6-1

Aircraft Systems

Chapter 6

Introduction

This chapter covers the primary systems found on most 

aircraft. These include the engine, propeller, induction, 

ignition, as well as the fuel, lubrication, cooling, electrical, 

landing gear, and environmental control systems. 

Powerplant

An aircraft engine, or powerplant, produces thrust to propel 

an aircraft. Reciprocating engines and turboprop engines 

work in combination with a propeller to produce thrust. 

Turbojet and turbofan engines produce thrust by increasing 

the velocity of air flowing through the engine. All of these 

powerplants also drive the various systems that support the 

operation of an aircraft. 
6-2

Figure 6-1. Radial engine.

Horizontally opposed engine

Opposed cylinders

Figure 6-2. Horizontally opposed engine.

Reciprocating Engines

Most small aircraft are designed with reciprocating 

engines. The name is derived from the back-and-forth, or 

reciprocating, movement of the pistons which produces the 

mechanical energy necessary to accomplish work.

Driven by a revitalization of the general aviation (GA) 

industry and advances in both material and engine design, 

reciprocating engine technology has improved dramatically 

over the past two decades. The integration of computerized 

engine management systems has improved fuel efficiency, 

decreased emissions, and reduced pilot workload. 

Reciprocating engines operate on the basic principle of 

converting chemical energy (fuel) into mechanical energy. 

This conversion occurs within the cylinders of the engine 

through the process of combustion. The two primary 

reciprocating engine designs are the spark ignition and the 

compression ignition. The spark ignition reciprocating engine 

has served as the powerplant of choice for many years. In 

an effort to reduce operating costs, simplify design, and 

improve reliability, several engine manufacturers are turning 

to compression ignition as a viable alternative. Often referred 

to as jet fuel piston engines, compression ignition engines 

have the added advantage of utilizing readily available and 

lower cost diesel or jet fuel. 

The main mechanical components of the spark ignition and 

the compression ignition engine are essentially the same. 

Both use cylindrical combustion chambers and pistons that 

travel the length of the cylinders to convert linear motion 

into the rotary motion of the crankshaft. The main difference 

between spark ignition and compression ignition is the 

process of igniting the fuel. Spark ignition engines use a 

spark plug to ignite a pre-mixed fuel/air mixture. (Fuel/air 

mixture is the ratio of the “weight” of fuel to the “weight” 

of air in the mixture to be burned.) A compression ignition 

engine first compresses the air in the cylinder, raising its 

temperature to a degree necessary for automatic ignition 

when fuel is injected into the cylinder. 

These two engine designs can be further classified as:

1. Cylinder arrangement with respect to the crankshaft—

radial, in-line, v-type, or opposed.

2. Operating cycle—two or four.

3. Method of cooling—liquid or air. 

Radial engines were widely used during World War II and 

many are still in service today. With these engines, a row or 

rows of cylinders are arranged in a circular pattern around 

the crankcase. The main advantage of a radial engine is the 

favorable power-to-weight ratio. [Figure 6-1]

In-line engines have a comparatively small frontal area, but 

their power-to-weight ratios are relatively low. In addition, 

the rearmost cylinders of an air-cooled, in-line engine 

receive very little cooling air, so these engines are normally 

limited to four or six cylinders. V-type engines provide 

more horsepower than in-line engines and still retain a small 

frontal area. 

Continued improvements in engine design led to the 

development of the horizontally-opposed engine which 

remains the most popular reciprocating engines used on 

smaller aircraft. These engines always have an even number 

of cylinders, since a cylinder on one side of the crankcase 

“opposes” a cylinder on the other side. [Figure 6-2] The 

majority of these engines are air cooled and usually are 

mounted in a horizontal position when installed on fixed-wing 

airplanes. Opposed-type engines have high power-to-weight 

ratios because they have a comparatively small, lightweight 

crankcase. In addition, the compact cylinder arrangement 

reduces the engine’s frontal area and allows a streamlined 

installation that minimizes aerodynamic drag. 
6-3

Exhaust valve Fuel injector

Piston

Forced air

2. Power stroke

1. Intake/compression

and exhaust

Figure 6-3. Two-stroke compression ignition.

Exhaust valve

Intake valve

Spark plug

Piston

Connecting rod Crankshaft

Crankcase

Cylinder

Figure 6-4. Main components of a spark ignition reciprocating 

engine.

Depending on the engine manufacturer, all of these 

arrangements can be designed to utilize spark or compression 

ignition, and operate on either a two- or four-stroke cycle.

In a two-stroke engine, the conversion of chemical energy 

into mechanical energy occurs over a two-stroke operating 

cycle. The intake, compression, power, and exhaust processes 

occur in only two strokes of the piston rather than the more 

common four strokes. Because a two-stroke engine has a 

power stroke each revolution of the crankshaft, it typically has 

higher power-to-weight ratio than a comparable four-stroke 

engine. Due to the inherent inefficiency and disproportionate 

emissions of the earliest designs, use of the two-stroke engine 

has been limited in aviation. 

Recent advances in material and engine design have 

reduced many of the negative characteristics associated 

with two-stroke engines. Modern two-stroke engines often 

use conventional oil sumps, oil pumps and full pressure fed 

lubrication systems. The use of direct fuel injection and 

pressurized air, characteristic of advanced compression 

ignition engines, make two-stroke compression ignition 

engines a viable alternative to the more common four-stroke 

spark ignition designs. [Figure 6-3] 

Spark ignition four-stroke engines remain the most common 

design used in general aviation today. [Figure 6-4] The 

main parts of a spark ignition reciprocating engine include 

the cylinders, crankcase, and accessory housing. The intake/

exhaust valves, spark plugs, and pistons are located in the 

cylinders. The crankshaft and connecting rods are located 

in the crankcase. The magnetos are normally located on the 

engine accessory housing.

In a four-stroke engine the conversion of chemical energy 

into mechanical energy occurs over a four stroke operating 

cycle. The intake, compression, power, and exhaust processes 

occur in four separate strokes of the piston. 

1. The intake stroke begins as the piston starts its 

downward travel. When this happens, the intake 

valve opens and the fuel/air mixture is drawn into the 

cylinder.

2. The compression stroke begins when the intake valve 

closes and the piston starts moving back to the top of 

the cylinder. This phase of the cycle is used to obtain 

a much greater power output from the fuel/air mixture 

once it is ignited.

3. The power stroke begins when the fuel/air mixture is 

ignited. This causes a tremendous pressure increase 

in the cylinder, and forces the piston downward away 

from the cylinder head, creating the power that turns 

the crankshaft.
6-4

1. Intake 2. Compression

3. Power 4. Exhaust

Intake valve Exhaust valve

Piston Spark plug

Crankshaft Connecting rod

Figure 6-5. The arrows in this illustration indicate the direction of 

motion of the crankshaft and piston during the four-stroke cycle.

4. The exhaust stroke is used to purge the cylinder of 

burned gases. It begins when the exhaust valve opens 

and the piston starts to move toward the cylinder head 

once again.

Even when the engine is operated at a fairly low speed, 

the four-stroke cycle takes place several hundred times 

each minute. [Figure 6-5] In a four-cylinder engine, each 

cylinder operates on a different stroke. Continuous rotation 

of a crankshaft is maintained by the precise timing of the 

power strokes in each cylinder. Continuous operation of the 

engine depends on the simultaneous function of auxiliary 

systems, including the induction, ignition, fuel, oil, cooling, 

and exhaust systems.

The latest advance in aircraft reciprocating engines was 

pioneered in the mid-1960s by Frank Thielert, who looked 

to the automotive industry for answers on how to integrate 

diesel technology into an aircraft engine. The advantage 

of a diesel-fueled reciprocating engine lies in the physical 

similarity of diesel and kerosene. Aircraft equipped with a 

diesel piston engine runs on standard aviation fuel kerosene 

which provides more independence, higher reliability, lower 

consumption, and operational cost saving. 

In 1999, Thielert formed Thielert Aircraft Engines (TAE) 

to design, develop, certify, and manufacture a brand-new 

Jet-A-burning diesel cycle engine (also known as jet-fueled 

piston engine) for the GA industry. By March 2001, the first 

prototype engine became the first certified diesel engine 

since World War II. TAE continues to design and develop 

diesel cycle engines and other engine manufacturers such as 

Société de Motorisations Aéronautiques (SMA) now offer 

jet-fueled piston engines as well. TAE engines can be found 

on the Diamond DA40 single and the DA42 Twin Star, the 

first diesel engine to be part of the type certificate of a new 

original equipment manufacturer (OEM) aircraft. 

These engines have also gained a toehold in the retrofit 

market with a supplemental type certificate (STC) to reengine the Cessna 172 models and the Piper PA-28 family. 

The jet-fueled piston engines technology has continued to 

progress and a full authority digital engine control (FADEC, 

discussed more fully later in the chapter) is standard on such 

equipped aircraft which minimizes complication of engine 

control. By 2007, various jet-fueled piston aircraft had logged 

well over 600,000 hours of service. 

Propeller

The propeller is a rotating airfoil, subject to induced drag, 

stalls, and other aerodynamic principles that apply to any 

airfoil. It provides the necessary thrust to pull, or in some 

cases push, the aircraft through the air. The engine power is 

used to rotate the propeller, which in turn generates thrust 

very similar to the manner in which a wing produces lift. 

The amount of thrust produced depends on the shape of the 

airfoil, the angle of attack of the propeller blade, and the 

revolutions per minute (rpm) of the engine. The propeller 

itself is twisted so the blade angle changes from hub to tip. 

The greatest angle of incidence, or the highest pitch, is at the 

hub while the smallest angle of incidence or smallest pitch 

is at the tip. [Figure 6-6]

The reason for the twist is to produce uniform lift from the 

hub to the tip. As the blade rotates, there is a difference in 

the actual speed of the various portions of the blade. The tip 

of the blade travels faster than the part near the hub, because 

the tip travels a greater distance than the hub in the same 

length of time. [Figure 6-7] Changing the angle of incidence 

(pitch) from the hub to the tip to correspond with the speed 

produces uniform lift throughout the length of the blade. A 
6-5

60 in.

40 in.

20 in.

Short travel distance—slow speed

—129 kno st

Moderate travel distance—moderate speed—259 knots

Greater travel distance—very high speed—389 knots

2,500 rpm

2,500 rpm

2,500 rpm

Figure 6-6. Changes in propeller blade angle from hub to tip.

Figure 6-7. Relationship of travel distance and speed of various 

portions of propeller blade.

Figure 6-8. Engine rpm is indicated on the tachometer.

