferrous metal particles in the oil. This plug should always be examined closely during inspections. The presence of metal particles may indicate gear or bearing failure.
  • A temperature bulb and an oil pressure fitting may be provided.

This system is typical of all engines using a wet-sump lubrication system. The bearing and drive gears in the accessory drive casing are lubricated by a splash system. The oil for the remaining points of lubrication leaves the pump under pressure. It passes through a filter to jet nozzles that direct the oil into the rotor bearings and couplings. Most wet-sump pressure systems are variable-pressure systems in which the pump outlet pressure depends on the engine RPM.

The scavenged oil is returned to the reservoir (sump) by gravity and pump suction. Oil from the front compressor bearing in the accessory-drive coupling shaft drains directly into the reservoir. Oil from the turbine coupling and the remaining rotor shaft bearings drains into a sump. The oil is then pumped by the scavenge element through a finger screen into the reservoir.


The oil system components used on gas turbine engines are–

  • Tanks.
  • Pressure pumps.
  • Scavenger pumps.
  • Filters.
  • Oil coolers.
  • Relief valves.
  • Breathers and pressurizing components.
  • Pressure and temperature gages lights.
  • Temperature-regulating valves.
  • Oil-jet nozzle.
  • Fittings, valves, and plumbing.
  • Chip detectors.

Not all of the units will be found in the oil system of any one engine. But a majority of the parts listed will be found in most engines.

Oil Tanks

Tanks can be either an airframe or engine-manufacturer-supplied unit. Usually constructed of welded sheet aluminum or steel, it provides a storage place for the oil. In most engines the tank is pressurized to ensure a constant supply of oil to the pressure pump. The tank can contain–

  • Venting system.
  • Deaerator to separate entrained air from the oil.
  • Oil level transmitter or dipstick.
  • Rigid or flexible oil pickup.
  • Coarse mesh screens.
  • Various oil and air inlets and outlets.

Pressure Pumps

Both gear- and Gerotor-type pumps are used in the lubricating system of the turbine engine. The gear-type pump consists of a driving and a driven gear. The engine-accessory section drives the rotation of the pump. Rotation causes the oil to pass around the outside of the gears in pockets formed by the gear teeth and the pump casing. The pressure developed is proportional to engine RPM up to the time the relief valve opens. After that any further increase in engine speed will not result in an oil pressure increase. The relief valve may be located in the pump housing or elsewhere in the pressure system for both types of pumps.

The Gerotor pump has two moving parts: an inner-toothed element meshing with an outer-toothed element. The inner element has one less tooth than the outer. The missing tooth provides a chamber to move the fluid from the intake to the discharge port. Both elements are mounted eccentrically to one another on the same shaft.

Scavenger Pumps

These pumps are similar to the pressure pumps but have a much larger total capacity. An engine is generally provided with several scavenger pumps to drain oil from various parts of the engine. Often one or two of the scavenger elements are incorporated in the same housing as the pressure pump (Figure 5-3). Different capacities can be provided for each system despite the common driving shaft speed. This is accomplished by varying the diameter or thickness of the gears to vary the volume of the tooth chamber. A vane-type pump may sometimes be used.


Oil Filters and Screens or Strainers

To prevent foreign matter from reaching internal parts of the engine, filters and screens or stainers are provided in the engine lubricating system. The three basic types of oil filters for the jet engine are the cartridge screen-disc and screen (Figures 5-4, 5-5 and 5-6). The cartridge filter is most commonly used and must be replaced periodically. The other two can be cleaned and reused. In the screen-disc filter there are a series of circular screen-type filters. Each filter is comprised of two layers of mesh forming a chamber between mesh layers. The filters are mounted on a common tube and arranged to provide a space between each circular element. Lube oil passes through the circular mesh elements and into the chamber between the two layers of mesh. This chamber is ported to the center of a common tube which directs oil out of the filter. Screens or strainers are placed at pressure oil inlets to bearings in the engine. This aids in preventing foreign matter from reaching the bearings.




To allow for oil flow in the event of filter blockage, all filters incorporate a bypass or relief valve as part of the filter or in the oil passages. When the pressure differential reaches a specified value (about 15 to 20 psi), the valve opens and allows oil to bypass the filter. Some filters incorporate a check valve. This prevents reverse flow or flow through the system when the engine is stopped Filtering characteristics vary, but most filters will stop particles of approximately 50 microns.

Magnetic Chip Detector

One or more magnetic chip detectors are installed on gas turbine engines. They are used to detect and attract ferrous material (metal with iron as its basic element) which may come from inside the engine. This ferrous material builds up until it bridges a gap. Whenever there is a requirement, the chip detectors may be collected and analyzed to determine the condition of the engine. Most engines utilize an electrical chip detector, located in the scavenger pump housing or in the accessory gearbox. Should the engine oil become contaminated with metal particles, the detector will catch some of them. This causes the warning light on the caution panel to come on.

Tubing, Hose, and Fittings

Tubing, hose, and fittings are used throughout the lubricating system. Their purpose is to connect apart into a system or to connect one part to another to complete a system.

Oil Pressure Indicating System

In a typical engine oil pressure indicating system the indicator receives inlet oil pressure indications from the oil pressure transmitter and provides readings in pounds per square inch Electrical power for oil pressure indicator and transmitter operation is supplied by the 28-volt AC system.

Oil-Pressure-Low Caution Light

Most gas turbine engine lubricating systems incorporate an engine oil-pressure-low caution light warning device into the system for safety purposes. The light is connected to a low-pressure switch. When pressure drops below a safe limit, the switch closes an electrical circuit causing the caution light to burn. Power is supplied by the 28-volt DC system.

Oil Temperature Indicating System

In a typical engine oil temperature indicating system, the indicator is connected to and receives temperature indications from an electrical resistance-type thermocouple or thermobulb. These are located in the pressure pump oil inlet side to the engine. Power to operate this circuit is supplied by the 28-volt DC system.

Oil Coolers

The oil cooler is used to reduce oil temperature by transmitting heat from the oil to another fluid usually fuel. Since the fuel flow through the cooler is much greater than the oil flow, the fuel is able to absorb a considerable amount of heat. This reduces the size and weight of the cooler. Thermostatic or pressure-sensitive valves control the oil temperature by determining whether the oil passes through or bypasses the cooler. Oil coolers are also cooled by air forced through them by a blower/fan.

Breathers and Pressurizing Systems

Internal oil leakage is kept to a minimum by pressurizing the bearing sump areas with air that is bled off the compressor (Figure 5-7). The airflow into the sump minimizes oil leakage across the seals in the reverse direction.


The oil scavenge pumps exceed the capacity of the lubrication pressure pump They are capable of handling considerably more oil than actually exists in the bearing sumps and gearboxes. Because the pumps area constant-displacement type, they make up for the lack of oil by pumping air from the sumps. Large quantities of air are delivered to the oil tank. Sump and tank pressures are maintained close to one another by a line which connects the two. If the sump pressure exceeds the tank pressure, the sump vent check valve opens, allowing the excess sump air to enter the oil tank. The valve allows flow only into the tank; oil or tank vapors cannot back up into the sump areas. Tank pressure is maintained little above ambient.

The scavenge pumps and sump-vent check valve functions result in relatively low air pressure in the sumps and gearboxes. These low internal sump pressures allow air to flow across the oil seals into the sumps. This airflow minimizes lube oil leakage across the seals. For this reason it is necessary to maintain sump pressures low enough to ensure seal-air leakage into the sumps. Under some conditions, the ability of the scavenge pumps to pump air forms a pressure low enough to cavitate the pumps or cause the sump to collapse. Under other conditions, too much air can enter the sump through worn seals.

If the seal leakage is not sufficient to maintain proper internal pressure, check valves in the sump and tank pressurizing valves open and allow ambient air to enter the system. Inadequate internal sump and gearbox pressure may be caused by seal leakage. If that occurs, air flows from the sumps, through the sump-vent check valve, the oil tank, the tank and sump pressurizing valves to the atmosphere. Tank pressure is always maintained a few pounds above ambient pressure by the sump and tank pressurizing valve.

The following addresses two types of lubrication systems used in the Army today: the General Electric T-701 turboshaft engine and the International/Solar T-62-series engine.






Flight instruments enable an airplane to be operated with maximum performance and enhanced safety, especially when flying long distances. Manufacturers provide the necessary flight instruments, but to use them effectively, pilots need to understand how they operate. This chapter covers the operational aspects of the pitot-static system and associated instruments, the vacuum system and associated instruments, and the magnetic compass.


There are two major parts of the pitot-static system: the impact pressure chamber and lines, and the static pressure chamber and lines. They provide the source of ambient air pressure for the operation of the altimeter, vertical speed indicator (vertical velocity indicator), and the airspeed indicator.



In this system, the impact air pressure (air striking the airplane because of its forward motion) is taken from a pitot tube, which is mounted in locations that provide minimum disturbance or turbulence caused by the motion of the airplane through the air. The static pressure (pressure of the still air) is usually taken from the static line attached to a vent or vents mounted flush with the side of the fuselage. This compensates for any possible variation in static pressure due to erratic changes in airplane attitude.

The openings of both the pitot tube and the static vent must be checked during the preflight inspection to assure that they are free from obstructions. Blocked or partially blocked openings should be cleaned by a certificated mechanic. Blowing into these openings is not recommended because this could damage the instruments.

As the airplane moves through the air, the impact pressure on the open pitot tube affects the pressure in the pitot chamber. Any change of pressure in the pitot chamber is transmitted through a line connected to the airspeed indicator, which utilizes impact pressure for its operation.


The static chamber is vented through small holes to the free undisturbed air, and as the atmospheric pressure increases or decreases, the pressure in the static chamber changes accordingly. Again, this pressure change is transmitted through lines to the instruments which utilize static pressure.

An alternate source for static pressure is provided in some airplanes in the event the static ports become blocked. This source usually is vented to the pressure inside the cockpit. Because of the venturi effect of the flow of air over the cockpit, this alternate static pressure is usually lower than the pressure provided by the normal static air source. When the alternate static source is used, the following differences in the instrument indications usually occur: the altimeter will indicate higher than the actual altitude, the airspeed will indicate greater than the actual airspeed, and the vertical speed will indicate a climb while in level flight. Consult the Airplane Flight Manual or Pilot’s Operating Handbook (AFM/POH) to determine the amount of error.

If the airplane is not equipped with an alternate static source, breaking the glass seal of the vertical speed indicator allows ambient air pressure to enter the static system. This makes the VSI unusable.



The altimeter measures the height of the airplane above a given pressure level. Since it is the only instrument that gives altitude information, the altimeter is one of the most vital instruments in the airplane. To use the altimeter effectively, its operation and how atmospheric pressure and temperature affect it must be thoroughly understood.

A stack of sealed aneroid (Aneroid—A sealed flexible container, which expands or contracts in relation to the surrounding air pressure. It is used in an altimeter or a barometer to measure the pressure of the air.) Wafers comprise the main component of the altimeter. These wafers expand and contract with changes in atmospheric pressure from the static source. The mechanical linkage translates these changes into pointer movements on the indicator.


The pressure altimeter is an aneroid barometer that measures the pressure of the atmosphere at the level where the altimeter is located, and presents an altitude indication in feet. The altimeter uses static pressure as its source of operation. Air is denser at sea level than aloft, so as altitude increases, atmospheric pressure decreases. This difference in pressure at various levels causes the altimeter to indicate changes in altitude.


The presentation of altitude varies considerably between different types of altimeters. Some have one pointer while others have two or more. Only the multi pointer type will be discussed in this handbook. The dial of a typical altimeter is graduated with numerals arranged clockwise from 0 to 9. Movement of the aneroid element is transmitted through gears to the three hands that indicate altitude. The shortest hand indicates altitude in tens of thousands of feet; the intermediate hand in thousands of feet; and the longest hand in hundreds of feet.

This indicated altitude is correct, however, only when the sea level barometric pressure is standard (29.92 inches of mercury), the sea level free air temperature is standard (+15°C or 59°F), and the pressure and temperature decrease at a standard rate with an increase in altitude. Adjustments for nonstandard conditions are accomplished by setting the corrected pressure into a barometric scale located on the face of the altimeter. Only after the altimeter is set does it indicate the correct altitude.


If no means were provided for adjusting altimeters to nonstandard pressure, flight could be hazardous. For example, if flying from a high-pressure area to a low-pressure area without adjusting the altimeter, the actual altitude of the airplane would be LOWER than the indicated altitude. An old saying, “HIGH TO LOW, LOOK OUT BELOW” is a way of remembering which condition is dangerous. When flying from a low-pressure area to a high-pressure area without adjusting the altimeter, the actual altitude of the airplane is HIGHER than the indicated altitude. 


Figure shows how variations in air temperature also affect the altimeter. On a warm day, a given mass of air expands to a larger volume than on a cold day, raising the pressure levels. For example, the pressure level where the altimeter indicates 5,000 feet is HIGHER on a warm day than under standard conditions. On a cold day, the reverse is true, and the pressure level where the altimeter indicates 5,000 feet is LOWER.  The adjustment to compensate for nonstandard pressure does not compensate for nonstandard temperature.