Engine r.p.m. is indicated on the tachometer

RPM HUNDREDS

25

3

5

I5

30

20

I0

35

25

3

5

I5

30

20

I0

35 HOURS

AVOID CONTINUOUS

OPERATION BETWEEN 2250 AND 2350 RPM

I

I0

propeller blade designed with the same angle of incidence 

throughout its entire length would be inefficient because as 

airspeed increases in flight, the portion near the hub would 

have a negative angle of attack while the blade tip would 

be stalled.

Small aircraft are equipped with either one of two types 

of propellers. One is the fixed pitch, and the other is the 

adjustable pitch.

Fixed-Pitch Propeller

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, which 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 
6-6

IN Hg

ALg.

MANIFOLD

PRESS

35

10 50

15

25

40

30

20

45

35

10 50

15

25

40

30

20

45

P

Figure 6-9. Engine power output is indicated on the manifold 

pressure gauge.

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.

Adjustable-Pitch Propeller

The adjustable-pitch propeller was the forerunner of the 

constant-speed propeller. It is a propeller with blades whose 

pitch can be adjusted on the ground with the engine not 

running, but which cannot be adjusted in flight. It is also 

referred to as a ground adjustable propeller. By the 1930s, 

pioneer aviation inventors were laying the ground work for 

automatic pitch-change mechanisms, which is why the term 

sometimes refers to modern constant-speed propellers that 

are adjustable in flight.

The first adjustable-pitch propeller systems provided only 

two pitch settings: low and high. Today, most adjustable-pitch 

propeller systems are capable of a range of pitch settings.

A constant-speed propeller is a controllable-pitch propeller 

whose pitch is automatically varied in flight by a governor 

maintaining constant rpm despite varying air loads. It is the 

most common type of adjustable-pitch propeller. The main 

advantage of a constant-speed propeller is that it converts 

a high percentage of brake horsepower (BHP) into thrust 

horsepower (THP) over a wide range of rpm and airspeed 

combinations. A constant-speed propeller is more efficient 

than other propellers because it allows selection of the most 

efficient engine rpm for the given conditions.

An aircraft with a constant-speed propeller has two controls: 

the throttle and the propeller control. The throttle controls 

power output and the propeller control regulates engine rpm. 

This in turn regulates propeller rpm which is registered on 

the tachometer.

Once a specific rpm is selected, a governor automatically 

adjusts the propeller blade angle as necessary to maintain 

the selected rpm. For example, after setting the desired rpm 

during cruising flight, an increase in airspeed or decrease in 

propeller load will cause the propeller blade angle to increase 

as necessary to maintain the selected rpm. A reduction in 

airspeed or increase in propeller load will cause the propeller 

blade angle to decrease.

The propeller’s constant-speed range, defined by the high 

and low pitch stops, is the range of possible blade angles for 

a constant-speed propeller. As long as the propeller blade 

angle is within the constant-speed range and not against 

either pitch stop, a constant engine rpm will be maintained. 

If the propeller blades contact a pitch stop, the engine rpm 

will increase or decrease as appropriate, with changes in 

airspeed and propeller load. For example, once a specific 

rpm has been selected, if aircraft speed decreases enough to 

rotate the propeller blades until they contact the low pitch 

stop, any further decrease in airspeed will cause engine rpm 

to decrease the same way as if a fixed-pitch propeller were 

installed. The same holds true when an aircraft equipped with 

a constant-speed propeller accelerates to a faster airspeed. As 

the aircraft accelerates, the propeller blade angle increases to 

maintain the selected rpm until the high pitch stop is reached. 

Once this occurs, the blade angle cannot increase any further 

and engine rpm increases.

On aircraft equipped with a constant-speed propeller, power 

output is controlled by the throttle and indicated by a manifold 

pressure gauge. The gauge measures the absolute pressure of 

the fuel/air mixture inside the intake manifold and is more 

correctly a measure of manifold absolute pressure (MAP). At 

a constant rpm and altitude, the amount of power produced 

is directly related to the fuel/air flow being delivered to the 

combustion chamber. As the throttle setting is increased, 

more fuel and air flows to the engine and MAP increases. 

When the engine is not running, the manifold pressure gauge 

indicates ambient air pressure (i.e., 29.92 inches mercury 

(29.92 "Hg)). When the engine is started, the manifold 

pressure indication will decrease to a value less than ambient 

pressure (i.e., idle at 12 "Hg). Engine failure or power loss is 

indicated on the manifold gauge as an increase in manifold 

pressure to a value corresponding to the ambient air pressure 

at the altitude where the failure occurred. [Figure 6-9]
6-7

The manifold pressure gauge is color coded to indicate the 

engine’s operating range. The face of the manifold pressure 

gauge contains a green arc to show the normal operating 

range, and a red radial line to indicate the upper limit of 

manifold pressure.

For any given rpm, there is a manifold pressure that should not 

be exceeded. If manifold pressure is excessive for a given rpm, 

the pressure within the cylinders could be exceeded, placing 

undue stress on the cylinders. If repeated too frequently, this 

stress can weaken the cylinder components and eventually 

cause engine failure. As a general rule, manifold pressure 

(inches) should be less than the rpm.

A pilot can avoid conditions that overstress the cylinders by 

being constantly aware of the rpm, especially when increasing 

the manifold pressure. Conform to the manufacturer’s 

recommendations for power settings of a particular engine to 

maintain the proper relationship between manifold pressure 

and rpm.

When both manifold pressure and rpm need to be changed, 

avoid engine overstress by making power adjustments in 

the proper order:

• When power settings are being decreased, reduce 

manifold pressure before reducing rpm. If rpm is 

reduced before manifold pressure, manifold pressure 

will automatically increase, possibly exceeding the 

manufacturer’s tolerances.

• When power settings are being increased, reverse the 

order—increase rpm first, then manifold pressure.

• To prevent damage to radial engines, minimize 

operating time at maximum rpm and manifold 

pressure, and avoid operation at maximum rpm and 

low manifold pressure. 

The engine and/or airframe manufacturer’s recommendations 

should be followed to prevent severe wear, fatigue, and 

damage to high-performance reciprocating engines.

Induction Systems

The induction system brings in air from the outside, mixes 

it with fuel, and delivers the fuel/air mixture to the cylinder 

where combustion occurs. Outside air enters the induction 

system through an intake port on the front of the engine 

cowling. This port normally contains an air filter that inhibits 

the entry of dust and other foreign objects. Since the filter 

may occasionally become clogged, an alternate source of 

air must be available. Usually, the alternate air comes from 

inside the engine cowling, where it bypasses a clogged air 

filter. Some alternate air sources function automatically, 

while others operate manually.

Two types of induction systems are commonly used in small 

aircraft engines:

1. The carburetor system, which mixes the fuel and air 

in the carburetor before this mixture enters the intake 

manifold

2. The fuel injection system, which mixes the fuel and air 

immediately before entry into each cylinder or injects 

fuel directly into each cylinder

Carburetor Systems

Carburetors are classified as either float type or pressure type. 

The float type of carburetor, complete with idling, accelerating, 

mixture control, idle cutoff, and power enrichment systems is 

probably the most common of all carburetor types. Pressure 

carburetors are usually not found on small aircraft. The basic 

difference between a float-type and a pressure-type carburetor 

is the delivery of fuel. The pressure-type carburetor delivers 

fuel under pressure by a fuel pump. 

In the operation of the float-type carburetor system, the 

outside air first flows through an air filter, usually located 

at an air intake in the front part of the engine cowling. This 

filtered air flows into the carburetor and through a venturi, a 

narrow throat in the carburetor. When the air flows through 

the venturi, a low-pressure area is created, which forces the 

fuel to flow through a main fuel jet located at the throat. The 

fuel then flows into the airstream where it is mixed with the 

flowing air. [Figure 6-10]

The fuel/air mixture is then drawn through the intake manifold 

and into the combustion chambers where it is ignited. The 

float-type carburetor acquires its name from a float, which 

rests on fuel within the float chamber. A needle attached to 

the float opens and closes an opening at the bottom of the 

carburetor bowl. This meters the correct amount of fuel into 

the carburetor, depending upon the position of the float, 

which is controlled by the level of fuel in the float chamber. 

When the level of the fuel forces the float to rise, the needle 

valve closes the fuel opening and shuts off the fuel flow to 

the carburetor. The needle valve opens again when the engine 

requires additional fuel. The flow of the fuel/air mixture to 

the combustion chambers is regulated by the throttle valve, 

which is controlled by the throttle in the flight deck.

The float-type carburetor has several distinct disadvantages. 

In the first place, imagine the effect that abrupt maneuvers 

have on the float action. In the second place, the fact that its 

fuel must be discharged at low pressure leads to incomplete 

vaporization and difficulty in discharging fuel into some 

types of supercharged systems. The chief disadvantage of 

the float carburetor, however, is its icing tendency. Since 

the float carburetor must discharge fuel at a point of low 
6-8

Float-type carburetor

FUEL

Fuel level is maintained

by a float-type device.

FLOAT CHAMBER

Fuel is received into

the carburetor through

the fuel inlet.

FUEL INLET

The mixture needle 

controls fuel to the 

discharge nozzle. 

Mixture needle position 

can be adjusted using 

the mixture control.

MIXTURE NEEDLE

Air enters the carburetor

through the air inlet.

AIR INLET

The blend of fuel and 

air is routed to the 

combustion chambers 

to be burned. 

FUEL/AIR MIXTURE

The flow of the fuel/air 

mixture is controlled by

the throttle valve. The

throttle valve is adjusted 

from the flight deck by

the throttle.

THROTTLE VALVE

Fuel is forced through 

the discharge nozzle 

into the venturi by

greater atmospheric

pressure in the float 

chamber.

DISCHARGE NOZZLE

The shape of the venturi 

creates an area of low 

pressure.

VENTURI

The air bleed allows air to be mixed 

with fuel being drawn out of the discharge nozzle to decrease fuel density 

and promote fuel vaporization.

AIR BLEED

Figure 6-10. Float-type carburetor.

pressure, the discharge nozzle must be located at the venturi 

throat, and the throttle valve must be on the engine side of the 

discharge nozzle. This means the drop in temperature due to 

fuel vaporization takes place within the venturi. As a result, 

ice readily forms in the venturi and on the throttle valve.

A pressure-type carburetor discharges fuel into the airstream 

at a pressure well above atmospheric. This results in better 

vaporization and permits the discharge of fuel into the 

airstream on the engine side of the throttle valve. With the 

discharge nozzle located at this point, the drop in temperature 

due to fuel vaporization takes place after the air has passed the 

throttle valve and at a point where engine heat tends to offset 

it. Thus, the danger of fuel vaporization icing is practically 

eliminated. The effects of rapid maneuvers and rough air on 

the pressure-type carburetors are negligible since their fuel 

chambers remain filled under all operating conditions

Mixture Control

Carburetors are normally calibrated at sea-level pressure, 

where the correct fuel-to-air mixture ratio is established 

with the mixture control set in the FULL RICH position. 

However, as altitude increases, the density of air entering the 

carburetor decreases, while the density of the fuel remains 

the same. This creates a progressively richer mixture, which 

can result in engine roughness and an appreciable loss of 

power. The roughness normally is due to spark plug fouling 

from excessive carbon buildup on the plugs. Carbon buildup 

occurs because the rich mixture lowers the temperature 

inside the cylinder, inhibiting complete combustion of the 

fuel. This condition may occur during the pretakeoff runup 

at high-elevation airports and during climbs or cruise flight 

at high altitudes. To maintain the correct fuel/air mixture, the 

mixture must be leaned using the mixture control. Leaning 

the mixture decreases fuel flow, which compensates for the 

decreased air density at high altitude.

During a descent from high altitude, the mixture must 

be enriched, or it may become too lean. An overly lean 

mixture causes detonation, which may result in rough 

engine operation, overheating, and a loss of power. The 

best way to maintain the proper mixture is to monitor the 

engine temperature and enrich the mixture as needed. Proper 

mixture control and better fuel economy for fuel-injected 

engines can be achieved by use of an exhaust gas temperature 

(EGT) gauge. Since the process of adjusting the mixture can 

vary from one aircraft to another, it is important to refer to 

the airplane flight manual (AFM) or the pilot’s operating 

handbook (POH) to determine the specific procedures for 

a given aircraft.

Carburetor Icing

As mentioned earlier, one disadvantage of the float-type 

carburetor is its icing tendency. Carburetor ice occurs due 
6-9

Incoming air

Venturi

Ice

To engine Fuel/air mixture

Ice

Ice

Figure 6-11. The formation of carburetor ice may reduce or block 

fuel/air flow to the engine.

Relative humidity

Outside air temperature

100%

50%

80%

60%

70%

90%

20°F/-7°C 32°F/0°C 70°F/21°C 100°F/38°C

High carburetor

icing potential

Carburetor icing possible

Figure 6-12. Although carburetor ice is most likely to form when 

the temperature and humidity are in ranges indicated by this chart, 

carburetor ice is possible under conditions not depicted.

by a decrease in manifold pressure, but no reduction in rpm. 

Propeller pitch is automatically adjusted to compensate for 

loss of power. Thus, a constant rpm is maintained. Although 

carburetor ice can occur during any phase of flight, it is 

particularly dangerous when using reduced power during 

a descent. Under certain conditions, carburetor ice could 

build unnoticed until power is added. To combat the effects 

of carburetor ice, engines with float-type carburetors employ 

a carburetor heat system.

Carburetor Heat

Carburetor heat is an anti-icing system that preheats the air 

before it reaches the carburetor, and is intended to keep the 

fuel/air mixture above the freezing temperature to prevent 

the formation of carburetor ice. Carburetor heat can be used 

to melt ice that has already formed in the carburetor if the 

accumulation is not too great, but using carburetor heat as a 

preventative measure is the better option. Additionally, the 

use of carburetor heat as an alternate air source can be used 

if the intake filter clogs such as in sudden or unexpected 

airframe icing conditions. The carburetor heat should be 

checked during the engine runup. When using carburetor 

heat, follow the manufacturer’s recommendations.

When conditions are conducive to carburetor icing during 

flight, periodic checks should be made to detect its presence. If 

detected, full carburetor heat should be applied immediately, 

and it should be left in the ON position until the pilot is 

certain all the ice has been removed. If ice is present, applying 

partial heat or leaving heat on for an insufficient time might 

aggravate the situation. In extreme cases of carburetor icing, 

even after the ice has been removed, full carburetor heat 

should be used to prevent further ice formation. If installed, 

a carburetor temperature gauge is useful in determining when 

to use carburetor heat.

to the effect of fuel vaporization and the decrease in air 

pressure in the venturi, which causes a sharp temperature 

drop in the carburetor. If water vapor in the air condenses 

when the carburetor temperature is at or below freezing, ice 

may form on internal surfaces of the carburetor, including 

the throttle valve. [Figure 6-11] 

The reduced air pressure, as well as the vaporization of fuel, 

contributes to the temperature decrease in the carburetor. Ice 

generally forms in the vicinity of the throttle valve and in the 

venturi throat. This restricts the flow of the fuel/air mixture 

and reduces power. If enough ice builds up, the engine may 

cease to operate. Carburetor ice is most likely to occur when 

temperatures are below 70 degrees Fahrenheit (°F) or 21 

degrees Celsius (°C) and the relative humidity is above 80 

percent. Due to the sudden cooling that takes place in the 

carburetor, icing can occur even with temperatures as high 

as 100 °F (38 °C) and humidity as low as 50 percent. This 

temperature drop can be as much as 60 to 70 °F (15 to 21 °C). 

Therefore, at an outside air temperature of 100 °F (37 °C), a 

temperature drop of 70 °F (21 °C) results in an air temperature 

in the carburetor of 30 °F (-1 °C). [Figure 6-12] 

The first indication of carburetor icing in an aircraft with 

a fixed-pitch propeller is a decrease in engine rpm, which 

may be followed by engine roughness. In an aircraft with a 

constant-speed propeller, carburetor icing is usually indicated 
6-10

Usually, the face of the gauge is calibrated in degrees Celsius, 

with a yellow arc indicating the carburetor air temperatures 

where icing may occur. This yellow arc typically ranges 

between -15 °C and +5 °C (5 °F and 41 °F). If the air 

temperature and moisture content of the air are such that 

carburetor icing is improbable, the engine can be operated 

with the indicator in the yellow range with no adverse effects. 

If the atmospheric conditions are conducive to carburetor 

icing, the indicator must be kept outside the yellow arc by 

application of carburetor heat.

Certain carburetor air temperature gauges have a red radial, 

which indicates the maximum permissible carburetor inlet 

air temperature recommended by the engine manufacturer. If 

present, a green arc indicates the normal operating range.

Outside Air Temperature Gauge

Most aircraft are also equipped with an outside air 

temperature (OAT) gauge calibrated in both degrees Celsius 

and Fahrenheit. It provides the outside or ambient air 

temperature for calculating true airspeed, and also is useful 

in detecting potential icing conditions.

Fuel Injection Systems

In a fuel injection system, the fuel is injected directly into 

the cylinders, or just ahead of the intake valve. The air 

intake for the fuel injection system is similar to that used 

in a 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 six 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 6-13]

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.

Whenever the throttle is closed during flight, the engine cools 

rapidly and vaporization of the fuel is less complete than if 

the engine is warm. Also, in this condition, the engine is more 

susceptible to carburetor icing. If carburetor icing conditions 

are suspected and closed-throttle operation anticipated, adjust 

the carburetor heat to the full ON position before closing the 

throttle and leave it on during the closed-throttle operation. 

The heat will aid in vaporizing the fuel and help prevent the 

formation of carburetor ice. Periodically, open the throttle 

smoothly for a few seconds to keep the engine warm; 

otherwise, the carburetor heater may not provide enough 

heat to prevent icing.

The use of carburetor heat causes a decrease in engine 

power, sometimes up to 15 percent, because the heated air 

is less dense than the outside air that had been entering the 

engine. This enriches the mixture. When ice is present in an 

aircraft with a fixed-pitch propeller and carburetor heat is 

being used, there is a decrease in rpm, followed by a gradual 

increase in rpm as the ice melts. The engine also should run 

more smoothly after the ice has been removed. If ice is not 

present, the rpm will decrease and then remain constant. 

When carburetor heat is used on an aircraft with a constantspeed propeller and ice is present, a decrease in the manifold 

pressure will be noticed, followed by a gradual increase. 

If carburetor icing is not present, the gradual increase in 

manifold pressure will not be apparent until the carburetor 

heat is turned off.

It is imperative for a pilot to recognize carburetor ice when 

it forms during flight because a loss of power, altitude, 

and/or airspeed will occur. These symptoms may sometimes 

be accompanied by vibration or engine roughness. Once a 

power loss is noticed, immediate action should be taken to 

eliminate ice already formed in the carburetor, and to prevent 

further ice formation. This is accomplished by applying full 

carburetor heat, which will cause a further reduction in power, 

and possibly engine roughness as melted ice goes through the 

engine. These symptoms may last from 30 seconds to several 

minutes, depending on the severity of the icing. During this 

period, the pilot must resist the temptation to decrease the 

carburetor heat usage. Carburetor heat must remain in the 

full-hot position until normal power returns.

Since the use of carburetor heat tends to reduce the output 

of the engine and to increase the operating temperature, 

carburetor heat should not be used when full power is required 

(as during takeoff) or during normal engine operation, except 

to check for the presence or to remove carburetor ice.

Carburetor Air Temperature Gauge

Some aircraft are equipped with a carburetor air temperature 

gauge, which is useful in detecting potential icing conditions. 
6-11

Fuel tank

Auxiliary fuel pump

Engine-driven fuel pump

Fuel/air control unit

Fuel lines Fuel manifold valve

Figure 6-13. Fuel injection system.

A fuel injection system is considered to be less susceptible to 

icing than the carburetor system, but impact icing on the air 

intake is a possibility in either system. Impact icing occurs 

when ice forms on the exterior of the aircraft, and blocks 

openings such as the air intake for the injection system.

Advantages of fuel injection:

• Reduction in evaporative icing

• Better fuel flow

• Faster throttle response

• Precise control of mixture

• Better fuel distribution

• Easier cold weather starts

Disadvantages:

• 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

Superchargers and Turbosuperchargers

To increase an engine’s horsepower, manufacturers have 

developed forced induction systems called supercharger 

and turbosupercharger systems. They both compress the 

intake air to increase its density. The key difference lies in 

the power supply. A supercharger relies on an engine-driven 

air pump or compressor, while a turbocharger gets its power 

from the exhaust stream that runs through a turbine, which in 

turn spins the compressor. Aircraft with these systems have 

a manifold pressure gauge, which displays MAP within the 

engine’s intake manifold.

On a standard day at sea level with the engine shut down, 

the manifold pressure gauge will indicate the ambient 

absolute air pressure of 29.92 "Hg. Because atmospheric 

pressure decreases approximately 1 "Hg per 1,000 feet of 

altitude increase, the manifold pressure gauge will indicate 

approximately 24.92 "Hg at an airport that is 5,000 feet above 

sea level with standard day conditions.