If terrain or obstacle clearance is a factor in selecting a cruising altitude, particularly at higher altitudes, remember to anticipate that a colder-than-standard temperature places the airplane LOWER than the altimeter indicates. Therefore, it is necessary to use a higher indicated altitude to provide adequate terrain clearance. Modify the memory aid to “HIGH TO LOW OR HOT TO COLD, LOOK OUT BELOW.”


There are two means by which the altimeter pointers can be moved. The first is a change in air pressure, while the other is an adjustment to the barometric scale. When the airplane climbs or descends, changing pressure within the altimeter case expands or contracts the aneroid barometer. This movement is transmitted through mechanical linkage to rotate the pointers. A decrease in pressure causes the altimeter to indicate an increase in altitude, and an increase in pressure causes the altimeter to indicate a decrease in altitude. Accordingly, if the airplane is flown from a pressure level of 28.75 a pressure level of 29.75 in. Hg., the altimeter would show a decrease of approximately 1,000 feet in altitude.

The other method of moving the pointers does not rely on changing air pressure, but the mechanical construction of the altimeter. Do not be confused by the fact that as the barometric pressure scale is moved, the indicator needles move in the same direction, which is opposite to the reaction the pointers have when air pressure changes. To illustrate this point, suppose the pilot lands at an airport with an elevation of 1,000 feet and the altimeter is correctly set to the current sea level pressure of 30.00 in. Hg. While the airplane is parked on the ramp, the pressure decreases to 29.50. The altimeter senses this as a climb and now indicates 1,500 feet.

When returning to the airplane, if the setting in the altimeter window is decreased to the current sea level pressure of 29.50, the indication will be reduced back down to 1,000 feet. Knowing the airplane’s altitude is vitally important to a pilot. The pilot must be sure that the airplane is flying high enough to clear the highest terrain or obstruction along the intended route. It is especially important to have accurate altitude information when visibility is restricted. To clear obstructions, the pilot must constantly be aware of the altitude of the airplane and the elevation of the surrounding terrain. To reduce the possibility of a midair collision, it is essential to maintain altitude in accordance with air traffic rules.


Altitude is vertical distance above some point or level used as a reference. There are as many kinds of altitude as there are reference levels from which altitude is measured, and each may be used for specific reasons. Pilots are mainly concerned with five types of altitudes:

Indicated Altitude—That altitude read directly from the altimeter (uncorrected) when it is set to the current altimeter setting.

True Altitude— The vertical distance of the airplane above sea level—the actual altitude. It is often expressed as feet above mean sea level (MSL). Airport, terrain, and obstacle elevations on aeronautical charts are true altitudes.

Absolute Altitude—the vertical distance of an airplane above the terrain, or above ground level (AGL).

Pressure Altitude— The altitude indicated when the altimeter setting window (barometric scale) is adjusted to 29.92. This is the altitude above the standard datum plane, which is a theoretical plane where air pressure. (Corrected to 15°C) equals 29.92 in. Hg. Pressure altitude is used to compute density altitude, true altitude, true airspeed, and other performance data.

Density Altitude—This altitude is pressure altitude corrected for variations from standard temperature. When conditions are standard, pressure altitude and density altitude are the same. If the temperature is above standard, the density altitude is higher than pressure altitude. If the temperature is below standard, the density altitude is lower than pressure altitude. This is an important altitude because it is directly related to the airplane’s performance.

As an example, consider an airport with a field elevation of 5,048 feet MSL where the standard temperature is 5°C. Under these conditions, pressure altitude and density altitude are the same—5,048 feet. If the temperature changes to 30°C, the density altitude increases to 7,855 feet. This means an airplane would perform on takeoff as though the field elevation were 7,855 feet at standard temperature. Conversely, a temperature of -25°C would result in a density altitude of 1,232 feet. An airplane would have much better performance under these conditions.

Instrument Check

To determine the condition of an altimeter, set the barometric scale to the altimeter setting transmitted by the local automated flight service station (AFSS) or any other reliable source. The altimeter pointers should indicate the surveyed elevation of the airport. If the indication is off more than 75 feet from the surveyed elevation, the instrument should be referred to a certificated instrument repair station for recalibration.



The airspeed indicator is a sensitive, differential pressure gauge which measures and shows promptly the difference between pitot or impact pressure, and static pressure, the undisturbed atmospheric pressure at level flight. These two pressures will be equal when the airplane is parked on the ground in calm air. When the airplane moves through the air, the pressure on the pitot line becomes greater than the pressure in the static lines. This difference in pressure is registered by the airspeed pointer on the face of the instrument, which is calibrated in miles per hour, knots, or both.


Pilots should understand the following speeds:

Indicated Airspeed (IAS)—The direct instrument reading obtained from the airspeed indicator, uncorrected for variations in atmospheric density, installation error, or instrument error. Manufacturers use this airspeed as the basis for determining airplane performance. Takeoff, landing, and stall speeds listed in the AFM or POH are indicated airspeeds and do not normally vary with altitude or temperature.

Calibrated Airspeed (CAS)—Indicated airspeed corrected for installation error and instrument error. Although manufacturers attempt to keep airspeed errors to a minimum, it is not possible to eliminate all errors throughout the airspeed operating range. At certain airspeeds and with certain flap settings, the installation and instrument errors may total several knots. This error is generally greatest at low airspeeds.

In the cruising and higher airspeed ranges, indicated airspeed and calibrated airspeed are approximately the same. Refer to the airspeed calibration chart to correct for possible airspeed errors.

True Airspeed (TAS)—Calibrated airspeed corrected for altitude and nonstandard temperature. Because air density decreases with an increase in altitude, an airplane has to be flown faster at higher altitudes to cause the same pressure difference between pitot impact pressure and static pressure. Therefore, for a given calibrated airspeed, true airspeed increases as altitude increases; or for a given true airspeed, calibrated airspeed decreases as altitude increases.

A pilot can find true airspeed by two methods. The most accurate method is to use a flight computer. With this method, the calibrated airspeed is corrected for temperature and pressure variation by using the airspeed correction scale on the computer. Extremely accurate electronic flight computers are also available. Just enter the CAS, pressure altitude, and temperature and the computer calculates the true airspeed. A second method, which is a “rule of thumb,” will provide the approximate true airspeed. Simply add 2 percent to the calibrated airspeed for each 1,000 feet of altitude.

Groundspeed (GS)— the actual speed of the airplane over the ground. It is true airspeed adjusted for wind. Groundspeed decreases with a headwind, and increases with a tailwind.


Airplanes weighing 12,500 pounds or less, manufactured after 1945, and certificated by the FAA, are required to have airspeed indicators marked in accordance with a standard color-coded marking.


• White arc—This arc is commonly referred to as the flap operating range since its lower limit represents the full flap stall speed and its upper limit provides the maximum flap speed. Approaches and landings are usually flown at speeds within the white arc.

• Lower limit of white arc (VS0)— The stalling speed or the minimum steady flight speed in the landing configuration. In small airplanes, this is the power-off stall speed at the maximum landing weight in the landing configuration (gear and flaps down).

• Upper limit of the white arc (VFE)—The maximum speed with the flaps extended.

• Green arc—This is the normal operating range of the airplane. Most flying occurs within this range.

• Lower limit of green arc (VS1)—The stalling speed or the minimum steady flight speed obtained in a specified configuration. For most airplanes, this is the power-off stall speed at the maximum takeoff weight in the clean configuration (gear up, if retractable, and flaps up).

• Upper limit of green arc (VNO)—The maximum structural cruising speed. Do not exceed this speed except in smooth air.

• Yellow arc—Caution range. Fly within this range only in smooth air, and then, only with caution.

• Red line (VNE)—Never-exceed speed. Operating above this speed is prohibited since it may result in damage or structural failure.


Some important airspeed limitations are not marked on the face of the airspeed indicator, but are found on placards and in the AFM or POH. These airspeeds include:

• Design maneuvering speed (VA)—This is the “rough air” speed and the maximum speed for abrupt maneuvers. If during flight, rough air or severe turbulence is encountered, reduce the airspeed to maneuvering speed or less to minimize stress on the airplane structure. It is important to consider weight when referencing this speed. For example, VA may be 100 knots when an airplane is heavily loaded, but only 90 knots when the load is light.

• Landing gear operating speed (VLO)—The maximum speed for extending or retracting the landing gear if using an airplane equipped with retractable landing gear.

• Landing gear extended speed (VLE)—The maximum speed at which an airplane can be safely flown with the landing gear extended.

• Best angle-of-climb speed (VX)—The airspeed at which an airplane gains the greatest amount of altitude in a given distance. It is used during a short-field takeoff to clear an obstacle.

• Best rate-of-climb speed (VY)— This airspeed provides the most altitude gain in a given period of time.

• Minimum control speed (VMC)— This is the minimum flight speed at which a light, twin-engine airplane can be satisfactorily controlled when an engine suddenly becomes inoperative and the remaining engine is at takeoff power.

• Best rate of climb with one engine inoperative (VYSE)— This airspeed provides the most altitude gain in a given period of time in a light, twin-engine airplane following an engine failure.

Instrument Check—Prior to takeoff, the airspeed indicator should read zero. However, if there is a strong wind blowing directly into the pitot tube, the airspeed indicator may read higher than zero. When beginning the takeoff, make sure the airspeed is increasing at an appropriate rate.

Vertical Speed Indicator (VSI)

The VSI, which is sometimes called a vertical velocity indicator (VVI), indicates whether the aircraft is climbing, descending, or in level flight. The rate of climb or descent is indicated in feet per minute (fpm). If properly calibrated, the VSI indicates zero in level flight. [Figure 7-5]


Figure 7-5. Vertical speed indicator (VSI).

Although the VSI operates solely from static pressure, it is a differential pressure instrument. It contains a diaphragm with connecting linkage and gearing to the indicator pointer inside an airtight case. The inside of the diaphragm is connected directly to the static line of the pitot-static system. The area outside the diaphragm, which is inside the instrument case, is also connected to the static line, but through a restricted orifice (calibrated leak).

Both the diaphragm and the case receive air from the static line at existing atmospheric pressure. The diaphragm receives unrestricted air while the case receives the static pressure via the metered leak. When the aircraft is on the ground or in level flight, the pressures inside the diaphragm and the instrument case are equal and the pointer is at the zero indication. When the aircraft climbs or descends, the pressure inside the diaphragm changes immediately, but due to the metering action of the restricted passage, the case pressure remains higher or lower for a short time, causing the diaphragm to contract or expand. This causes a pressure differential that is indicated on the instrument needle as a climb or descent.

When the pressure differential stabilizes at a definite ratio, the needle indicates the rate of altitude change.

The VSI displays two different types of information:

  • Trend information shows an immediate indication of an increase or decrease in the aircraft’s rate of climb or descent.
  • Rate information shows a stabilized rate of change in altitude.

The trend information is the direction of movement of the VSI needle. For example, if an aircraft is maintaining level flight and the pilot pulls back on the control yoke causing the nose of the aircraft to pitch up, the VSI needle moves upward to indicate a climb. If the pitch attitude is held constant, the needle stabilizes after a short period (6–9 seconds) and indicates the rate of climb in hundreds of fpm. The time period from the initial change in the rate of climb, until the VSI displays an accurate indication of the new rate, is called the lag. Rough control technique and turbulence can extend the lag period and cause erratic and unstable rate indications. Some aircraft are equipped with an instantaneous vertical speed indicator (IVSI), which incorporates accelerometers to compensate for the lag in the typical VSI. [Figure 7-6]


Figure 7-6. An IVSI incorporates accelerometers to help the instrument immediately indicate changes in vertical speed.



Several flight instruments utilize the properties of a gyroscope for their operation. The most common instruments containing gyroscopes are the turn coordinator, heading indicator, and the attitude indicator. To understand how these instruments operate requires knowledge of the instrument power systems, gyroscopic principles, and the operating principles of each instrument.


Any spinning object exhibits gyroscopic properties. A wheel or rotor designed and mounted to utilize these properties is called a gyroscope. Two important design characteristics of an instrument gyro are great weight for its size, or high density, and rotation at high speed with low friction bearings.

There are two general types of mountings; the type used depends upon which property of the gyro is utilized. A freely or universally mounted gyroscope is free to rotate in any direction about its center of gravity. Such a wheel is said to have three planes of freedom. The wheel or rotor is free to rotate in any plane in relation to the base and is so balanced that with the gyro wheel at rest, it will remain in the position in which it is placed. Restricted or semi rigidly mounted gyroscopes are those mounted so that one of the planes of freedom is held fixed in relation to the base.

There are two fundamental properties of gyroscopic action—rigidity in space and precession.


Rigidity in space refers to the principle that a gyroscope remains in a fixed position in the plane in which it is spinning. By mounting this wheel, or gyroscope, on a set of gimbal rings(Gimbal Ring—A type of support that allows an object, such as a gyroscope, to remain in an upright condition when its base is tilted), the gyro is able to rotate freely in any direction. Thus, if the gimbal rings are tilted, twisted, or otherwise moved, the gyro remains in the plane in which it was originally spinning.