As a normally aspirated aircraft climbs, it eventually reaches 

an altitude where the MAP is insufficient for a normal climb. 

That altitude limit is the aircraft’s service ceiling, and it is 

directly affected by the engine’s ability to produce power. 

If the induction air entering the engine is pressurized, or 

boosted, by either a supercharger or a turbosupercharger, 

the aircraft’s service ceiling can be increased. With these 

systems, an aircraft can fly at higher altitudes with the 

advantage of higher true airspeeds and the increased ability 

to circumnavigate adverse weather.
6-12

Figure 6-14. Power output of normally aspirated engine compared 

to a single-stage, two-speed supercharged engine. Brake horsepower

Sea level Density altitude

Two-speed supercharged engine

High blower

Low blower 

Normally aspirated engine

Superchargers

A supercharger is an engine-driven air pump or compressor 

that provides compressed air to the engine to provide 

additional pressure to the induction air so the engine can 

produce additional power. It increases manifold pressure 

and forces the fuel/air mixture into the cylinders. The higher 

the manifold pressure, the more dense the fuel/air mixture, 

and the more power an engine can produce. With a normally 

aspirated engine, it is not possible to have manifold pressure 

higher than the existing atmospheric pressure. A supercharger 

is capable of boosting manifold pressure above 30 "Hg.

For example, at 8,000 feet a typical engine may be able to 

produce 75 percent of the power it could produce at mean sea 

level (MSL) because the air is less dense at the higher altitude. 

The supercharger compresses the air to a higher density 

allowing a supercharged engine to produce the same manifold 

pressure at higher altitudes as it could produce at sea level. 

Thus, an engine at 8,000 feet MSL could still produce 25 "Hg 

of manifold pressure whereas without a supercharger it could 

produce only 22 "Hg. Superchargers are especially valuable 

at high altitudes (such as 18,000 feet) where the air density 

is 50 percent that of sea level. The use of a supercharger in 

many cases will supply air to the engine at the same density 

it did at sea level. With a normally aspirated engine, it is not 

possible to have manifold pressure higher than the existing 

atmospheric pressure. A supercharger is capable of boosting 

manifold pressure above 30 "Hg.

The components in a supercharged induction system are similar 

to those in a normally aspirated system, with the addition of 

a supercharger between the fuel metering device and intake 

manifold. A supercharger is driven by the engine through a 

gear train at one speed, two speeds, or variable speeds. In 

addition, superchargers can have one or more stages. Each 

stage also provides an increase in pressure and superchargers 

may be classified as single stage, two stage, or multistage, 

depending on the number of times compression occurs.

An early version of a single-stage, single-speed supercharger 

may be referred to as a sea-level supercharger. An engine 

equipped with this type of supercharger is called a sea-level 

engine. With this type of supercharger, a single gear-driven 

impeller is used to increase the power produced by an engine at 

all altitudes. The drawback with this type of supercharger is a 

decrease in engine power output with an increase in altitude.

Single-stage, single-speed superchargers are found on many 

high-powered radial engines and use an air intake that faces 

forward so the induction system can take full advantage of the 

ram air. Intake air passes through ducts to a carburetor, where 

fuel is metered in proportion to the airflow. The fuel/air charge 

is then ducted to the supercharger, or blower impeller, which 

accelerates the fuel/air mixture outward. Once accelerated, the 

fuel/air mixture passes through a diffuser, where air velocity 

is traded for pressure energy. After compression, the resulting 

high pressure fuel/air mixture is directed to the cylinders.

Some of the large radial engines developed during World 

War II have a single-stage, two-speed supercharger. With 

this type of supercharger, a single impeller may be operated 

at two speeds. The low impeller speed is often referred to 

as the low blower setting, while the high impeller speed is 

called the high blower setting. On engines equipped with a 

two-speed supercharger, a lever or switch in the flight deck 

activates an oil-operated clutch that switches from one speed 

to the other.

Under normal operations, takeoff is made with the 

supercharger in the low blower position. In this mode, the 

engine performs as a ground-boosted engine, and the power 

output decreases as the aircraft gains altitude. However, once 

the aircraft reaches a specified altitude, a power reduction is 

made, and the supercharger control is switched to the high 

blower position. The throttle is then reset to the desired 

manifold pressure. An engine equipped with this type of 

supercharger is called an altitude engine. [Figure 6-14]

Turbosuperchargers

The most efficient method of increasing horsepower in an 

engine is by use of a turbosupercharger or turbocharger. 

Installed on an engine, this booster uses the engine’s exhaust 

gases to drive an air compressor to increase the pressure of 

the air going into the engine through the carburetor or fuel 

injection system to boost power at higher altitude.
6-13

EXHAUST GAS DISCHARGE

Pressurized air from the

turbocharger is supplied

to the cylinders.

INTAKE MANIFOLD

Intake air is ducted to the

turbocharger where it is

compressed.

AIR INTAKE This controls the amount 

of exhaust through the 

turbine. Waste gate 

position is actuated by

engine oil pressure.

WASTE GATE

The turbocharger incorporates a turbine, 

which is driven by exhaust gases, and a compressor that pressurizes

the incoming air. 

TURBOCHARGER

This regulates airflow 

to the engine.

THROTTLE BODY

Exhaust gas is ducted 

through the exhaust manifold and is used to turn 

the turbine which drives 

the compressor.

EXHAUST MANIFOLD

Figure 6-15. Turbocharger components.

The major disadvantage of the gear-driven supercharger––use 

of a large amount of the engine’s power output for the amount 

of power increase produced––is avoided with a turbocharger, 

because turbochargers are powered by an engine’s exhaust 

gases. This means a turbocharger recovers energy from hot 

exhaust gases that would otherwise be lost.

A second advantage of turbochargers over superchargers is 

the ability to maintain control over an engine’s rated sealevel horsepower from sea level up to the engine’s critical 

altitude. Critical altitude is the maximum altitude at which 

a turbocharged engine can produce its rated horsepower. 

Above the critical altitude, power output begins to decrease 

like it does for a normally aspirated engine.

Turbochargers increase the pressure of the engine’s induction 

air, which allows the engine to develop sea level or greater 

horsepower at higher altitudes. A turbocharger is comprised 

of two main elements: a compressor and turbine. The 

compressor section houses an impeller that turns at a high rate 

of speed. As induction air is drawn across the impeller blades, 

the impeller accelerates the air, allowing a large volume of 

air to be drawn into the compressor housing. The impeller’s 

action subsequently produces high-pressure, high-density 

air, which is delivered to the engine. To turn the impeller, 

the engine’s exhaust gases are used to drive a turbine wheel 

that is mounted on the opposite end of the impeller’s drive 

shaft. By directing different amounts of exhaust gases to flow 

over the turbine, more energy can be extracted, causing the 

impeller to deliver more compressed air to the engine. The 

waste gate, essentially an adjustable butterfly valve installed 

in the exhaust system, is used to vary the mass of exhaust gas 

flowing into the turbine. When closed, most of the exhaust 

gases from the engine are forced to flow through the turbine. 

When open, the exhaust gases are allowed to bypass the 

turbine by flowing directly out through the engine’s exhaust 

pipe. [Figure 6-15]

Since the temperature of a gas rises when it is compressed, 

turbocharging causes the temperature of the induction air to 

increase. To reduce this temperature and lower the risk of 

detonation, many turbocharged engines use an intercooler. 

This small heat exchanger uses outside air to cool the hot 

compressed air before it enters the fuel metering device.

System Operation

On most modern turbocharged engines, the position of 

the waste gate is governed by a pressure-sensing control 

mechanism coupled to an actuator. Engine oil directed into 

or away from this actuator moves the waste gate position. 

On these systems, the actuator is automatically positioned to 

produce the desired MAP simply by changing the position 

of the throttle control.
6-14

Other turbocharging system designs use a separate manual 

control to position the waste gate. With manual control, 

the manifold pressure gauge must be closely monitored to 

determine when the desired MAP has been achieved. Manual 

systems are often found on aircraft that have been modified 

with aftermarket turbocharging systems. These systems 

require special operating considerations. For example, if the 

waste gate is left closed after descending from a high altitude, 

it is possible to produce a manifold pressure that exceeds the 

engine’s limitations. This condition, called an overboost, 

may produce severe detonation because of the leaning effect 

resulting from increased air density during descent.

Although an automatic waste gate system is less likely 

to experience an overboost condition, it can still occur. If 

takeoff power is applied while the engine oil temperature is 

below its normal operating range, the cold oil may not flow 

out of the waste gate actuator quickly enough to prevent an 

overboost. To help prevent overboosting, advance the throttle 

cautiously to prevent exceeding the maximum manifold 

pressure limits.

A pilot flying an aircraft with a turbocharger should be aware 

of system limitations. For example, a turbocharger turbine 

and impeller can operate at rotational speeds in excess of 

80,000 rpm while at extremely high temperatures. To achieve 

high rotational speed, the bearings within the system must be 

constantly supplied with engine oil to reduce the frictional 

forces and high temperature. To obtain adequate lubrication, 

the oil temperature should be in the normal operating range 

before high throttle settings are applied. In addition, allow 

the turbocharger to cool and the turbine to slow down before 

shutting the engine down. Otherwise, the oil remaining in 

the bearing housing will boil, causing hard carbon deposits 

to form on the bearings and shaft. These deposits rapidly 

deteriorate the turbocharger’s efficiency and service life. For 

further limitations, refer to the AFM/POH.

High Altitude Performance

As an aircraft equipped with a turbocharging system climbs, 

the waste gate is gradually closed to maintain the maximum 

allowable manifold pressure. At some point, the waste gate 

will be fully closed and further increases in altitude will 

cause the manifold pressure to decrease. This is the critical 

altitude, which is established by the aircraft or engine 

manufacturer. When evaluating the performance of the 

turbocharging system, be aware that if the manifold pressure 

begins decreasing before the specified critical altitude, the 

engine and turbocharging system should be inspected by 

a qualified aviation maintenance technician to verify the 

system’s proper operation.

Ignition System

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 
6-15

Upper magneto wires

Lower magneto wires

4

�

�

�

Upper spark plugs Lower spark plugs

Left magneto Right magneto

Figure 6-16. Ignition system components.

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.