Precession is the tilting or turning of a gyro in response to a deflective force. The reaction to this force does not occur at the point where it was applied; rather, it occurs at a point that is 90° later in the direction of rotation. This principle allows the gyro to determine a rate of turn by sensing the amount of pressure created by a change in direction. The rate at which the gyro processes is inversely proportional to the speed of the rotor and proportional to the deflective force. Precession can also create some minor errors in some instruments.



In some airplanes, all the gyros are vacuum, pressure, or electrically operated; in others, vacuum or pressure systems provide the power for the heading and attitude indicators, while the electrical system provides the power for the turn coordinator. Most airplanes have at least two sources of power to ensure at least one source of bank information if one power source fails. The vacuum or pressure system spins the gyro by drawing a stream of air against the rotor vanes to spin the rotor at high speed, much like the operation of a waterwheel or turbine. The amount of vacuum or pressure required for instrument operation varies, but is usually between 4.5 and 5.5 in. Hg.

One source of vacuum for the gyros is a vane-type engine-driven pump that is mounted on the accessory case of the engine. Pump capacity varies in different airplanes, depending on the number of gyros. A typical vacuum system consists of an engine-driven vacuum pump, relief valve, air filter, gauge, and tubing necessary to complete the connections. The gauge is mounted in the airplane’s instrument panel and indicates the amount of pressure in the system (vacuum is measured in inches of mercury less than ambient pressure).


As shown in figure, air is drawn into the vacuum system by the engine-driven vacuum pump. It first goes through a filter, which prevents foreign matter from entering the vacuum or pressure system. The air then moves through the attitude and heading indicators, where it causes the gyros to spin. A relief valve prevents the vacuum pressure, or suction, from exceeding prescribed limits. After that, the air is expelled overboard or used in other systems, such as for inflating pneumatic deicing boots. It is important to monitor vacuum pressure during flight, because the attitude and heading indicators may not provide reliable information when suction pressure is low. The vacuum, or suction, gauge generally is marked to indicate the normal range. Some airplanes are equipped with a warning light that illuminates when the vacuum pressure drops below the acceptable level.


Engine instruments that indicate oil pressure, oil temperature, engine speed, exhaust gas temperature, and fuel flow are common to both turbine and reciprocating engines. However, there are some instruments that are unique to turbine engines. These instruments provide indications of engine pressure ratio, turbine discharge pressure, and torque. In addition, most gas turbine engines have multiple temperature-sensing instruments, called thermo couples, that provide pilots with temperature readings in and around the turbine section.
An engine pressure ratio (EPR) gauge is used to indicate the power output of a turbojet/turbofan engine. EPR is the ratio of turbine discharge to compressor inlet pressure. Pressure measurements are recorded by probes installed in the engine inlet and at the exhaust. Once collected, the data is sent to a differential pressure transducer, which is indicated on a cockpit EPR gauge.
EPR system design automatically compensates for the effects of airspeed and altitude. However, changes in ambient temperature do require a correction to be applied to EPR indications to provide accurate engine power settings.


A limiting factor in a gas turbine engine is the temperature of the turbine section. The temperature of a turbine section must be monitored closely to prevent overheating the turbine blades and other exhaust section components. One common way of monitoring the temperature of a turbine section is with an exhaust gas temperature (EGT) gauge. EGT is an engine operating limit used to monitor overall engine operating conditions.
Variations of EGT systems bear different names based on the location of the temperature sensors. Common turbine temperature sensing gauges include the turbine inlet temperature (TIT) gauge, turbine outlet temperature (TOT) gauge, inter stage turbine temperature (ITT) gauge, and turbine gas temperature (TGT) gauge.
Turboprop/turbo shaft engine power output is measured by the torque meter. Torque is a twisting force applied to a shaft. The torque meter measures power applied to the shaft. Turboprop and turbo shaft engines are designed to produce torque for driving a propeller.
Torque meters are calibrated in percentage units, foot-pounds, or pounds per square inch.

Exhaust Gas Temperature (EGT)

A limiting factor in a gas turbine engine is the temperature of the turbine section. The temperature of a turbine section must be monitored closely to prevent overheating the turbine blades and other exhaust section components. One common way of monitoring the temperature of a turbine section is with an EGT gauge. EGT is an engine operating limit used to monitor overall engine operating conditions.
Variations of EGT systems bear different names based on the location of the temperature sensors. Common turbine temperature sensing gauges include the turbine inlet temperature (TIT) gauge, turbine outlet temperature (TOT) gauge, inter-stage turbine temperature (ITT) gauge, and turbine gas temperature (TGT) gauge.


N1 represents the rotational speed of the low pressure compressor and is presented on the indicator as a percentage of design r.p.m. After start the speed of the low pressure compressor is governed by the N1 turbine wheel. The N1 turbine wheel is connected to the low pressure compressor through a concentric shaft.
N2 represents the rotational speed of the high pressure compressor and is presented on the indicator as a percentage of design r.p.m. The high pressure compressor is governed by the N2 turbine wheel. The N2 turbine wheel is connected to the high pressure compressor through a concentric shaft.

Fuel Gauges

The fuel quantity gauges indicate the amount of fuel measured by a sensing unit in each fuel tank and is displayed in gallons or pounds. Aircraft certification rules require accuracy in fuel gauges only when they read “empty.” Any reading other than “empty” should be verified. Do not depend solely on the accuracy of the fuel quantity gauges. Always visually check the fuel level in each tank during the preflight inspection, and then compare it with the corresponding fuel quantity indication.
If a fuel pump is installed in the fuel system, a fuel pressure gauge is also included. This gauge indicates the pressure in the fuel lines. The normal operating pressure can be found in the AFM/POH or on the gauge by color coding.


A propeller with fixed blade angles is a fixed-pitch propeller. The pitch of this propeller is set by the manufacturer and cannot be changed. Since a fixed-pitch propeller achieves the best efficiency only at a given combination of airspeed and rpm, the pitch setting is ideal for neither cruise nor climb. Thus, the aircraft suffers a bit in each performance category. The fixed-pitch propeller is used when low weight, simplicity, and low cost are needed.
There are two types of fixed-pitch propellers: climb and cruise. Whether the airplane has a climb or cruise propeller installed depends upon its intended use. The climb propeller has a lower pitch, therefore less drag. Less drag results in higher rpm and more horsepower capability, this increases performance during takeoffs and climbs, but decreases performance during cruising flight.

The cruise propeller has a higher pitch, therefore more drag. More drag results in lower rpm and less horsepower capability, which decreases performance during takeoffs and climbs, but increases efficiency during cruising flight.

The propeller is usually mounted on a shaft, which may be an extension of the engine crankshaft. In this case, the rpm of the propeller would be the same as the crankshaft rpm. On some engines, the propeller is mounted on a shaft geared to the engine crankshaft. In this type, the rpm of the propeller is different than that of the engine.


In a fixed-pitch propeller, the tachometer is the indicator of engine power. [Figure 6-8] A tachometer is calibrated in hundreds of rpm and gives a direct indication of the engine and propeller rpm. The instrument is color coded, with a green arc denoting the maximum continuous operating rpm. Some tachometers have additional markings to reflect engine and/or propeller limitations. The manufacturer’s recommendations should be used as a reference to clarify any misunderstanding of tachometer markings.

The rpm is regulated by the throttle, which controls the fuel/air flow to the engine. At a given altitude, the higher the tachometer reading, the higher the power output of the engine.
When operating altitude increases, the tachometer may not show correct power output of the engine. For example, 2,300 rpm at 5,000 feet produces less horsepower than 2,300 rpm at sea level because power output depends on air density. Air density decreases as altitude increases and a decrease in air density (higher density altitude) decreases the power output of the engine. As altitude changes, the position of the throttle must be changed to maintain the same rpm. As altitude is increased, the throttle must be opened further to indicate the same rpm as at a lower altitude.

Outside Air Temperature (OAT) Gauge

The outside air temperature (OAT) gauge is a simple and effective device mounted so that the sensing element is exposed to the outside air. The sensing element consists of a bimetallic-type thermometer in which two dissimilar materials are welded together in a single strip and twisted into a helix. One end is anchored into protective tube and the other end is affixed to the pointer, which reads against the calibration on a circular face. OAT gauges are calibrated in degrees °C, °F, or both. An accurate air temperature provides the pilot with useful information about temperature lapse rate with altitude change.  [Figure 7-38]


Figure 7-38. Outside air temperature (OAT) gauge.




At atmospheric condition we have sufficient pressure (3 p.s.i.) to breath freely. As the altitude increases pressure drops. Till 8,000 ft altitude the pressure variation won’t affect us but beyond that we will get affect (hypoxia), so in order to have same pressure to breath freely in the cabin of aircraft, we need to go for maintaining the pressure artificially.

It may be done by

  • I. Cabin Superchargers
  • II. Engine –Driven Compressors
  • III. Compressed air from engine

As like pressure temperature also drops as the altitude increases, but up to certain layer of atmosphere which is stratosphere. Most of the aircraft are operating within Stratosphere. So we need to have a constant temperature in our cabin, so that passengers feel the comfort of their travel.

It can be done by

  • i. Air cycle cooling system
  • ii. Vapor cycle cooling system

Most of the aircraft have mingled the pressurization and air conditioning system in order to simplify the system.




1. Left engine

2. Right engine

3. Flow limiter

4. Primary heat exchanger

5. Primary heat exchanger bypass valve

6. Shut off valve

7. Compressor

8. Secondary heat exchanger

9. Water separator

10. Secondary heat exchanger bypass valve

11. Refrigeration unit

1. Air from the compressor sections of the two engines is taken for air conditioning and pressurization. We are making a cross connection so that it can supply uniform flow of air.

2. Air is going to the flow limiter. Flow limiter limits the amount of compressed air to entry into the system. Suppose if there is any pipeline ruptured in air cycle system, then flow limiter won’t allow the compressed air to enter into the air cycle system.

3. Partial amount of air from the flow limiter goes to the primary heat exchanger. Primary heat exchanger utilizes ram air for cooling purpose. The compressed air taken from the engine compressor may be at temp range of 200 deg to 400 deg Celsius. We are cooling the air without reducing much pressure by using heat exchanger, and heat exchanger works on Convective type.

4. Another amount of compressed air from flow limiter goes to mingle with the heat exchanger outlet air to make a constant temperature air of 300 °F. This constant temperature can be attained only by proper operation of Primary heat exchanger (PHE) bypass valve. For example if heat exchanger outlet air is at 200°F, but I need an output of 300°F at outlet portion of Primary heat exchanger bypass valve, so we are opening the bypass valve for some designated time and mixing the hot(directed from flow limiter) and cold (PHE outlet ) flows.

5 .This 300°F temperature air is going to split for three Purpose

a) For Anti-icing

b) To supply hot air to cabin(if required)

c) To refrigeration unit

6. This 300°F air can be directly used for anti-icing and de-icing purpose. This air will be taken by the tubes and will be sprayed on the leading edge through suitable arrangements. And one part of the 300°F air flow is directed to the cabin for hot air supply.

7. Remaining part of the air is directed towards refrigeration unit for further cooling. On the way there is main control that is “Main shut off Valve”. This can be directly controlled by the pilot.

8. After main shut off valve there is Refrigeration unit, this contains

a) Compressor

b) Turbine

c) Water Separator

Both turbine and compressor are connected by the same shaft.

9. The air flows to the compressor region; there it strikes the compressor blade and makes rotating at initial. After compressor starts it compresses the air and Pressurizes the air so some amount of heat may be added to the air. Then Air is going to the Secondary heat exchanger (SHE), where the heat from the air is taken by the ram air by convection. The air has been cooled now.

10. The output of SHE goes to the Turbine. Since because of the turbine, here the cold gas is allowed to expand, so that pressure drops, temperature again drops, it may be on some minus °F sometimes.

11. After expansion in the turbine, air which is in a circular motion is allowed to go for the water separator region for separation of water particles in the cold air. Since water particles are denser than air they get attached to the walls of the water separator due to centrifugal force. In some specified place there are some holes made on the water separator to drain the water particles attached.

12. In the cabin we need only a temperature range of 60°F-125°F (15°C-51°C) and a pressure of 3p.s.i. which is suitable for human.

13. By opening and closing the refrigerant bypass valve, we can mingle the pure cold air and 300°F air, to make a possible living temperature for human beings.

14. After maintaining to proper temperature and pressure the air is allowed to go to the cabin by suitable pipelines.

15. If there is any problem on the total system means, we can directly mix the hot air supply with ram air (which is taken near SHE) and maintain the proper temperature by proper mixing. But this method is only for emergency purpose. Please note when using this ram air method of cooling pressurization should be done separately by cabin superchargers or whatever the device builds up pressure.


Vapor cycle cooling system in Aircraft

Physical principle

Liquids can be vaporized at any temperature by changing the pressure acting on it. To clearly understand this concept, we will take an example of water contained in a vessel. When the vessel is at atmospheric pressure, the water will be boiling at 100°C when heating. If I am pressurizing the vessel to more than the atmospheric pressure, then water will not boil at 100°C. If I am creating a drop in pressure in the vessel by a vacuum pump, then water will boil at a temperature less than 100°C.