Oil Systems

The engine oil system performs several important functions:

• Lubrication of the engine’s moving parts

• Cooling of the engine by reducing friction

• Removing heat from the cylinders

• Providing a seal between the cylinder walls and 

pistons

• Carrying away contaminants

Reciprocating engines use either a wet-sump or a dry-sump 

oil system. In a wet-sump system, the oil is located in a sump, 

which is an integral part of the engine. In a dry-sump system, 

the oil is contained in a separate tank, and circulated through 

the engine by pumps. [Figure 6-17]

The main component of a wet-sump system is the oil pump, 

which draws oil from the sump and routes it to the engine. 

After the oil passes through the engine, it returns to the 

sump. In some engines, additional lubrication is supplied 

by the rotating crankshaft, which splashes oil onto portions 

of the engine.

An oil pump also supplies oil pressure in a dry-sump system, 

but the source of the oil is located external to the engine in a 

separate oil tank. After oil is routed through the engine, it is 

pumped from the various locations in the engine back to the 

oil tank by scavenge pumps. Dry-sump systems allow for a 

greater volume of oil to be supplied to the engine, which makes 

them more suitable for very large reciprocating engines. 

The oil pressure gauge provides a direct indication of the oil 

system operation. It ensures the pressure in pounds per square 

inch (psi) of the oil supplied to the engine. Green indicates 

the normal operating range, while red indicates the minimum 

and maximum pressures. There should be an indication of 

oil pressure during engine start. Refer to the AFM/POH for 

manufacturer limitations.
6-16

Figure 6-18. Always check the engine oil level during the preflight 

inspection.

Figure 6-17. Wet-sump oil system.

Engine

and 

Accessory 

Bearings

200

245

I50

I00

75

°F P

S

I

II5

60

I00

20

0 OIL

P

R

E

S

S

T

E

M

P

Oil cooler and filter

Oil pressure relief valve

Oil filter cap and dipstick

High pressure oil screen

Oil pump

Low pressure oil screen

Oil sump

Pressure oil from 

oil pump

Sump oil and return 

oil from relief valve

The oil temperature gauge measures the temperature of oil. 

A green area shows the normal operating range and the red 

line indicates the maximum allowable temperature. Unlike 

oil pressure, changes in oil temperature occur more slowly. 

This is particularly noticeable after starting a cold engine, 

when it may take several minutes or longer for the gauge to 

show any increase in oil temperature.

Check oil temperature periodically during flight especially 

when operating in high or low ambient air temperature. High 

oil temperature indications may signal a plugged oil line, a low 

oil quantity, a blocked oil cooler, or a defective temperature 

gauge. Low oil temperature indications may signal improper 

oil viscosity during cold weather operations.

The oil filler cap and dipstick (for measuring the oil quantity) 

are usually accessible through a panel in the engine cowling. If 

the quantity does not meet the manufacturer’s recommended 

operating levels, oil should be added. The AFM/POH or 

placards near the access panel provide information about 

the correct oil type and weight, as well as the minimum and 

maximum oil quantity. [Figure 6-18] 

Engine Cooling Systems

The burning fuel within the cylinders produces intense 

heat, most of which is expelled through the exhaust system. 

Much of the remaining heat, however, must be removed, or 

at least dissipated, to prevent the engine from overheating. 

Otherwise, the extremely high engine temperatures can lead 

to loss of power, excessive oil consumption, detonation, and 

serious engine damage.

While the oil system is vital to the internal cooling of the 

engine, an additional method of cooling is necessary for the 

engine’s external surface. Most small aircraft are air cooled, 

although some are liquid cooled. 

Air cooling is accomplished by air flowing into the engine 

compartment through openings in front of the engine 

cowling. Baffles route this air over fins attached to the engine 

cylinders, and other parts of the engine, where the air absorbs 

the engine heat. Expulsion of the hot air takes place through 
6-17

Figure 6-19. Outside air aids in cooling the engine.

Fixed cowl opening

Baffle

Baffle

Cylinders

Air inlet

one or more openings in the lower, aft portion of the engine 

cowling. [Figure 6-19]

The outside air enters the engine compartment through an 

inlet behind the propeller hub. Baffles direct it to the hottest 

parts of the engine, primarily the cylinders, which have fins 

that increase the area exposed to the airflow.

The air cooling system is less effective during ground 

operations, takeoffs, go-arounds, and other periods of highpower, low-airspeed operation. Conversely, high-speed 

descents provide excess air and can shock cool the engine, 

subjecting it to abrupt temperature fluctuations.

Operating the engine at higher than its designed temperature 

can cause loss of power, excessive oil consumption, and 

detonation. It will also lead to serious permanent damage, 

such as scoring the cylinder walls, damaging the pistons 

and rings, and burning and warping the valves. Monitoring 

the flight deck engine temperature instruments will aid in 

avoiding high operating temperature.

Under normal operating conditions in aircraft not equipped 

with cowl flaps, the engine temperature can be controlled 

by changing the airspeed or the power output of the engine. 

High engine temperatures can be decreased by increasing the 

airspeed and/or reducing the power.

The oil temperature gauge gives an indirect and delayed 

indication of rising engine temperature, but can be used for 

determining engine temperature if this is the only means 

available.

Most aircraft are equipped with a cylinder-head temperature 

gauge which indicates a direct and immediate cylinder 

temperature change. This instrument is calibrated in degrees 

Celsius or Fahrenheit, and is usually color coded with a 

green arc to indicate the normal operating range. A red line 

on the instrument indicates maximum allowable cylinder 

head temperature.

To avoid excessive cylinder head temperatures, increase 

airspeed, enrich the mixture, and/or reduce power. Any of 

these procedures help to reduce the engine temperature. On 

aircraft equipped with cowl flaps, use the cowl flap positions 

to control the temperature. Cowl flaps are hinged covers that 

fit over the opening through which the hot air is expelled. If 

the engine temperature is low, the cowl flaps can be closed, 

thereby restricting the flow of expelled hot air and increasing 

engine temperature. If the engine temperature is high, the cowl 

flaps can be opened to permit a greater flow of air through the 

system, thereby decreasing the engine temperature.

Exhaust Systems

Engine exhaust systems vent the burned combustion gases 

overboard, provide heat for the cabin, and defrost the 

windscreen. An exhaust system has exhaust piping attached 

to the cylinders, as well as a muffler and a muffler shroud. 

The exhaust gases are pushed out of the cylinder through 

the exhaust valve and then through the exhaust pipe system 

to the atmosphere.

For cabin heat, outside air is drawn into the air inlet and is 

ducted through a shroud around the muffler. The muffler is 

heated by the exiting exhaust gases and, in turn, heats the 

air around the muffler. This heated air is then ducted to the 

cabin for heat and defrost applications. The heat and defrost 

are controlled in the flight deck, and can be adjusted to the 

desired level. 

Exhaust gases contain large amounts of carbon monoxide, 

which is odorless and colorless. Carbon monoxide is deadly, 

and its presence is virtually impossible to detect. The exhaust 

system must be in good condition and free of cracks.

Some exhaust systems have an EGT probe. This probe 

transmits the EGT to an instrument in the flight deck. The 

EGT gauge measures the temperature of the gases at the 

exhaust manifold. This temperature varies with the ratio 

of fuel to air entering the cylinders and can be used as a 
6-18

Figure 6-21. Normal combustion and explosive combustion.

Combustion

Normal combustion Explosion

Figure 6-20. Typical starting circuit.

Electrical system schematic

M

A

I

N

B

U

S

Starter

Battery

A

L

T

B

A

T

Battery

contactor

(solenoid)

Starter

contactor

External

power

relay

+

+

OFF

L

R B

S

Ignition switch

External

power plug

basis for regulating the fuel/air mixture. The EGT gauge 

is highly accurate in indicating the correct mixture setting. 

When using the EGT to aid in leaning the fuel/air mixture, 

fuel consumption can be reduced. For specific procedures, 

refer to the manufacturer’s recommendations for leaning 

the mixture.

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 onboard 

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 bar 

through the battery solenoid. Both the starter and the starter 

switch draw current from the main bus bar, 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 6-20]

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 aircraft should be in an area where 

the propeller will not stir up gravel or dust.

Combustion

During normal combustion, the fuel/air mixture burns in a 

very controlled and predictable manner. In a spark ignition 

engine the process occurs in a fraction of a second. The 

mixture actually begins to burn at the point where it is ignited 

by the spark plugs, then burns away from the plugs until it 

is all consumed. This type of combustion causes a smooth 

build-up of temperature and pressure and ensures that the 

expanding gases deliver the maximum force to the piston at 

exactly the right time in the power stroke. [Figure 6-21]
6-19

Detonation is an uncontrolled, explosive ignition of the 

fuel/air mixture within the cylinder’s combustion chamber. 

It causes excessive temperatures and pressures which, if not 

corrected, can quickly lead to failure of the piston, cylinder, 

or valves. In less severe cases, detonation causes engine 

overheating, roughness, or loss of power.

Detonation is characterized by high cylinder head temperatures 

and is most likely to occur when operating at high power 

settings. Common operational causes of detonation are:

• Use of a lower fuel grade than that specified by the 

aircraft manufacturer.

• Operation of the engine with extremely high manifold 

pressures in conjunction with low rpm.

• Operation of the engine at high power settings with 

an excessively lean mixture.

• Maintaining extended ground operations or steep 

climbs in which cylinder cooling is reduced.

Detonation may be avoided by following these basic 

guidelines during the various phases of ground and flight 

operations:

• Make sure the proper grade of fuel is used.

• Keep the cowl flaps (if available) in the full-open 

position while on the ground to provide the maximum 

airflow through the cowling.

• Use an enriched fuel mixture, as well as a shallower 

climb angle to increase cylinder cooling during takeoff 

and initial climb.

• Avoid extended, high power, steep climbs.

• Develop the habit of monitoring the engine instruments 

to verify proper operation according to procedures 

established by the manufacturer.

Preignition occurs when the fuel/air mixture ignites prior 

to the engine’s normal ignition event. Premature burning 

is usually caused by a residual hot spot in the combustion 

chamber, often created by a small carbon deposit on a spark 

plug, a cracked spark plug insulator, or other damage in the 

cylinder that causes a part to heat sufficiently to ignite the 

fuel/air charge. Preignition causes the engine to lose power, 

and produces high operating temperature. As with detonation, 

preignition may also cause severe engine damage, because 

the expanding gases exert excessive pressure on the piston 

while still on its compression stroke.

Detonation and preignition often occur simultaneously and 

one may cause the other. Since either condition causes high 

engine temperature accompanied by a decrease in engine 

performance, it is often difficult to distinguish between the 

two. Using the recommended grade of fuel and operating 

the engine within its proper temperature, pressure, and rpm 

ranges reduce the chance of detonation or preignition.

Full Authority Digital Engine Control 

(FADEC)

FADEC is a system consisting of a digital computer and 

ancillary components that control an aircraft’s engine 

and propeller. First used in turbine-powered aircraft, and 

referred to as full authority digital electronic control, these 

sophisticated control systems are increasingly being used in 

piston powered aircraft.

In a spark ignition reciprocating engine the FADEC uses 

speed, temperature, and pressure sensors to monitor the 

status of each cylinder. A digital computer calculates the 

ideal pulse for each injector and adjusts ignition timing as 

necessary to achieve optimal performance. In a compression 

ignition engine the FADEC operates similarly and performs 

all of the same functions, excluding those specifically related 

to the spark ignition process. 

FADEC systems eliminate the need for magnetos, carburetor 

heat, mixture controls, and engine priming. A single throttle 

lever is characteristic of an aircraft equipped with a FADEC 

system. The pilot simply positions the throttle lever to a 

desired detent such as start, idle, cruise power, or max power, 

and the FADEC system adjusts the engine and propeller 

automatically for the mode selected. There is no need for the 

pilot to monitor or control the air/fuel mixture.

During aircraft starting, the FADEC primes the cylinders, 

adjusts the mixture, and positions the throttle based on engine 

temperature and ambient pressure. During cruise flight, the 

FADEC constantly monitors the engine and adjusts fuel flow, 

and ignition timing individually in each cylinder. This precise 

control of the combustion process often results in decreased 

fuel consumption and increased horsepower. 

FADEC systems are considered an essential part of the 

engine and propeller control, and may be powered by the 

aircraft’s main electrical system. In many aircraft FADEC 

uses power from a separate generator connected to the engine. 

In either case, there must be a backup electrical source 

available because failure of a FADEC system could result 

in a complete loss of engine thrust. To prevent loss of thrust, 

two separate and identical digital channels are incorporated 

for redundancy, each channel capable of providing all engine 

and propeller functions without limitations. 
6-20

Figure 6-22. Eclipse 500 VLJ.

Turbine Engines

An aircraft turbine engine consists of an air inlet, compressor, 

combustion chambers, a turbine section, and exhaust. Thrust 

is produced by increasing the velocity of the air flowing 

through the engine. Turbine engines are highly desirable 

aircraft powerplants. They are characterized by smooth 

operation and a high power-to-weight ratio, and they 

use readily available jet fuel. Prior to recent advances in 

material, engine design, and manufacturing processes, the 

use of turbine engines in small/light production aircraft was 

cost prohibitive. Today, several aviation manufacturers are 

producing or plan to produce small/light turbine-powered 

aircraft. These smaller turbine-powered aircraft typically 

seat between three and seven passengers and are referred to 

as very light jets (VLJs) or microjets. [Figure 6-22] 

Types of Turbine Engines

Turbine engines are classified according to the type of 

compressors they use. There are three types of compressors—

centrifugal flow, axial flow, and centrifugal-axial flow. 

Compression of inlet air is achieved in a centrifugal flow 

engine by accelerating air outward perpendicular to the 

longitudinal axis of the machine. The axial-flow engine 

compresses air by a series of rotating and stationary 

airfoils moving the air parallel to the longitudinal axis. The 

centrifugal-axial flow design uses both kinds of compressors 

to achieve the desired compression.

The path the air takes through the engine and how power is 

produced determines the type of engine. There are four types 

of aircraft turbine engines—turbojet, turboprop, turbofan, 

and turboshaft.

Turbojet

The turbojet engine consists of four sections: compressor, 

combustion chamber, turbine section, and exhaust. The 

compressor section passes inlet air at a high rate of speed to 

the combustion chamber. The combustion chamber contains 

the fuel inlet and igniter for combustion. The expanding 

air drives a turbine, which is connected by a shaft to the 

compressor, sustaining engine operation. The accelerated 

exhaust gases from the engine provide thrust. This is a basic 

application of compressing air, igniting the fuel-air mixture, 

producing power to self-sustain the engine operation, and 

exhaust for propulsion. [Figure 6-23] 

Turbojet engines are limited in range and endurance. They 

are also slow to respond to throttle applications at slow 

compressor speeds.

Turboprop

A turboprop engine is a turbine engine that drives a propeller 

through a reduction gear. The exhaust gases drive a power 

turbine connected by a shaft that drives the reduction gear 

assembly. Reduction gearing is necessary in turboprop engines 

because optimum propeller performance is achieved at much 

slower speeds than the engine’s operating rpm. Turboprop 

engines are a compromise between turbojet engines and 

reciprocating powerplants. Turboprop engines are most 

efficient at speeds between 250 and 400 mph and altitudes 

between 18,000 and 30,000 feet. They also perform well at 

the slow airspeeds required for takeoff and landing, and are 

fuel efficient. The minimum specific fuel consumption of the 

turboprop engine is normally available in the altitude range 

of 25,000 feet to the tropopause. [Figure 6-24] 

Turbofan

Turbofans were developed to combine some of the best 

features of the turbojet and the turboprop. Turbofan engines 

are designed to create additional thrust by diverting a 

secondary airflow around the combustion chamber. The 

turbofan bypass air generates increased thrust, cools the 

engine, and aids in exhaust noise suppression. This provides 

turbojet-type cruise speed and lower fuel consumption.

The inlet air that passes through a turbofan engine is usually 

divided into two separate streams of air. One stream passes 

through the engine core, while a second stream bypasses the 

engine core. It is this bypass stream of air that is responsible 

for the term “bypass engine.” A turbofan’s bypass ratio refers 

to the ratio of the mass airflow that passes through the fan 

divided by the mass airflow that passes through the engine 

core. [Figure 6-25] 

Turboshaft

The fourth common type of jet engine is the turboshaft. 

[Figure 6-26] It delivers power to a shaft that drives 

something other than a propeller. The biggest difference 

between a turbojet and turboshaft engine is that on a 
6-21

Figure 6-23. Turbojet engine.

Turbojet engine components

Fuel injector

Inlet

Compressor Combustion chamber

Turbine

Hot gases

Nozzle

Figure 6-24. Turboprop engine.

Turboprop engine components

Gear box Compressor Combustion chamber

Turbine

Exhaust

Fuel injector

Inlet

Prop

Figure 6-25. Turbofan engine.

Turbofan engine components

Inlet Fuel injector

Compressor Combustion chamber

Turbine

Hot gases

Nozzle

Duct fan

Secondary air stream

Primary air stream

Figure 6-26. Turboshaft engine.

Turboprop engine components

Inlet Compressor Combustion chamber Exhaust

Free (power) turbine

Compressor turbine

Power shaft
6-22

Figure 6-27. Dual-spool axial-flow compressor.

Dual-spool axial-flow compressor

Low pressure compressor drive shaft

Low pressure

compressor (N1)

High pressure

compressor (N2)

High pressure compressor drive shaft

turboshaft engine, most of the energy produced by the 

expanding gases is used to drive a turbine rather than produce 

thrust. Many helicopters use a turboshaft gas turbine engine. 

In addition, turboshaft engines are widely used as auxiliary 

power units on large aircraft.

Turbine Engine Instruments

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 thermocouples, which provide pilots with 

temperature readings in and around the turbine section. 

Engine Pressure Ratio (EPR) 

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 

flight deck EPR gauge.

EPR system design automatically compensates for the effects 

of airspeed and altitude. Changes in ambient temperature 

require a correction be applied to EPR indications to provide 

accurate engine power settings.

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, interstage turbine temperature (ITT) gauge, and 

turbine gas temperature (TGT) gauge.

Torquemeter

Turboprop/turboshaft engine power output is measured by the 

torquemeter. Torque is a twisting force applied to a shaft. The 

torquemeter measures power applied to the shaft. Turboprop 

and turboshaft engines are designed to produce torque for 

driving a propeller. Torquemeters are calibrated in percentage 

units, foot-pounds, or psi.

N1 Indicator

N1 represents the rotational speed of the low pressure 

compressor and is presented on the indicator as a percentage 

of design rpm. 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 Indicator

N2 represents the rotational speed of the high pressure 

compressor and is presented on the indicator as a percentage 

of design rpm. 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. 

[Figure 6-27] 

Turbine Engine Operational Considerations

The great variety of turbine engines makes it impractical to 

cover specific operational procedures, but there are certain 

operational considerations common to all turbine engines. 

They are engine temperature limits, foreign object damage, 

hot start, compressor stall, and flameout.

Engine Temperature Limitations

The highest temperature in any turbine engine occurs at the 

turbine inlet. Turbine inlet temperature is therefore usually 

the limiting factor in turbine engine operation.

Thrust Variations

Turbine engine thrust varies directly with air density. As air 

density decreases, so does thrust. Additionally, because air 
6-23

Figure 6-28. Comparison of normal and distorted airflow into the 

compressor section.

Comparison of normal and distorted airflow

Normal inlet airflow

Distorted inlet airflow

density decreases with an increase in temperature, increased 

temperatures will also result in decreased thrust. While both 

turbine and reciprocating powered engines are affected to 

some degree by high relative humidity, turbine engines will 

experience a negligible loss of thrust, while reciprocating 

engines a significant loss of brake horsepower.

Foreign Object Damage (FOD)

Due to the design and function of a turbine engine’s air inlet, 

the possibility of ingestion of debris always exists. This 

causes significant damage, particularly to the compressor 

and turbine sections. When ingestion of debris occurs, it is 

called foreign object damage (FOD). Typical FOD consists 

of small nicks and dents caused by ingestion of small objects 

from the ramp, taxiway, or runway, but FOD damage caused 

by bird strikes or ice ingestion also occur. Sometimes FOD 

results in total destruction of an engine.

Prevention of FOD is a high priority. Some engine inlets 

have a tendency to form a vortex between the ground and 

the inlet during ground operations. A vortex dissipater may 

be installed on these engines. Other devices, such as screens 

and/or deflectors, may also be utilized. Preflight procedures 

include a visual inspection for any sign of FOD.

Turbine Engine Hot/Hung Start

When the EGT exceeds the safe limit of an aircraft, it 

experiences a “hot start.” It is caused by too much fuel 

entering the combustion chamber, or insufficient turbine 

rpm. Any time an engine has a hot start, refer to the AFM/

POH or an appropriate maintenance manual for inspection 

requirements.

If the engine fails to accelerate to the proper speed after 

ignition or does not accelerate to idle rpm, a hung or false start 

has occurred. A hung start may be caused by an insufficient 

starting power source or fuel control malfunction.

Compressor Stalls

Compressor blades are small airfoils and are subject to the 

same aerodynamic principles that apply to any airfoil. A 

compressor blade has an angle of attack which is a result of 

inlet air velocity and the compressor’s rotational velocity. 