Basic law of thermodynamics states that heat will always flow from hot body to cold body. If I need a reverse of this, I have to add some work.

Typical System operation:


1. We are having the FREON as refrigerant in the vapor cycle cooling system. It has a boiling point of 4°C.

2. At the receiver I am having high pressure, so that FREON will have high boiling point.

3. When the vapor cycle system is switched on, the compressor starts delivering the pressure and thus making flow.

4. The highly pressurized FREON at the receiver is in liquid phase. When the Freon flows through the circuit, first it expands at the Expansion valve. So pressure has been dropped (i.e. Boiling point decreased).

5. The less pressure Freon then goes to the evaporator stage. Evaporator will be exposed to Cabin. We blow the warm air of cabin over the evaporator coils by fan, and thus doing a forced convection.

6. The heat transferred to the Freon makes it to change the phase which is from liquid to vapor.

7. The less pressure Freon vapor is then compressed by the Compressor and thus it delivers high temperature high pressure Freon vapor.

8. Now this high pressure and high temperature Freon vapor enters the Condenser coils where the cool air from atmosphere will be blown over the coils (here too making a forced convection). Condenser will be exposed to the Atmosphere. Because of heat transfer the Freon losses heat and returns to liquid phase.

9. Then it goes to the receiver (high pressure low temperature Freon liquid)

10. The cycle continues as stated. Before doing any type of maintenance activities to the vapor cycle system, we have to purge the system with inert gas in a open atmosphere.

11. Freon is colorless, odorless, and non toxic; however, being heavier than air, it will displace oxygen and cause suffocation. When heated over an open flame, it converts to phosgene which is deadly!

12. To know the Freon level in the circuit a sight glass arrangement will be employed between Receiver to Expansion valve. If the unit requires additional refrigerant, bubbles will be present in the sight glass otherwise steady.


When using this method air conditioning, pressurization have to be done separately, either by using Cabin superchargers or Engine driven compressor or by using your future design that delivers pressure with least energy taken as input.



Because fire is one of the most dangerous threats to an aircraft, the potential fire zones of modern multi-engine aircraft are protected by a fixed fire protection system. A ‘fire zone’ is an area of region of an aircraft designed by the manufacturer to require fire detection and or fire extinguishing equipment and a high degree of inherent fire resistance. The term ‘fixed’ describes a permanently installed system in contrast to any type of portable fire extinguishing equipment, such as hand-held Co2 fire extinguisher.

Fire suppression system includes

I. Fire detection system

II. Smoke detection system

III. Fire extinguishing system

Fire detection system should signal the presence of fire. Units of the system are installed in locations where there are greatest possibilities of a fire. The common fire detection systems are

v Thermal Switch

v Thermocouple

v Continuous loop Systems



Systron donner

Thermal Switch:

Thermal switches are heat –sensitive units that complete electrical circuits at a certain temperature. They are connected in parallel with each other. If the temperature rises above a set value in any one section of the circuit, the thermal switch will close the electrical circuit and gives indication to the pilot through indication lights.

Thermo-couple systems:

A thermocouple gives the rate of temperature rise.

A Thermocouple is the junction of two dissimilar metals which generates a small electric current that varies according to the temperature of the junction. For this reason it does not require an external power source. The dissimilar metals can be constantan and iron, Alumel and Chromel, or some other combination of metals or alloys which will produce the required results. The complete thermocouple circuit consists of the ‘cold’ junction, the ‘hot’ junction, electric leads (made from the same material as the thermocouple), and a galvanometer-type indicating instrument as illustrated below.


Constantan is chosen as one of the metals because its resistance is affected very little by changes in temperature. When the hot junction is at higher or lower temperature than the cold junction, a current will flow through the circuit and instrument. The value of current is depending on the difference in temperature between the two junctions.

Continuous loop systems:

A continuous loop detector or sensing system permits more complete coverage of a fire hazard area than any type of detectors. There is no rate of heat rise sensitivity in a continuous loop system.

Kidde System


In the kidde continuous loop system, two wires are imbedded in a special ceramic core within a inconel tube. One of the two wires in the kidde sensing system is welded to the case at each end and acts as an internal ground. The second wire is a hot lead (above ground potential) that provides a current signal when ceramic core material changes its resistance with change in temperature. If any fire is near this instrument causes the ceramic core resistance to drop and thereby wire will conduct more current than usual, from that we can find the fire.

Fenwal System:

It uses a single wire surrounded by a continuous string of ceramic beads in an inconel tube. The beads in the Fenwal detector are wetted with a eutectic salt which posses the characteristics of suddenly lowering its electrical resistance as the sensing element reaches its alarm temperature.




If the volume of a gas is held constant, its pressure increases as temperature increases; thus the helium between the two tubing walls exerts a pressure proportional to the average absolute temperature along the entire length of the tube. One end of the tube is connected to a small chamber containing a metal diaphragm switch. One side of the diaphragm is therefore exposed to the sensor tube pressure and one to the ambient pressure. When system pressure exceeds some predetermined value, diaphragm will activate the electrical contacts which in-turn activates the alarm. When the temperature is reduced to the normal levels after action has been taken to suppress the fire, the gas absorption material reabsorbs the gas and pressure drops, thus opening the switch and cutting off the alarm.


Smoke detection System

Pure air just contains 78% of nitrogen, 21% of oxygen and 1% rest of other gases. When there is fire prevailing in the aircraft the air will get dusted because of reaction between Carbon and oxygen thus forms CO, CO2 and other gases. In order to eliminate the fire, smoke has to be identified first. It can be done by

a) Light detection

b) Light Refraction

c) Ionization

d) Solid state semiconductor

Light detection type:

We all know photo voltaic cell converts light energy into electrical energy. The pure light rays will be fall on photo voltaic panel and produces some electrical energy at normal. When the area is surrounded by smoke the intensity of light rays falling on the panel gets differ. From that we gets alarm.


Light Refraction type smoke detector:

Here a controlled mass of air is taken for smoke detection by use of fan or whatever blowing device. Source of light is the Electric bulb. It is made to pass it rays from a single point in order to satisfy our requirement. The light which fall on the wall of the detection chamber gets reflected and the reflected rays which falls on photo voltaic panel induces some current based on the intensity of rays fallen. When smoke particles comes into picture, they itself reflect the rays before colliding by the wall and makes different electrical energy by making different intensity, thus alarm gets activated.


Ionization type smoke detector:

The Radioactive material (Americium) emits Alpha rays continuously. The Alpha rays makes reaction with the clean cabin air (under normal conditions) and ionize the incoming air ingredients to Positive and Negative charges. Those Positive charged ions are attracted by negative charged pole, and Negative charged ions are attracted by positive charged poles. Thus it gives some reading. When smoked air comes into picture the amount of attraction on poles varies and thus we are getting smoke detection alarm activated.


Solid State Smoke Detector:


A solid filament coated with semi-conductor material is along the length of the fuselage to detect the fire. An electrical current is given over the filament by the wrap around coils. At normal conditions the filament will have some resistance and it produces some output value. When smoke comes into picture, the semi conductor has ability to absorb carbon particles, so that resistance of the filament varies, thus getting different output from that of we got on normal.

If suppose the atmospheric air itself getting too much carbon particles, then we may commit a false smoke warning, so that in order to activate alarm we are just comparing the filament exposed to atmospheric air and cabin.



After finding out the fired place or region we have to use suitable extinguishing agent in order to maintain safety. Types of fire and the suitable extinguishing agent filled extinguisher are figure below.


For Aircraft fire extinguishing operation, we are having extinguishing agent stored at an arrangement called Fire Bottle. It will be routed to the region in need of extinguishing by suitable hoses and Sprayed through Nozzles. The Common shapes of fire bottle are Cylindrical and Spherical.

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If the aircraft gets fire in the ground, we may use ordinary Fire extinguishers. But make sure that the fire extinguisher type which you use should be specified as per the figure shown at first. For example if you use water filled extinguisher for a fire caused by electrical sources, then it is useless. After selection of specified extinguisher spray the extinguishing agent in the base of the fire by opening the Safety pin and by pressing the trigger.

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Anti-Ice and Deice Systems

Anti-icing equipment is designed to prevent the formation of ice, while deicing equipment is designed to remove ice once it has formed. These systems protect the leading edge of wing and tail surfaces, pitot and static port openings, fuel tank vents, stall warning devices, windshields, and propeller blades. Ice detection lighting may also be installed on some aircraft to determine the extent of structural icing during night flights.

Most light aircraft have only a heated pitot tube and are not certified for flight in icing. These light aircraft have limited cross-country capability in the cooler climates during late fall, winter, and early spring. Noncertificated aircraft must exit icing conditions immediately. Refer to the AFM/POH for details.

Airfoil Anti-Ice and Deice

Inflatable deicing boots consist of a rubber sheet bonded to the leading edge of the airfoil. When ice builds up on the leading edge, an engine-driven pneumatic pump inflates the rubber boots. Many turboprop aircraft divert engine bleed air to the wing to inflate the rubber boots. Upon inflation, the ice is cracked and should fall off the leading edge of the wing. Deicing boots are controlled from the flight deck by a switch and can be operated in a single cycle or allowed to cycle at automatic, timed intervals. [Figure 6-48]


Figure 6-48. Deicing boots on the leading edge of the wing.

In the past it was believed that if the boots were cycled too soon after encountering ice, the ice layer would expand instead of breaking off, resulting in a condition referred to as ice “bridging.” Consequently, subsequent deice boot cycles would be ineffective at removing the ice buildup. Although some residual ice may remain after a boot cycle, “bridging” does not occur with any modern boots. Pilots can cycle the boots as soon as an ice accumulation is observed. Consult the AFM/POH for information on the operation of deice boots on an aircraft.

Many deicing boot systems use the instrument system suction gauge and a pneumatic pressure gauge to indicate proper boot operation. These gauges have range markings that indicate the operating limits for boot operation. Some systems may also incorporate an annunciator light to indicate proper boot operation.

Proper maintenance and care of deicing boots are important for continued operation of this system. They need to be carefully inspected during preflight.

Another type of leading edge protection is the thermal anti-ice system. Heat provides one of the most effective methods for preventing ice accumulation on an airfoil. High performance turbine aircraft often direct hot air from the compressor section of the engine to the leading edge surfaces. The hot air heats the leading edge surfaces sufficiently to prevent the formation of ice. A newer type of thermal anti-ice system referred to as thermawing uses electrically heated graphite foil laminate applied to the leading edge of the wing and horizontal stabilizer. Thermawing systems typically have two zones of heat application. One zone on the leading edge receives continuous heat; the second zone further aft receives heat in cycles to dislodge the ice allowing aerodynamic forces to remove it. Thermal anti-ice systems should be activated prior to entering icing conditions.

An alternate type of leading edge protection that is not as common as thermal anti-ice and deicing boots is known as a weeping wing. The weeping-wing design uses small holes located in the leading edge of the wing to prevent the formation and build-up of ice. An antifreeze solution is pumped to the leading edge and weeps out through the holes. Additionally, the weeping wing is capable of deicing an aircraft. When ice has accumulated on the leading edges, application of the antifreeze solution chemically breaks down the bond between the ice and airframe, allowing aerodynamic forces to remove the ice. [Figure 6-48]


Figure 6-48. TKS weeping wing anti-ice/deicing system.

Windscreen Anti-Ice

There are two main types of windscreen anti-ice systems. The first system directs a flow of alcohol to the windscreen. If used early enough, the alcohol will prevent ice from building up on the windscreen. The rate of alcohol flow can be controlled by a dial in the flight deck according to procedures recommended by the aircraft manufacturer.

Another effective method of anti-icing equipment is the electric heating method. Small wires or other conductive material is imbedded in the windscreen. The heater can be turned on by a switch in the flight deck, causing an electrical current to be passed across the shield through the wires to provide sufficient heat to prevent the formation of ice on the windscreen. The heated windscreen should only be used during flight. Do not leave it on during ground operations, as it can overheat and cause damage to the windscreen. Warning: the electrical current can cause compass deviation errors by as much as 40°.

Propeller Anti-Ice

Propellers are protected from icing by the use of alcohol or electrically heated elements. Some propellers are equipped with a discharge nozzle that is pointed toward the root of the blade. Alcohol is discharged from the nozzles, and centrifugal force drives the alcohol down the leading edge of the blade. The boots are also grooved to help direct the flow of alcohol. This prevents ice from forming on the leading edge of the propeller. Propellers can also be fitted with propeller anti-ice boots. The propeller boot is divided into two sections—the inboard and the outboard sections. The boots are imbedded with electrical wires that carry current for heating the propeller. The prop anti-ice system can be monitored for proper operation by monitoring the prop anti-ice ammeter. During the preflight inspection, check the propeller boots for proper operation. If a boot fails to heat one blade, an unequal blade loading can result, and may cause severe propeller vibration. [Figure 6-49]


Figure 6-49. Prop ammeter and anti-ice boots.