These two forces combine to form a vector, which defines the 

airfoil’s actual angle of attack to the approaching inlet air. 

A compressor stall is an imbalance between the two vector 

quantities, inlet velocity and compressor rotational speed. 

Compressor stalls occur when the compressor blades’ angle 

of attack exceeds the critical angle of attack. At this point, 

smooth airflow is interrupted and turbulence is created with 

pressure fluctuations. Compressor stalls cause air flowing 

in the compressor to slow down and stagnate, sometimes 

reversing direction. [Figure 6-28] 

Compressor stalls can be transient and intermittent or steady 

and severe. Indications of a transient/intermittent stall are 

usually an intermittent “bang” as backfire and flow reversal 

take place. If the stall develops and becomes steady, strong 

vibration and a loud roar may develop from the continuous 

flow reversal. Often, the flight deck gauges do not show a 

mild or transient stall, but they do indicate a developed stall. 

Typical instrument indications include fluctuations in rpm and 

an increase in exhaust gas temperature. Most transient stalls 

are not harmful to the engine and often correct themselves 

after one or two pulsations. The possibility of severe engine 

damage from a steady state stall is immediate. Recovery must 

be accomplished by quickly reducing power, decreasing the 

aircraft’s angle of attack, and increasing airspeed.

Although all gas turbine engines are subject to compressor 

stalls, most models have systems that inhibit them. One 

system uses a variable inlet guide vane (VIGV) and variable 

stator vanes, which direct the incoming air into the rotor 

blades at an appropriate angle. To prevent air pressure stalls, 

operate the aircraft within the parameters established by the 

manufacturer. If a compressor stall does develop, follow the 

procedures recommended in the AFM/POH.

Flameout

A flameout occurs in the operation of a gas turbine engine in 

which the fire in the engine unintentionally goes out. If the 

rich limit of the fuel/air ratio is exceeded in the combustion 

chamber, the flame will blow out. This condition is often 

referred to as a rich flameout. It generally results from very 

fast engine acceleration, in which an overly rich mixture 

causes the fuel temperature to drop below the combustion 
6-24

Figure 6-29. Engine net thrust versus aircraft speed and drag. Points 

A through F are explained in the text below. Net thrust

 Airspeed

Aircraft drag

A B C D E F

Reciprocating

Turboprop

Turbofan

Turbojet

temperature. It may also be caused by insufficient airflow 

to support combustion.

A more common flameout occurrence is due to low fuel 

pressure and low engine speeds, which typically are 

associated with high-altitude flight. This situation may also 

occur with the engine throttled back during a descent, which 

can set up the lean-condition flameout. A weak mixture can 

easily cause the flame to die out, even with a normal airflow 

through the engine. 

Any interruption of the fuel supply can result in a 

flameout. This may be due to prolonged unusual attitudes, 

a malfunctioning fuel control system, turbulence, icing or 

running out of fuel.

Symptoms of a flameout normally are the same as those 

following an engine failure. If the flameout is due to a 

transitory condition, such as an imbalance between fuel 

flow and engine speed, an airstart may be attempted once 

the condition is corrected. In any case, pilots must follow 

the applicable emergency procedures outlined in the AFM/

POH. Generally these procedures contain recommendations 

concerning altitude and airspeed where the airstart is most 

likely to be successful.

Performance Comparison

It is possible to compare the performance of a reciprocating 

powerplant and different types of turbine engines. For 

the comparison to be accurate, thrust horsepower (usable 

horsepower) for the reciprocating powerplant must be used 

rather than brake horsepower, and net thrust must be used 

for the turbine-powered engines. In addition, aircraft design 

configuration and size must be approximately the same. 

When comparing performance, the following definitions 

are useful:

Brake horsepower (BHP)—the horsepower actually delivered 

to the output shaft. Brake horsepower is the actual usable 

horsepower. 

Net thrust—the thrust produced by a turbojet or turbofan 

engine.

Thrust horsepower (THP)—the horsepower equivalent of the 

thrust produced by a turbojet or turbofan engine. 

Equivalent shaft horsepower (ESHP)—with respect to 

turboprop engines, the sum of the shaft horsepower (SHP) 

delivered to the propeller and THP produced by the exhaust 

gases.

Figure 6-29 shows how four types of engines compare in net 

thrust as airspeed is increased. This figure is for explanatory 

purposes only and is not for specific models of engines. The 

following are the four types of engines:

• Reciprocating powerplant

• Turbine, propeller combination (turboprop)

• Turbine engine incorporating a fan (turbofan)

• Turbojet (pure jet)

By plotting the performance curve for each engine, a 

comparison can be made of maximum aircraft speed variation 

with the type of engine used. Since the graph is only a means 

of comparison, numerical values for net thrust, aircraft speed, 

and drag are not included.

Comparison of the four powerplants on the basis of net thrust 

makes certain performance capabilities evident. In the speed 

range shown to the left of line A, the reciprocating powerplant 

outperforms the other three types. The turboprop outperforms 

the turbofan in the range to the left of line C. The turbofan 

engine outperforms the turbojet in the range to the left of 

line F. The turbofan engine outperforms the reciprocating 

powerplant to the right of line B and the turboprop to the 

right of line C. The turbojet outperforms the reciprocating 

powerplant to the right of line D, the turboprop to the right 

of line E, and the turbofan to the right of line F.
6-25

Figure 6-30. Gravity-feed and fuel-pump systems.

Fuel systems

Gravity-feed system

Fuel-pump system

Engine-driven pump

Right tank

Right tank

 Left tank

 Left tank

BOTH

LEFT

RIGHT

OFF

BOTH

LEFT

RIGHT

OFF

Selector valve

Carburetor

Primer

Strainer

Vent

Selector valve

Carburetor

Electric pump

Strainer

Primer

The points where the aircraft drag curve intersects the net 

thrust curves are the maximum aircraft speeds. The vertical 

lines from each of the points to the baseline of the graph 

indicate that the turbojet aircraft can attain a higher maximum 

speed than aircraft equipped with the other types of engines. 

Aircraft equipped with the turbofan engine will attain a higher 

maximum speed than aircraft equipped with a turboprop or 

reciprocating powerplant.

Airframe Systems 

Fuel, electrical, hydraulic, and oxygen systems make up the 

airframe systems.

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 aircraft: gravity-feed and fuel-pump systems.

Gravity-Feed System 

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, 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 aircraft is such that gravity cannot be used 

to transfer fuel, fuel pumps are installed. For example, on 

low-wing airplanes, the fuel tanks in the wings are located 

below the carburetor. [Figure 6-30] 

Fuel-Pump System

Aircraft with fuel-pump systems have two fuel pumps. The 

main pump system is engine driven with an electrically driven 

auxiliary pump 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 flight deck.

Fuel Primer

Both gravity-feed and fuel-pump systems may incorporate a 

fuel primer into the system. The fuel primer is used to draw 

fuel from the tanks to vaporize fuel directly into the cylinders 

prior to starting the engine. During cold weather, when engines 

are difficult to start, the fuel primer helps 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 the aircraft.

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.
6-26

Figure 6-31. Fuel selector valve.

Fuel selector valve

 LEFT

 19 gal

LEVEL

 FLIGHT

 ONLY

RIGHT 

19 gal 

LEVEL

FLIGHT

ONLY 

OFF

TAKEOFF

LANDING

ALL FLIGHT 38 GAL ATTITUDES

BOTH

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.

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 6-31]

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 and cause vapor lock, which makes it difficult to 

restart the engine. On fuel-injected engines, the fuel becomes 

so hot it vaporizes in the fuel line, not allowing fuel to reach 

the cylinders.

Fuel Strainers, Sumps, and Drains

After leaving the fuel tank and before it enters the carburetor, 

the fuel passes through a strainer which removes any moisture 

and other sediments 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 a low point in a fuel 

system and/or fuel tank. The fuel system may contain sump, 

fuel strainer, and fuel tank drains, 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, more water in the fuel tanks is probable 

and they should be drained until there is no evidence of water. 

Never take off until all water and contaminants have been 

removed from the engine fuel system. 

Because of the variation in fuel systems, become thoroughly 

familiar with the systems that apply to the aircraft being flown. 

Consult the AFM/POH for specific operating procedures.

Fuel Grades

Aviation gasoline (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 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 grade lower than recommended. This can cause 

the cylinder head temperature and engine oil temperature to 

exceed their normal operating ranges, which may result in 

detonation.

Several grades of AVGAS 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/POH, on placards in the flight deck, and 

next to the filler caps. 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 (FAA).
6-27

Figure 6-32. Aviation fuel color-coding system.

Aviation fuel color-coding system

80

AVGAS

100

AVGAS

100LL

AVGAS

JET

A

AVGAS

100LL

AVGAS

80

AVGAS

100 JET A

RED GREEN BLUE COLORLESS

OR STRAW

The current method identifies AVGAS for aircraft with 

reciprocating engines 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 6-32] 

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 AVGAS is 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

Accidents attributed to powerplant failure from fuel 

contamination have often 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 has 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 after the last 

flight of the day to prevent moisture condensation within the 

tank. To prevent fuel contamination, avoid refueling from 

cans and drums. 

In remote areas or in emergency situations, there may be no 

alternative to refueling from sources with inadequate anticontamination systems. While a chamois skin and funnel 

may be the only possible means of filtering fuel, using 

them is hazardous. Remember, 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 aircraft 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. Because both the aircraft and 

refueler have different static charges, bonding both components 

to each other is critical. By bonding both components to each 

other, the static differential charge is equalized. The refueling 

nozzle should be bonded to the aircraft before refueling begins 

and should remain bonded 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. Bond drum to aircraft or nozzle to aircraft before the 

fuel cap is removed

When disconnecting, reverse the order.
6-28

Figure 6-33. On this master switch, the left half is for the alternator 

and the right half is for the battery.

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.

Electrical System

Most aircraft are equipped with either a 14- or a 28-volt direct 

current electrical system. A basic aircraft electrical system 

consists of the following components:

• Alternator/generator

• Battery

• Master/battery switch

• Alternator/generator switch

• Bus bar, fuses, and circuit breakers

• Voltage regulator

• Ammeter/loadmeter

• Associated electrical wiring

Engine-driven alternators or generators supply electric 

current to the electrical system. They also maintain a 

sufficient electrical charge in the battery. Electrical energy 

stored in a battery provides a source of electrical power for 

starting the engine and a limited supply of electrical power 

for use in the event the alternator or generator fails.

Most direct-current generators will not produce a sufficient 

amount of electrical current at low engine rpm to operate the 

entire electrical system. During operations at low engine rpm, 

the electrical needs must be drawn from the battery, which 

can quickly be depleted.