Other Anti-Ice and Deice Systems

Pitot and static ports, fuel vents, stall-warning sensors, and other optional equipment may be heated by electrical elements. Operational checks of the electrically heated systems are to be checked in accordance with the AFM /POH.

Operation of aircraft anti-icing and deicing systems should be checked prior to encountering icing conditions. Encounters with structural ice require immediate action. Anti-icing and deicing equipment are not intended to sustain long-term flight in icing conditions.


Fuel systems

The fuel system is designed to provide an uninterrupted flow of clean fuel from the fuel tanks to the engine. The fuel must be available to the engine under all conditions of engine power, altitude, attitude, and during all approved flight maneuvers. Two common classifications apply to fuel systems in small airplanes – gravity-feed and fuel-pump systems.


The gravity-feed system utilizes the force of gravity to transfer the fuel from the tanks to the engine – for example, on high-wing airplanes where the fuel tanks are installed in the wings. This places the fuel tanks above the carburetor, and the fuel is gravity fed through the system and into the carburetor. If the design of the airplane is such that gravity cannot be used to transfer fuel, fuel pumps are installed – for example, on low-wing airplanes where the fuel tanks in the wings are located below the carburetor.


Figure 15: Gravity-feed and fuel-pump systems.

Fuel pumps

Airplanes with fuel pump systems have two fuel pumps. The main pump system is engine driven, and an electrically driven auxiliary pump is provided for use in engine starting and in the event the engine pump fails. The auxiliary pump, also known as a boost pump, provides added reliability to the fuel system. The electrically driven auxiliary pump is controlled by a switch in the cockpit.

Fuel primer

Both gravity fed and pump systems may incorporate a fuel primer into the system. The primer is used to draw fuel from the tanks to vaporize it directly into the cylinders prior to starting the engine. This is particularly helpful during cold weather, when engines are hard to start because there is not enough heat available to vaporize the fuel in the carburetor. It is important to lock the primer in place when it is not in use. If the knob is free to move, it may vibrate out during flight and can cause an excessively rich mixture. To avoid overpriming, read the priming instructions for your airplane.

Fuel tanks

The fuel tanks, normally located inside the wings of an airplane, have a filler opening on top of the wing through which they can be filled. A filler cap covers this opening. The tanks are vented to the outside to maintain atmospheric pressure inside the tank. They may be vented through the filler cap or through a tube extending through the surface of the wing. Fuel tanks also include an overflow drain that may stand alone or be collocated with the fuel tank vent. This allows fuel to expand with increases in temperature without damage to the tank itself. If the tanks have been filled on a hot day, it is not unusual to see fuel coming from the overflow drain.

Fuel gauges

The fuel quantity gauges indicate the amount of fuel measured by a sensing unit in each fuel tank and is displayed in gallons or pounds. Aircraft certification rules only require accuracy in fuel gauges when they read “empty.” Any reading other than “empty” should be verified. Do not depend solely on the accuracy of the fuel quantity gauges. Always visually check the fuel level in each tank during the preflight inspection, and then compare it with the corresponding fuel quantity indication.

If a fuel pump is installed in the fuel system, a fuel pressure gauge is also included. This gauge indicates the pressure in the fuel lines. The normal operating pressure can be found in the AFM/POH, or on the gauge by color coding.

Fuel selectors

The fuel selector valve allows selection of fuel from various tanks. A common type of selector valve contains four positions: LEFT, RIGHT, BOTH, and OFF. Selecting the LEFT or RIGHT position allows fuel to feed only from that tank, while selecting the BOTH position feeds fuel from both tanks. The LEFT or RIGHT position may be used to balance the amount of fuel remaining in each wing tank.


Figure 16: Fuel selector valve.

Fuel placards will show any limitations on fuel tank usage, such as “level flight only” and/or “both” for landings and takeoffs.

Regardless of the type of fuel selector in use, fuel consumption should be monitored closely to ensure that a tank does not run completely out of fuel. Running a fuel tank dry will not only cause the engine to stop, but running for prolonged periods on one tank causes an unbalanced fuel load between tanks. Running a tank completely dry may allow air to enter the fuel system, which may cause vapor lock. When this situation develops, it may be difficult to restart the engine. On fuel-injected engines, the fuel may become so hot it vaporizes in the fuel line, not allowing fuel to reach the cylinders.

Fuel strainers, sumps, and drains

After the fuel selector valve, the fuel passes through a strainer before it enters the carburetor. This strainer removes moisture and other sediments that might be in the system. Since these contaminants are heavier than aviation fuel, they settle in a sump at the bottom of the strainer assembly. A sump is defined as a low point in a fuel system and/or fuel tank. The fuel system may contain sump, fuel strainer, and fuel tank drains, some of which may be collocated.

The fuel strainer should be drained before each flight.

Fuel samples should be drained and checked visually for water and contaminants. Water in the sump is hazardous because in cold weather the water can freeze and block fuel lines. In warm weather, it can flow into the carburetor and stop the engine. If water is present in the sump, it is likely there is more water in the fuel tanks, and you should continue to drain them until there is no evidence of water. In any event, never take off until you are certain that all water and contaminants have been removed from the engine fuel system.

Because of the variation in fuel systems, you should become thoroughly familiar with the systems that apply to your airplane. Consult the AFM or POH for specific operating procedures.

Fuel grades

Aviation gasoline, or AVGAS, is identified by an octane or performance number (grade), which designates the antiknock value or knock resistance of the fuel mixture in the engine cylinder. The higher the grade of gasoline, the more pressure the fuel can withstand without detonating. Lower grades of fuel are used in lower-compression engines because these fuels ignite at a lower temperature. Higher grades are used in higher-compression engines, because they must ignite at higher temperatures, but not prematurely. If the proper grade of fuel is not available, use the next higher grade as a substitute. Never use a lower grade. This can cause the cylinder head temperature and engine oil temperature to exceed their normal operating range, which may result in detonation.

Several grades of aviation fuel are available. Care must be exercised to ensure that the correct aviation grade is being used for the specific type of engine. The proper fuel grade is stated in the AFM or POH, on placards in the cockpit, and next to the filler caps. Due to its lead content, auto gas should NEVER be used in aircraft engines unless the aircraft has been modified with a Supplemental Type Certificate (STC) issued by the Federal Aviation Administration.

The current method to identify aviation gasoline for aircraft with reciprocating engines is by the octane and performance number, along with the abbreviation AVGAS. These aircraft use AVGAS 80, 100, and 100LL. Although AVGAS 100LL performs the same as grade 100, the “LL” indicates it has a low lead content.

Fuel for aircraft with turbine engines is classified as JET A, JET A-1, and JET B. Jet fuel is basically kerosene and has a distinctive kerosene smell.

Since use of the correct fuel is critical, dyes are added to help identify the type and grade of fuel.


Figure 17: Aviation fuel color-coding system.

In addition to the color of the fuel itself, the color-coding system extends to decals and various airport fuel handling equipment. For example, all aviation gasolines are identified by name, using white letters on a red background. In contrast, turbine fuels are identified by white letters on a black background.

Fuel contamination

Of the accidents attributed to powerplant failure from fuel contamination, most have been traced to:

  • Inadequate preflight inspection by the pilot.
  • Servicing aircraft with improperly filtered fuel from small tanks or drums.
  • Storing aircraft with partially filled fuel tanks.
  • Lack of proper maintenance.

Fuel should be drained from the fuel strainer quick drain and from each fuel tank sump into a transparent container, and then checked for dirt and water. When the fuel strainer is being drained, water in the tank may not appear until all the fuel has been drained from the lines leading to the tank. This indicates that water remains in the tank, and is not forcing the fuel out of the fuel lines leading to the fuel strainer. Therefore, drain enough fuel from the fuel strainer to be certain that fuel is being drained from the tank. The amount will depend on the length of fuel line from the tank to the drain. If water or other contaminants are found in the first sample, drain further samples until no trace appears.

Water may also remain in the fuel tanks after the drainage from the fuel strainer had ceased to show any trace of water. This residual water can be removed only by draining the fuel tank sump drains.

Water is the principal fuel contaminant. Suspended water droplets in the fuel can be identified by a cloudy appearance of the fuel or by the clear separation of water from the colored fuel, which occurs after the water has settled to the bottom of the tank. As a safety measure, the fuel sumps should be drained before every flight during the preflight inspection.

Fuel tanks should be filled after each flight, or at least after the last flight of the day to prevent moisture condensation within the tank. Another way to prevent fuel contamination is to avoid refueling from cans and drums. Refueling from cans or drums may result in fuel contamination.

The use of a funnel and chamois skin when refueling from cans or drums is hazardous under any conditions, and should be discouraged. In remote areas or in emergency situations, there may be no alternative to refueling from sources with inadequate anticontamination systems, and a chamois and funnel may be the only possible means of filtering fuel. However, the use of a chamois will not always ensure decontaminated fuel.

Worn-out chamois will not filter water; neither will a new, clean chamois that is already water-wet or damp.

Most imitation chamois skins will not filter water.

Refueling procedures

Static electricity is formed by the friction of air passing over the surfaces of an airplane in flight and by the flow of fuel through the hose and nozzle during refueling.

Nylon, dacron, or wool clothing is especially prone to accumulate and discharge static electricity from the person to the funnel or nozzle. To guard against the possibility of static electricity igniting fuel fumes, a ground wire should be attached to the aircraft before the fuel cap is removed from the tank. The refueling nozzle then should be grounded to the aircraft before refueling is begun, and should remain grounded throughout the refueling process. When a fuel truck is used, it should be grounded prior to the fuel nozzle contacting the aircraft.

If fueling from drums or cans is necessary, proper bonding and grounding connections are important.

Drums should be placed near grounding posts, and the following sequence of connections observed:

  1. Drum to ground.
  2. Ground to aircraft.
  3. Drum to aircraft.
  4. Nozzle to aircraft before the fuel cap is removed.

When disconnecting, reverse the order.

The passage of fuel through a chamois increases the charge of static electricity and the danger of sparks. The aircraft must be properly grounded and the nozzle, chamois filter, and funnel bonded to the aircraft. If a can is used, it should be connected to either the grounding post or the funnel. Under no circumstances should a plastic bucket or similar nonconductive container be used in this operation.

Jet engine Fuel system:



Starting system

Most small aircraft use a direct-cranking electric starter system. This system consists of a source of electricity, wiring, switches, and solenoids to operate the starter and a starter motor. Most aircraft have starters that automatically engage and disengage when operated, but some older aircraft have starters that are mechanically engaged by a lever actuated by the pilot.

The starter engages the aircraft flywheel, rotating the engine at a speed that allows the engine to start and maintain operation.

Electrical power for starting is usually supplied by an on-board battery, but can also be supplied by external power through an external power receptacle. When the battery switch is turned on, electricity is supplied to the main power bus through the battery solenoid. Both the starter and the starter switch draw current from the main bus, but the starter will not operate until the starting solenoid is energized by the starter switch being turned to the “start” position. When the starter switch is released from the “start” position, the solenoid removes power from the starter motor. The starter motor is protected from being driven by the engine through a clutch in the starter drive that allows the engine to run faster than the starter motor.


Figure 18: Typical starting circuit.

When starting an engine, the rules of safety and courtesy should be strictly observed. One of the most important is to make sure there is no one near the propeller. In addition, the wheels should be chocked and the brakes set, to avoid hazards caused by unintentional movement. To avoid damage to the propeller and property, the airplane should be in an area where the propeller will not stir up gravel or dust.

Ignition System: Spark Ignition Engine


In a spark ignition engine the ignition system provides a spark that ignites the fuel/air mixture in the cylinders and is made up of magnetos, spark plugs, high-tension leads, and the ignition switch. [Figure 6-16]

A magneto uses a permanent magnet to generate an electrical current completely independent of the aircraft’s electrical system. The magneto generates sufficiently high voltage to jump a spark across the spark plug gap in each cylinder. The system begins to fire when the starter is engaged and the crankshaft begins to turn. It continues to operate whenever the crankshaft is rotating.

Most standard certificated aircraft incorporate a dual ignition system with two individual magnetos, separate sets of wires, and spark plugs to increase reliability of the ignition system. Each magneto operates independently to fire one of the two spark plugs in each cylinder. The firing of two spark plugs improves combustion of the fuel/air mixture and results in a slightly higher power output. If one of the magnetos fails, the other is unaffected. The engine will continue to operate normally, although a slight decrease in engine power can be expected. The same is true if one of the two spark plugs in a cylinder fails.
The operation of the magneto is controlled in the flight deck by the ignition switch. The switch has five positions:

  1. OFF
  2. R (right)
  3. L (left)
  4. BOTH
  5. START

With RIGHT or LEFT selected, only the associated magneto is activated. The system operates on both magnetos with BOTH selected.
A malfunctioning ignition system can be identified during the pretakeoff check by observing the decrease in rpm that occurs when the ignition switch is first moved from BOTH to RIGHT, and then from BOTH to LEFT. A small decrease in engine rpm is normal during this check. The permissible decrease is listed in the AFM or POH. If the engine stops running when switched to one magneto or if the rpm drop exceeds the allowable limit, do not fly the aircraft until the problem is corrected. The cause could be fouled plugs, broken or shorted wires between the magneto and the plugs, or improperly timed firing of the plugs. It should be noted that “no drop” in rpm is not normal, and in that instance, the aircraft should not be flown.
Following engine shutdown, turn the ignition switch to the OFF position. Even with the battery and master switches OFF, the engine can fire and turn over if the ignition switch is left ON and the propeller is moved because the magneto requires no outside source of electrical power. Be aware of the potential for serious injury in this situation.
Even with the ignition switch in the OFF position, if the ground wire between the magneto and the ignition switch becomes disconnected or broken, the engine could accidentally start if the propeller is moved with residual fuel in the cylinder. If this occurs, the only way to stop the engine is to move the mixture lever to the idle cutoff position, then have the system checked by a qualified aviation maintenance technician.

Fuel injection systems

In a fuel injection system, the fuel is injected either directly into the cylinders, or just ahead of the intake valve. A fuel injection system is considered to be less susceptible to icing than the carburetor system. Impact icing on the air intake, however, is a possibility in either system. Impact icing occurs when ice forms on the exterior of the airplane, and blocks openings such as the air intake for the injection system.

The air intake for the fuel injection system is similar to that used in the carburetor system, with an alternate air source located within the engine cowling. This source is used if the external air source is obstructed. The alternate air source is usually operated automatically, with a backup manual system that can be used if the automatic feature malfunctions.

A fuel injection system usually incorporates these basic components—an engine-driven fuel pump, a fuel/air control unit, fuel manifold (fuel distributor), discharge nozzles, an auxiliary fuel pump, and fuel pressure/flow indicators.


Figure 10: Fuel injection system.

The auxiliary fuel pump provides fuel under pressure to the fuel/air control unit for engine starting and/or emergency use. After starting, the engine-driven fuel pump provides fuel under pressure from the fuel tank to the fuel/air control unit. This control unit, which essentially replaces the carburetor, meters fuel based on the mixture control setting, and sends it to the fuel manifold valve at a rate controlled by the throttle. After reaching the fuel manifold valve, the fuel is distributed to the individual fuel discharge nozzles. The discharge nozzles, which are located in each cylinder head, inject the fuel/air mixture directly into each cylinder intake port.

Some of the advantages of fuel injection are:

  • Reduction in evaporative icing.
  • Better fuel flow.
  • Faster throttle response.
  • Precise control of mixture.
  • Better fuel distribution.
  • Easier cold weather starts.

Disadvantages usually include:

  • Difficulty in starting a hot engine.
  • Vapor locks during ground operations on hot days.
  • Problems associated with restarting an engine that quits because of fuel starvation.

LUBRICATION SYSTEM.—The oil lubrication systems of modern gas turbine engines vary in design and plumbing. However, most systems have units that perform similar functions. In a majority of cases, a pressure pump or system furnishes oil to lubricate and cool several parts of the engine. A scavenging system returns the oil to the tank for reuse. Overheating is a problem in gas turbine engines. Overheating is more severe after the engine stops than while it is running. Oil flow, which normally cools the bearings, stops. The heat stored in the turbine wheel now raises the temperature of the bearings much higher than when the engine was running. The oil moves heat away from these bearings to prevent overheating. Most systems include a heat exchanger to cool the oil. Many systems have pressurized sumps and a pressurized oil tank. This equipment ensures a constant head pressure to the pressure lubrication pump to prevent pump cavitation at high altitudes. Oil consumption is relatively low in a gas turbine engine compared to a piston-type engine. Oil consumption in the turbine engine primarily depends

upon the efficiency of the seals. However, oil can be lost through internal leakage, and, in some engines, by malfunctioning of the pressurizing or venting system. Oil sealing is very important in a jet engine. Any wetting of the blades or vanes by oil vapor causes accumulation of dust or dirt. Since oil consumption is so low, oil tanks are made small to decrease weight and storage problems. The main parts of the turbine requiring lubrication and cooling are the main bearings and accessory drive gears. Therefore, lubrication of the gas turbine engine is simple. In some engines the oil operates the servomechanism of fuel controls and controls the position of the variable-area exhaust nozzle vanes. Because each engine bearing gets its oil from a metered or calibrated opening, the lubrication system is known as the calibrated type. With few exceptions, the lubricating system is of the dry sump design. This design carries the bulk of the oil in an airframe or engine-supplied separate tank. In the wet sump system, the oil is carried in the engine itself. All gas turbine engine lubrication systems normally use synthetic oil. Figure 6-18 shows components that usually make up the dry sump oil system of a gas turbine engine.




Hydraulic system

The word hydraulics is based on the Greek word for water, and originally meant the study of water at rest and in motion. Today the meaning has been expanded to include the physical behavior of all liquids, including hydraulic fluid. With the use of incompressible phenomenon of liquid we can easily make a hydraulic system.


As per Pascal’s law “Pressure applied to any part of a confined liquid is transmitted with undiminished intensity to every other parts” .The basic idea behind any hydraulic system is very simple: Force that is applied at one point is transmitted to another point using an incompressible fluid. The fluid is almost always an oil of some sort. The force is almost always multiplied in the process.

In this drawing, two pistons (red) fit into two glass cylinders filled with oil (light blue) and connected to one another with an oil-filled pipe. If you apply a downward force to one piston (the left one in this drawing), then the force is transmitted to the second piston through the oil in the pipe. Since oil is in-compressible, the efficiency is very good — almost all of the applied force appears at the second piston. The great thing about hydraulic systems is that the pipe connecting the two cylinders can be any length and shape, allowing it to snake through all sorts of things separating the two pistons. The pipe can also fork, so that one master cylinder can drive more than one slave cylinder if desired.


There are multiple applications for hydraulic use in airplanes, depending on the complexity of the airplane. For example, hydraulics is often used on small airplanes to operate wheel brakes, retractable landing gear, and some constant speed propellers. On large airplanes, hydraulics is used for flight control surfaces, wing flaps, spoilers, and other systems.

A basic hydraulic system consists of a reservoir, pump (either hand, electric, or engine driven), a filter to keep the fluid clean, selector valve to control the direction of flow, relief valve to relieve excess pressure, and an actuator. The hydraulic fluid is pumped through the system to an actuator or servo.

Servos can be either single-acting or double-acting servos based on the needs of the system. This means that the fluid can be applied to one or both sides of the servo, depending on the servo type, and therefore provides power in one direction with a single-acting servo. A servo is a cylinder with a piston inside that turns fluid power into work and creates the power needed to move an aircraft system or flight control. The selector valve allows the fluid direction to be controlled. This is necessary for operations like the extension and retraction of landing gear where the fluid must work in two different directions. The relief valve provides an outlet for the system in the event of excessive fluid pressure in the system. Each system incorporates different components to meet the individual needs of different aircraft.

A mineral-based fluid is the most widely used type for small airplanes. This type of hydraulic fluid, which is a kerosene-like petroleum product, has good lubricating Properties, as well as additives to inhibit foaming and prevent the formation of corrosion. It is quite stable chemically, has very little viscosity change with temperature, and is dyed for identification. Since several types of hydraulic fluids are commonly used, make sure your airplane is serviced with the type specified by the manufacturer.

The three types of gas-charged accumulators you’ll encounter on hydraulic systems are bladder, piston and diaphragm. Accumulators are used to store the fluid under given pressure.

The most popular of these is the bladder type. Bladder accumulators feature fast response (less than 25 milliseconds), a maximum gas compression ratio of around 4:1 and a maximum flow rate of 15 liters (4 gallons) per second, although “high-flow” versions up to 38 liters (10 gallons) per second are available. Bladder accumulators also have good dirt tolerance; they are mostly unaffected by particle contamination in the hydraulic fluid.


Piston accumulators, on the other hand, can handle much higher gas compression ratios (up to 10:1) and flow rates as high as 215 liters (57 gallons) per second. Unlike bladder accumulators, whose preferred mounting position is vertical to prevent the possibility of fluid getting trapped between the bladder and the shell, piston accumulators can be mounted in any position.

But, piston accumulators also require a higher level of fluid cleanliness than bladder units, have slower response times (greater than 25 milliseconds) – especially at lower pressures – and exhibit hysteresis. This is explained by the static friction of the piston seal which has to be overcome, and the necessary acceleration and deceleration of the piston mass.

Diaphragm accumulators have most of the advantages of bladder-type units but can handle gas compression ratios up to 8:1. They are limited to smaller volumes, and their performance can sometimes be affected by gas permeation across the diaphragm.

Maintenance Considerations

When charging the gas end of a bladder or diaphragm accumulator, the nitrogen gas should always be admitted very slowly. If the high-pressure nitrogen is allowed to expand rapidly as it enters the bladder, it can chill the bladder’s polymeric material to the point where immediate brittle failure occurs. Rapid pre-charging can also force the bladder underneath the poppet at the oil-end, causing it to be cut. If pre-charge pressure is too high or minimum system pressure is reduced without a corresponding reduction in pre-charge pressure, the operation of the accumulator will be affected and damage may also result. Excessive pre-charge of a bladder accumulator can drive the bladder into the poppet assembly during discharge, causing damage to the poppet assembly and/or the bladder. This is a common cause of bladder failure.


Low or no pre-charge also can have drastic consequences for bladder accumulators. It can result in the bladder being crushed into the top of the shell by system pressure. This can cause the bladder to extrude into or be punctured by the gas valve. In this scenario, only one such cycle is required to destroy the bladder.

Similarly, excessively high or low pre-charge of a piston accumulator can cause the piston to bottom out at the end of its stroke, resulting in damage to the piston and its seal. The good news is that, if this happens, an audible warning will result. Even though piston accumulators can be damaged by improper charging, they are much more tolerant of it than bladder accumulators.

Artificial feel devices

With purely mechanical flight control systems, the aerodynamic forces on the control surfaces are transmitted through the mechanisms and are felt directly by the pilot, allowing tactile feedback of airspeed. With hydro mechanical flight control systems, however, the load on the surfaces cannot be felt and there is a risk of overstressing the aircraft through excessive control surface movement. To overcome this problem, artificial feel systems can be used.

For example, for the controls of the RAF’s Avro Vulcan jet bomber and the RCAF’s Avro Canada CF-105 Arrow supersonic interceptor (both 1950s-era designs), the required force feedback was achieved by a spring device.The fulcrum of this device was moved in proportion to the square of the air speed (for the elevators) to give increased resistance at higher speeds. For the controls of the American Vought F-8 Crusader and the LTV A-7 Corsair II warplanes, a ‘bob-weight’ was used in the pitch axis of the control stick, giving force feedback that was proportional to the airplane’s normal acceleration.

Stick shaker

A stick shaker is a device (available in some hydraulic aircraft) that is fitted into the control column, which shakes the control column when the aircraft is about to stall. Also in some aircraft like the McDonnell Douglas DC-10 there is/was a back-up electrical power supply that the pilot can turn on to re-activates the stick shaker in case the hydraulic connection to the stick shaker is lost.

Types of Hydraulic Fluids:

       When adding fluid to the system, use the specified type of fluid in the manufactures manual. There are 3 types of fluids are currently being used in civil aircraft

ü Vegetable base hydraulic fluid

ü Mineral base hydraulic fluid

ü Phosphate ester base hydraulic fluid


1. Ease of installation

2. Simple inspection needed & requires minimum maintenance

Air in the System:

It is important that a hydraulic system contains no air bubbles. You may have heard about the need to “bleed the air out of the brake lines” of your car. If there is an air bubble in the system, then the force applied to the first piston gets used compressing the air in the bubble rather than moving the second piston, which has a big effect on the efficiency of the system.

Control Surface deflection using hydraulic system

· The piston rod can only produce Reciprocating motion.

· Reciprocating motion can be converted to Radial or oblique motion by the use of mechanical linkages.

Pneumatic system

Pneumatic is a branch of technology, which deals with the study and application of pressurized gas to effect mechanical motion.

Pneumatic systems are extensively used in industry, where factories are commonly plumbed with compressed air or compressed inert gases. This is because a centrally located and electrically powered compressor that powers cylinders and other pneumatic devices through solenoid valves is often able to provide motive power in a cheaper, safer, more flexible, and more reliable way than a large number of electric motors and actuators.

Pneumatic also has applications in dentistry, construction, mining, and other areas.



Pump that compresses air, raising air pressure to above ambient pressure for use in pneumatic systems.

Check valve:

One-way valve – allows pressurized air to enter the pneumatic system, but prevents backflow of air toward the Compressor when Compressor is stopped (prevent loss of pressure.


· Stores compressed air,

· Prevents surges in pressure

· Prevents constant Compressor operation (“duty cycles” of Compressor)


Directional Valve: (Selector valve)

ü Controls pressurized air flow from Accumulator (source to user equipment via selected port

ü Some valves are one way – shut tight

ü Some valves are two way, allowing free exhaust from the port not selected

ü valves can be actuated manually or electrically.



ü Converts energy stored in compressed air into mechanical motion

ü Example is a linear piston (piston limited to moving in two opposing directions)

ü Other examples are alternate tools including: rotary actuators, air tools, expanding bladders, etc



Pneumatic uses in Aircraft

  • Powers engine Suction System for Heading indicators and Attitude indicators.
  • Actuates Landing Gear (some aircraft)
  • Emergency Brakes (some aircraft)
  • Cabin Pressure (for pressurized aircraft)



Airplane brakes are located on the main wheels and are applied by either a hand control or by foot pedals (toe or heel). Foot pedals operate independently and allow for differential braking. During ground operations, differential braking can supplement nosewheel/tailwheel steering.


Aircraft Landing Gear

The landing gear forms the principal support of an aircraft on the surface. The most common type of landing gear consists of wheels, but aircraft can also be equipped with floats for water operations or skis for landing on snow. [Figure 6-37]


The landing gear on small aircraft consists of three wheels: two main wheels (one located on each side of the fuselage) and a third wheel positioned either at the front or rear of the airplane. Landing gear employing a rear-mounted wheel is called conventional landing gear. Airplanes with conventional landing gear are often referred to as tailwheel airplanes. When the third wheel is located on the nose, it is called a nosewheel, and the design is referred to as a tricycle gear. A steerable nosewheel or tailwheel permits the airplane to be controlled throughout all operations while on the ground.

Tricycle Landing Gear Airplanes

A tricycle gear airplane has three advantages:

  1. It allows more forceful application of the brakes during landings at high speeds without causing the aircraft to nose over.
  2. It permits better forward visibility for the pilot during takeoff, landing, and taxiing.
  3. It tends to prevent ground looping (swerving) by providing more directional stability during ground operation since the aircraft’s center of gravity (CG) is forward of the main wheels. The forward CG keeps the airplane moving forward in a straight line rather than ground looping.

Nosewheels are either steerable or castering. Steerable nosewheels are linked to the rudders by cables or rods, while castering nosewheels are free to swivel. In both cases, the aircraft is steered using the rudder pedals. Aircraft with a castering nosewheel may require the pilot to combine the use of the rudder pedals with independent use of the brakes.

Tailwheel Landing Gear Airplanes


Tailwheel landing gear aircraft have two main wheels attached to the airframe ahead of its CG that support most of the weight of the structure. A tailwheel at the very back of the fuselage provides a third point of support. This arrangement allows adequate ground clearance for a larger propeller and is more desirable for operations on unimproved fields. [Figure 6-38]

With the CG located behind the main gear, directional control of this type aircraft becomes more difficult while on the ground. This is the main disadvantage of the tailwheel landing gear. For example, if the pilot allows the aircraft to swerve while rolling on the ground at a low speed, he or she may not have sufficient rudder control and the CG will attempt to get ahead of the main gear which may cause the airplane to ground loop.
Lack of good forward visibility when the tailwheel is on or near the ground is a second disadvantage of tailwheel landing gear aircraft. These inherent problems mean specific training is required in tailwheel aircraft.

Fixed and Retractable Landing Gear

Landing gear can also be classified as either fixed or retractable. A fixed gear always remains extended and has the advantage of simplicity combined with low maintenance. A retractable gear is designed to streamline the airplane by allowing the landing gear to be stowed inside the structure during cruising flight. [Figure 6-39]


Flight Control System

Flight Control System


The architecture of the flight control system, essential for all flight operations, has significantly changed throughout the years. Soon after the first flights, articulated surfaces were introduced for basic control, operated by the pilot through a system of cables and pulleys. This technique survived for decades and is now still used for small airplanes.

The introduction of larger airplanes and the increase of flight envelopes made the muscular effort of the pilot, in many conditions, not sufficient to contrast the aerodynamic hinge moments consequent to the surface deflection; the first solution to this problem was the introduction of aerodynamic balances and tabs, but further grow of the aircraft sizes and flight enveops brought to the need of powered systems to control the articulated aerodynamic surfaces.

Nowadays two great categories of flight control systems can be found: a full mechanical control on gliders and small general aviation, and a powered, or servo-assisted, control on large or combat aircraft.

One of the great additional effects after the introduction of servomechanisms is the possibility of using active control technology, working directly on the flight control actuators, for a series of benefits:

• compensation for deficiencies in the aerodynamics of the basic airframe;

• stabilisation and control of unstable airplanes, that have commonly higher performances;

• flight at high angles of attack;

• automatic stall and spinning protection;

• gust alleviation.


Fig. 6.1 – Flight control surfaces on airliner

A further evolution of the servo-assisted control is the fly-by-wire technique, based on signal processing of the pilot’s demand before conversion into actuator control.

The number and type of aerodynamic surfaces to be controlled changes with aircraft category. Fig. 6.1 shows the classic layout for a conventional airliner. Aircraft have a number of different control surfaces:

those indicated in red form the primary flight control, i.e. pitch, roll and yaw control, basically obtained by deflection of elevators, ailerons and rudder (and combinations of them); those indicated in blue form the secondary flight control: high-lift and lift-dump devices, airbrakes, tail trimming, etc.

Modern aircraft have often particular configurations, typically as follows:

• elevons on delta wings, for pitch and roll control, if there is no horizontal tail;

• flaperons, or trailing edge flaps-ailerons extended along the entire span:

• tailerons, or stabillisers-ailerons (independently controlled);

• swing wings, with an articulation that allows sweep angle variation;

• canards, with additional pitch control and stabilization

Primary flight control capability is essential for safety, and this aspect is dramatically emphasized in the modern unstable (military) airplanes, which could be not controlled without the continued operation of the primary flight control surfaces. For this reason the actuation system in charge of primary control has a high redundancy and reliability, and is capable of operating close to full performance after one or more failures.

Secondary actuation system failure can only introduce flight restriction, like a flapless landing or reduction in the max angle of attack; therefore it is not necessary to ensure full operation after failures.

Conventional Systems/Direct mechanical control

As mentioned in the introduction, the linkage from cabin to control surface can be fully mechanical if the aircraft size and its flight envelop allow; in this case the hinge moment generated by the surface deflection is low enough to be easily contrasted by the muscular effort of the pilot.


Fig 6.2 – Push-pull rod system for elevator control

Two types of mechanical systems are used: push-pull rods and cable-pulley.

In the first case a sequence of rods link the control surface to the cabin input. Bell-crank levers are used to change the direction of the rod routings: fig. 6.2 sketches the push-pull control rod system between the elevator and the cabin control column; the bell-crank lever is here necessary to alter the direction of the transmission and to obtain the conventional coupling between stick movement and elevator deflection (column fwd = down deflection of surface and pitch down control).

From this simplified description the main requirements of a push-pull rod system are clear. First of all the linkage must be stiff, to avoid any unwanted deflection during flight and due to fuselage elasticity. Second, axial instability during compression must be excluded; the instability load P for a rod is given by:



E = Young modulus;

I = cross-section moment of inertia;

λ = reference length.

The reference length is linked to the real length of the rod, meaning that to increase the instability load the length must be decreased, or the rods must be frequently constrained by slide guides, or the routing must be interrupted with bell-cranks.

Finally a modal analysis of the system layout is sometimes necessary, because vibrations


Fig 6.3 – Cables and pulleys system for elevator control

of the rods can introduce oscillating deflections of the surface; this problem is particularly important on helicopters, because vibrations generated by the main rotor can induce a dramatic resonance of the flight control rods.

The same operation described before can be done by a cable-pulley system, where couples of cables are used in place of the rods. In this case pulleys are used to alter the direction of the lines, equipped with idlers to reduce any slack due to structure elasticity, cable strands relaxation or thermal expansion. Often the cable-pulley solution is preferred, because is more flexible and allows reaching more remote areas of the airplane. An example is sketched in fig. 6.3, where the cabin column is linked via a rod to a quadrant, which the cables are connected to.

Fully powered Flight Controls

To actuate the control Surface the pilot has to give full effort. This is very tough to actuate the control surfaces through simple mechanical linkages. One can feel the equal toughness when raising the hand perpendicular to the airflow on riding a motorbike.

In this type of flight control system we will have






The cable

To transmit the power


Cable connector

To connect the cable



To adjust the Cable length



To guide the Cable



To guide the in radial direction


Push pull rod

To go for and aft as per requirement


Control stick

To make orders for the remaining circuit




The most basic flight control system designs are mechanical and date back to early aircraft. They operate with a collection of mechanical parts such as rods, cables, pulleys, and sometimes chains to transmit the forces of the flight deck controls to the control surfaces. Mechanical flight control systems are still used today in small general and sport category aircraft where the aerodynamic forces are not excessive. When the pilot pushes the control stick forward/backward the cable is getting tensed through the linkages and it causes the Control surface to move respectively.

Power actuated systems

Hydraulic control

When the pilot’s action is not directly sufficient for a the control, the main option is a powered system that assists the pilot.

A few control surfaces on board are operated by electrical motors: as already discussed in a previous chapter, the hydraulic system has demonstrated to be a more suitable solution for actuation in terms of reliability, safety, weight per unit power and flexibility, with respect to the electrical system, then becoming the common tendency on most modern airplanes: the pilot, via the cabin components, sends a signal, or demand, to a valve that opens ports through which high pressure hydraulic fluid flows and operates one or more actuators.

The valve, that is located near the actuators, can be signalled in two different ways: mechanically or electrically; mechanical signalling is obtained by push-pull rods, or more commonly by cables and pulleys; electrical signalling is a solution of more modern and sophisticated vehicles and will be later on discussed.

The basic principle of the hydraulic control is simple, but two aspects must be noticed when a powered control is introduced:

1. the system must control the surface in a proportional way, i.e. the surface response (deflection) must be function to the pilot’s demand (stick deflection, for instance);

2. the pilot that with little effort acts on a control valve must have a feedback on the manoeuvre intensity.

The first problem is solved by using (hydraulic) servo-mechanisms, where the components are linked in such a way to introduce an actuator stroke proportional to the pilot’s demand; many examples can be made, two of them are sketched in fig. 6.4, the second one including also the hydraulic circuit necessary for a correct operation.

In both cases the control valve housing is solid with the cylinder and the cabin column



Fig. 6.4 – Classic hydraulic servomechanisms

has a mechanical linkage to drive the valve spool.

In the first case, the cylinder is hinged to the aircraft and, due to valve spool displacement and ports opening, the piston is moved in one direction or the other; the piston rod is also linked to the valve spool stick, in such a way that the piston movement brings the spool back towards its neutral position; when this is reached, the actuator stops, then obtaining a deflection that is proportional to the demand.

In the second case the piston is constrained to the aircraft; the cabin column controls the valve spool stick; this will result in a movement of the cylinder, and this brings the valve housing again towards the valve neutral position, then resulting in a stroke proportional to the pilot’s demand. The hydraulic circuit also includes an emergency valve on the delivery segment to the control valve; if the delivery pressure drops, due for instance to a pump or engine failure, the emergency valve switches to the other position and links all the control valve inlets to the tank; this operation hydraulically unlocks the system, allowing the pilot for manual actuation of the cylinder.

It is clear now that the pilot, in normal hydraulic operating conditions, is requested for a very low effort, necessary to contrast the mechanical frictions of the linkage and the movement of the control valve: the pilot is then no more aware of the load condition being imposed to the aircraft.

For this reason an artificial feel is introduced in powered systems, acting directly on the cabin control stick or pedals. The simplest solution is a spring system, then responding to the pilot’s demand with a force proportional to the stick deflection; this solution has of course the limit to be not sensitive to the actual flight conditions. A more sophisticated artificial feel is the so-called Q feel. This system receives data from the pitot-static probes, reading the dynamic pressure, or the difference between total (pt) and static (ps) pressure, that is proportional to the aircraft speed v through the air density ρ:


This signal is used to modulate a hydraulic cylinder that increases the stiffness in the artificial feel system, in such a way that the pilot is given a contrast force in the pedals or stick that is also proportional to the aircraft speed.

Pneumatic control system

Pneumatic is a branch of technology, which deals with the study and application of pressurized gas to effect mechanical motion.

Pneumatic systems are extensively used in industry, where factories are commonly plumbed with compressed air or compressed inert gases. This is because a centrally located and electrically powered compressor that powers cylinders and other pneumatic devices through solenoid valves is often able to provide motive power in a cheaper, safer, more flexible, and more reliable way than a large number of electric motors and actuators.

Pneumatic also has applications in dentistry, construction, mining, and other areas.



Pump that compresses air, raising air pressure to above ambient pressure for use in pneumatic systems.

Check valve:

One-way valve – allows pressurized air to enter the pneumatic system, but prevents backflow of air toward the Compressor when Compressor is stopped (prevent loss of pressure.


· Stores compressed air,

· Prevents surges in pressure

· Prevents constant Compressor operation (“duty cycles” of Compressor)


Directional Valve: (Selector valve)

ü Controls pressurized air flow from Accumulator (source to user equipment via selected port

ü Some valves are one way – shut tight

ü Some valves are two way, allowing free exhaust from the port not selected

ü valves can be actuated manually or electrically.



ü Converts energy stored in compressed air into mechanical motion

ü Example is a linear piston (piston limited to moving in two opposing directions)

ü Other examples are alternate tools including: rotary actuators, air tools, expanding bladders, etc



Pneumatic uses in Aircraft

  • Powers engine Suction System for Heading indicators and Attitude indicators.
  • Actuates Landing Gear (some aircraft)
  • Emergency Brakes (some aircraft)
  • Cabin Pressure (for pressurized aircraft)



In the 70’s the fly-by-wire architecture was developed, starting as an analogue technique and later on, in most cases, transformed into digital. It was first developed for military aviation, where it is now a common solution; the supersonic Concorde can be considered a first and isolated civil aircraft equipped with a (analogue) fly-by-wire system, but in the 80’s the digital technique was imported from military into civil aviation by Airbus, first with the A320, then followed by A319, A321, A330, A340, Boeing 777 and A380 (scheduled for 2005).

This architecture is based on computer signal processing and is schematically shown in fig. 6.5: the pilot’s demand is first of all transduced into electrical signal in the cabin and sent to a group of independent computers (Airbus architecture substitute the cabin control column with a side stick); the computers sample also data concerning the flight conditions and servo-valves and actuators positions; the pilot’s demand is then processed and sent to the actuator, properly tailored to the actual flight status.

The flight data used by the system mainly depend on the aircraft category; in general the following data are sampled and processed:

• pitch, roll, yaw rate and linear accelerations

angle of attack and sideslip;

• airspeed/mach number, pressure altitude and radio altimeter indications;

• stick and pedal demands;

• other cabin commands such as landing gear condition, thrust lever position, etc.

The full system has high redundancy to restore the level of reliability of a mechanical or hydraulic system, in the form of multiple (triplex or quadruplex) parallel and independent lanes to generate and transmit the signals, and independent computers that process them; in many cases both hardware and software are different, to make the generation of a common error extremely remote, increase fault tolerance and isolation; in some cases the multiplexing of the digital computing and signal transmission is supported with an analogue or mechanical back-up system, to achieve adequate system reliability.


Fig. 6.5 – Fly-by-wire system

between military and civil aircraft; some of the most important benefits are as follows:

• flight envelope protection (the computers will reject and tune pilot’s demands that might exceed the airframe load factors);

• increase of stability and handling qualities across the full flight envelope, including the possibility of flying unstable vehicles;

• turbulence suppression and consequent decrease of fatigue loads and increase of passenger comfort;

• use of thrust vectoring to augment or replace lift aerodynamic control, then extending the aircraft flight envelope;

• drag reduction by an optimised trim setting;

• higher stability during release of tanks and weapons;

• easier interfacing to auto-pilot and other automatic flight control systems;

• weight reduction (mechanical linkages are substituted by wirings);

• maintenance reduction;

• reduction of airlines’ pilot training costs (flight handling becomes very similar in an whole aircraft family).

the flight mode: ground, take-off, flight and flare. Transition between modes is smooth and the pilot is not affected in its ability to control the aircraft: in ground mode the pilot has control on the nose wheel steering as a function of speed, after lift-off the envelope protection is gradually introduced and in flight mode the aircraft is fully protected by exceeding the maximum negative and positive load factors (with and without high lift devices extracted), angle of attack, stall, airspeed/Mach number, pitch attitude, roll rate, bank angle etc; finally, when the aircraft approaches to ground the control is gradually switched to flare mode, where automatic trim is deactivated and modified flight laws are used for pitch control.

The control software is one of the most critical aspects of fly-by-wire. It is developed in accordance to very strict rules, taking into account the flight control laws, and extensive testing is performed to reduce the probability of error. The risk of aircraft loss due to flight control failure is 2×10-6 per flight hour for a sophisticated military airplane, that anyway has the ejection seat as ultimate solution; the risk is reduced to 10-9 per flight hour for a civil airplane, were occupants cannot evacuate the airplane during flight.

Fig. 6.6 shows, as example, the fly-by-wire layout for the Airbus 340. Three groups of personal computers are used on board: three for primary control (FCPC), two for secondary control (FCSC) and two for high lift devices control (SFCC). The primary and secondary computers are based on different hardware; computers belonging to the same group have different software.

Two additional personal computers are used to store flight data.


Fig. 6.6 – A340 fly-by-wire layout, including hydraulic system indications

In the drawing the computer group and hydraulic system that control each surface are indicated (there are three independent hydraulic systems on the A340, commonly indicated as Blue, Yellow and Green). The leading edge flaps are linked together, and so are the trailing edge flaps, and then they are controlled by hydraulic units in the fuselage.

The drawing shows a considerable redundancy of the flight control system: the inboard and outboard ailerons, elevators and rudder are controlled by both the primary and secondary computers and operated by the three hydraulic sub-systems; the high lift devices are controlled by their specific computers and operated by the three hydraulic systems (Blue and Green for the leading edge,Yellow and Green for the trailing edge); the vertical stabiliser, having a secondary role, is controlled only by the secondary computers and operated by two hydraulic sub-systems. Thanks to this layout, first of all, in case of double hydraulic sub-system fault, the aircraft can be basically controlled with one hydraulic sub-system. Moreover, in case of total power black-out, the pilot can control the rudder and elevators by a mechanical back-up system, since the capability of this aircraft to land safely has been demonstrated with only limited pitch and yaw control.

Fly-by-wire architecture is inevitable for some aircraft categories: fig. 6.7 shows a typically unstable aircraft and a tilt rotor aircraft.


Northrop B-2


Bell-Boeing V-22

Fig. 6.7 – Need of fly-y-wire architecture for unstable (B-2) and thrust vectoring (V-22) airplanes

Auto pilot System

An autopilot is a mechanical, electrical, or hydraulic system used to guide a vehicle without assistance from a human being. An autopilot can refer specifically to aircraft, self-steering gear for boats, or auto guidance of space craft and missiles. The autopilot of an aircraft is sometimes referred to as “George”, after one of the key contributors to its development.

Today, autopilots are sophisticated systems that perform the same duties as a highly trained pilot. In fact, for some in-flight routines and procedures, autopilots are even better than a pair of human hands. They don’t just make flights smoother -they make them safer and more efficient. We’ll look at how autopilots work by examining their main components, how they work together — and what happens if they fail.

Autopilots and Avionics

In the world of aircraft, the autopilot is more accurately described as the automatic flight control system (AFCS). An AFCS is part of an aircraft’s avionics – the electronic systems, equipment and devices used to control key systems of the plane and its flight. In addition to flight control systems, avionics include electronics for communications, navigation, collision avoidance and weather. The original use of an AFCS was to provide pilot relief during tedious stages of flight, such as high-altitude cruising. Advanced autopilots can do much more, carrying out even highly precise maneuvers, such as landing an aircraft in conditions of zero visibility.

Although there is great diversity in autopilot systems, most can be classified according to the number of parts, or surfaces, they control. To understand this discussion, it helps to be familiar with the three basic control surfaces that affect an airplane’s attitude.

Autopilots can control any or all of these surfaces. A single-axis autopilot manages just one set of controls, usually the ailerons. This simple type of autopilot is known as a “wing leveler” because, by controlling roll, it keeps the aircraft wings on an even keel.

A two-axis autopilot manages elevators and ailerons. Finally, a three-axis autopilot manages all three basic control systems: ailerons, elevators and rudder.

The invention of autopilot

Famous inventor and engineer Elmer Sperry patented the gyrocompass in 1908, but it was his son, Lawrence Burst Sperry, who first flight-tested such a device in an aircraft. The younger Sperry’s autopilot used four gyroscopes to stabilize the airplane and led to many flying firsts, including the first night flight in the history of aviation. In 1932, the Sperry Gyroscope Company developed the automatic pilot that Wiley Post would use in his first solo flight around the world.

Autopilot Parts

The heart of a modern automatic flight control system is a computer with several high-speed processors. To gather the intelligence required to control the plane, the processors communicate with sensors located on the major control surfaces. They can also collect data from other airplane systems and equipment, including gyroscopes, accelerometers, altimeters, compasses and airspeed indicators.

The processors in the AFCS then take the input data and, using complex calculations, compare it to a set of control modes. A control mode is a setting entered by the pilot that defines a specific detail of the flight. For example, there is a control mode that defines how an aircraft’s altitude will be maintained. There are also control modes that maintain airspeed, heading and flight path.

These calculations determine if the plane is obeying the commands set up in the control modes. The processors then send signals to various servomechanism units. A servomechanism, or servo for short, is a device that provides mechanical control at a distance. One servo exists for each control surface included in the autopilot system. The servos take the computer’s instructions and use motors or hydraulics to move the craft’s control surfaces, making sure the plane maintains its proper course and attitude.


The above illustration shows how the basic elements of an autopilot system are related. For simplicity, only one control surface — the rudder — is shown, although each control surface would have a similar arrangement. Notice that the basic schematic of an autopilot looks like a loop, with sensors sending data to the autopilot computer, which processes the information and transmits signals to the servo, which moves the control surface, which changes the attitude of the plane, which creates a new data set in the sensors, which starts the whole process again. This type of feedback loop is central to the operation of autopilot systems.

Autopilot Control Systems

An autopilot is an example of a control system. Control systems apply an action based on a measurement and almost always have an impact on the value they are measuring. A classic example of a control system is the negative feedback loop that controls the thermostat in your home. Such a loop works like this:

1. Its summertime and a homeowner set his thermostat to a desired room temperature   say 78°F.

2. The thermostat measures the air temperature and compares it to the preset value.

3. Over time, the hot air outside the house will elevate the temperature inside the house. When the temperature inside exceeds 78°F, the thermostat sends a signal to the air conditioning unit.

4. The air conditioning unit clicks on and cools the room.

5. When the temperature in the room returns to 78°F, another signal is sent to the air conditioner, which shuts off.

It’s called a negative feedback loop because the result of a certain action (the air conditioning unit clicking on) inhibits further performance of that action. All negative feedback loops require a receptor, a control center and an effector. In the example above, the receptor is the thermometer that measures air temperature. The control center is the processor inside the thermostat. And the effector is the air conditioning unit.

Automated flight control systems work the same way. Let’s consider the example of a pilot who has activated a single-axis autopilot — the so-called wing leveler we mentioned earlier.


1. The pilot sets a control mode to maintain the wings in a level position.

2. However, even in the smoothest air, a wing will eventually dip.

3. Position sensors on the wing detect this deflection and send a signal to the autopilot computer.

4. The autopilot computer processes the input data and determines that the wings are no longer level.

5. The autopilot computer sends a signal to the servos that control the aircraft’s ailerons. The signal is a very specific command telling the servo to make a precise adjustment.

a) Each servo has a small electric motor fitted with a slip clutch that, through a bridle cable, grips the aileron cable. When the cable moves, the control surfaces move accordingly.

b) As the ailerons are adjusted based on the input data, the wings move back toward level.

c) The autopilot computer removes the command when the position sensor on the wing detects that the wings are once again level.

d) The servos cease to apply pressure on the aileron cables.

This loop, shown above in the block diagram, works continuously, many times a second, much more quickly and smoothly than a human pilot could. Two- and three-axis autopilots obey the same principles, employing multiple processors that control multiple surfaces. Some airplanes even have auto thrust computers to control engine thrust. Autopilot and auto thrust systems can work together to perform very complex maneuvers.

Autopilot Failure

Autopilots can and do fail. A common problem is some kind of servo failure, either because of a bad motor or a bad connection. A position sensor can also fail, resulting in a loss of input data to the autopilot computer. Fortunately, autopilots for manned aircraft are designed as a failsafe — that is, no failure in the automatic pilot can prevent effective employment of manual override. To override the autopilot, a crew member simply has to disengage the system, either by flipping a power switch or, if that doesn’t work, by pulling the autopilot circuit breaker.

Some airplane crashes have been blamed on situations where pilots have failed to disengage the automatic flight control system. The pilots end up fighting the settings that the autopilot is administering; unable to figure out why the plane won’t do what they’re asking it to do. This is why flight instruction programs stress practicing for just such a scenario. Pilots must know how to use every feature of an AFCS, but they must also know how to turn it off and fly without it. They also have to adhere to a rigorous maintenance schedule to make sure all sensors and servos are in good working order. Any adjustments or fixes in key systems may require that the autopilot be tweaked. For example, a change made to gyro instruments will require realignment of the settings in the autopilot’s computer.

Modern Autopilot Systems

Many modern autopilots can receive data from a Global Positioning System (GPS) receiver installed on the aircraft. A GPS receiver can determine airplane’s position in space by calculating its distance from three or more satellites in the GPS network. Armed with such positioning information, an autopilot can do more than keep a plane straight and level — it can execute a flight plan.

Most commercial jets have had such capabilities for a while, but even ­smaller planes are incorporating sophisticated autopilot systems. New Cessna 182s and 206s are leaving the factory with the Garmin G1000 integrated cockpit, which includes a digital electronic autopilot combined with a flight director. The Garmin G1000 delivers essentially all the capabilities and modes of a jet avionics system, bringing true automatic flight control to a new generation of general aviation planes.Wiley Post could have only dreamed of such technology back in 1933.