Alternators have several advantages over generators. 

Alternators produce sufficient current to operate the entire 

electrical system, even at slower engine speeds, by producing 

alternating current, which is converted to direct current. The 

electrical output of an alternator is more constant throughout 

a wide range of engine speeds.

Some aircraft have receptacles to which an external ground 

power unit (GPU) may be connected to provide electrical 

energy for starting. These are very useful, especially 

during cold weather starting. Follow the manufacturer’s 

recommendations for engine starting using a GPU.

The electrical system is turned on or off with a master switch. 

Turning the master switch to the ON position provides 

electrical energy to all the electrical equipment circuits 

except the ignition system. Equipment that commonly uses 

the electrical system for its source of energy includes:

• Position lights

• Anticollision lights

• Landing lights

• Taxi lights

• Interior cabin lights

• Instrument lights

• Radio equipment

• Turn indicator

• Fuel gauges

• Electric fuel pump

• Stall warning system

• Pitot heat

• Starting motor

Many aircraft are equipped with a battery switch that controls 

the electrical power to the aircraft in a manner similar 

to the master switch. In addition, an alternator switch is 

installed which permits the pilot to exclude the alternator 

from the electrical system in the event of alternator failure. 

[Figure 6-33] 

With the alternator half of the switch in the OFF position, the 

entire electrical load is placed on the battery. All nonessential 

electrical equipment should be turned off to conserve battery 

power.
6-29

Figure 6-34. Electrical system schematic.

Electrical system schematic

Low-voltage warning light

ALT

To wing flap circuit

breaker

Magnetos

Starter

contactor

Clock

Ammeter

Alternator

control unit

Circuit breaker (auto-reset)

Circuit breaker (push to reset)

Circuit breaker (pull—off,

push to reset)

Fuse Diode 

 Resistor

Capacitor (Noise Filter)

To inst 

LTS circuit 

breaker FUEL IND.

BCN PITOT

PULL

OFF STROBE 

RADIO FAN

LDG LTS

FLAP

INST LTS

STBY VAC

RADIO 1

RADIO 2

RADIO 3

NAV

DOME

RADIO 4

G

F

B

Alternator

B

Master

switch

Pull off

Alternator field

circuit breaker

Flight hour

recorder

Oil pressure

switch

To fuel quantity indicators

To flashing beacon

To pitot heat

To radio cooling fan

To strobe lights

To landing and taxi lights

To ignition switch

To wing flap system

To red doorpost maplight

To low-voltage warning light

To instrument, radio, compass 

and post lights

To oil temperature gauge

To turn coordinator

To low-vacuum warning light

Switch/circuit breaker to

standby vacuum pump

To white doorpost light

To audio muting relay

To control wheel maplight

To navigation lights

To dome light

To radio

To radio

To radio or transponder

and encoding altimeter

To radio

P

R

I

M

A

R

Y

B

U

S

A

V

I

O

N

I

C

S

B

U

S

Low volt out

Power in

 Sense (+)

 Field

 Sense (-)

Ground

Starter

Battery

A

L

T

B

A

T

+ 60 - 60

-30 +30 0

A M P

L R

Battery

contactor

Ground service

plug receptacle

CODE

A bus bar is used as a terminal in the aircraft electrical system 

to connect the main electrical system to the equipment using 

electricity as a source of power. This simplifies the wiring 

system and provides a common point from which voltage can 

be distributed throughout the system. [Figure 6-34] 

Fuses or circuit breakers are used in the electrical system to 

protect the circuits and equipment from electrical overload. 

Spare fuses of the proper amperage limit should be carried 

in the aircraft to replace defective or blown fuses. Circuit 

breakers have the same function as a fuse but can be manually 

reset, rather than replaced, if an overload condition occurs in 

the electrical system. Placards at the fuse or circuit breaker 

panel identify the circuit by name and show the amperage 

limit.
6-30

Figure 6-36. Basic hydraulic system.

Basic hydraulic system

Pump

Selector valve

System relief valve

Double acting cylinder Motion

Hydraulic fluid supply

Return fluid

Hydraulic pressure

Figure 6-35. Ammeter and loadmeter.

+ 60 - 60

-30 +30 0

A M P

ALT AMPS

0 30 60

Ammeter

Loadmeter

An ammeter is used to monitor the performance of the aircraft 

electrical system. The ammeter shows if the alternator/

generator is producing an adequate supply of electrical power. 

It also indicates whether or not the battery is receiving an 

electrical charge.

Ammeters are designed with the zero point in the center 

of the face and a negative or positive indication on either 

side. [Figure 6-35] When the pointer of the ammeter is 

on the plus side, it shows the charging rate of the battery. 

A minus indication means more current is being drawn 

from the battery than is being replaced. A full-scale minus 

deflection indicates a malfunction of the alternator/generator. 

A full-scale positive deflection indicates a malfunction of 

the regulator. In either case, consult the AFM or POH for 

appropriate action to be taken.

Not all aircraft are equipped with an ammeter. Some have 

a warning light that, when lighted, indicates a discharge in 

the system as a generator/alternator malfunction. Refer to the 

AFM or POH for appropriate action to be taken.

Another electrical monitoring indicator is a loadmeter. 

This type of gauge has a scale beginning with zero and 

shows the load being placed on the alternator/generator. 

[Figure 6-35] The loadmeter reflects the total percentage of 

the load placed on the generating capacity of the electrical 

system by the electrical accessories and battery. When all 

electrical components are turned off, it reflects only the 

amount of charging current demanded by the battery.

A voltage regulator controls the rate of charge to the battery 

by stabilizing the generator or alternator electrical output. The 

generator/alternator voltage output should be higher than the 

battery voltage. For example, a 12-volt battery would be fed 

by a generator/alternator system of approximately 14 volts. 

The difference in voltage keeps the battery charged.

Hydraulic Systems

There are multiple applications for hydraulic use in aircraft, 

depending on the complexity of the aircraft. 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. [Figure 6-36]

The hydraulic fluid is pumped through the system to an 

actuator or 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. Servos can be 

either single-acting or double-acting, 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. A 

single-acting servo provides power in one direction. The 

selector valve allows the fluid direction to be controlled. 

This is necessary for operations such as the extension and 

retraction of landing gear during which 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 hydraulic fluid is the most widely used type 

for small aircraft. This type of hydraulic fluid, a kerosene-like 

petroleum product, has good lubricating properties, as well 

as additives to inhibit foaming and prevent the formation of 

corrosion. It is chemically stable, has very little viscosity 

change with temperature, and is dyed for identification. 

Since several types of hydraulic fluids are commonly used, 
6-31

Figure 6-37. The landing gear supports the airplane during the 

takeoff run, landing, taxiing, and when parked.

Figure 6-38. Tailwheel landing gear.

an aircraft must be serviced with the type specified by the 

manufacturer. Refer to the AFM/POH or the Maintenance 

Manual.

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] 
6-32

Figure 6-40. High performance airplane pressurization system.

High performance airplane pressurization system CODE

Ambient air

Compressor discharge air

Pressurization air

Pre-heated ambient air

Conditioned pressurization air

Pressurized cabin

Air scoops

Turbocharger

compressor

section

Heat shroud

Heat exchanger

Forward 

air outlets

Cabin heat

valve

Floor level outlets

Safety/dump valve

Flow control

venturi

To cabin altitude

controller

Outflow valve

Figure 6-39. Fixed (left) and retractable (right) gear airplanes.

Brakes

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.

Pressurized Aircraft

Aircraft are flown at high altitudes for two reasons. First, an 

aircraft flown at high altitude consumes less fuel for a given 

airspeed than it does for the same speed at a lower altitude 

because the aircraft is more efficient at a high altitude. 

Second, bad weather and turbulence may be avoided by flying 

in relatively smooth air above the storms. Many modern 

aircraft are being designed to operate at high altitudes, taking 

advantage of that environment. In order to fly at higher 

altitudes, the aircraft must be pressurized. It is important 

for pilots who fly these aircraft to be familiar with the basic 

operating principles.

In a typical pressurization system, the cabin, flight 

compartment, and baggage compartments are incorporated 

into a sealed unit capable of containing air under a pressure 

higher than outside atmospheric pressure. On aircraft powered 

by turbine engines, bleed air from the engine compressor 

section is used to pressurize the cabin. Superchargers may 

be used on older model turbine-powered aircraft to pump 

air into the sealed fuselage. Piston-powered aircraft may use 

air supplied from each engine turbocharger through a sonic 

venturi (flow limiter). Air is released from the fuselage by 

a device called an outflow valve. By regulating the air exit, 

the outflow valve allows for a constant inflow of air to the 

pressurized area. [Figure 6-40] 

A cabin pressurization system typically maintains a cabin 

pressure altitude of approximately 8,000 feet at the maximum 

designed cruising altitude of an aircraft. This prevents rapid 

changes of cabin altitude that may be uncomfortable or cause 

injury to passengers and crew. In addition, the pressurization 

system permits a reasonably fast exchange of air from 

the inside to the outside of the cabin. This is necessary to 

eliminate odors and to remove stale air. [Figure 6-41] 

Pressurization of the aircraft cabin is an accepted method of 

protecting occupants against the effects of hypoxia. Within a 

pressurized cabin, occupants can be transported comfortably 
6-33

Figure 6-41. Standard atmospheric pressure chart.

Altitude (ft) Pressure (psi)

Atmosphere pressure

Sea level

2,000

4,000

6,000

8,000

10,000

12,000

14,000

16,000

18,000

20,000

22,000

24,000

26,000

28,000

30,000

14.7

13.7

12.7

11.8

10.9

10.1

9.4

8.6

8.0

7.3

6.8

6.2

5.7

5.2

4.8

4.4

At an altitude of 28,000 

feet, standard atmospheric pressure is 4.8 

psi. By adding this

pressure to the cabin

pressure differential of

6.1 psi difference (psid),

a total air pressure of

10.9 psi is obtained.

The altitude at which the 

standard air pressure is

equal to 10.9 psi can be 

found at 8,000 feet.

and safely for long periods of time, particularly if the cabin 

altitude is maintained at 8,000 feet or below, where the use 

of oxygen equipment is not required. The flight crew in this 

type of aircraft must be aware of the danger of accidental 

loss of cabin pressure and be prepared to deal with such an 

emergency whenever it occurs.

The following terms will aid in understanding the operating 

principles of pressurization and air conditioning systems:

• Aircraft altitude—the actual height above sea level at 

which the aircraft is flying

• Ambient temperature—the temperature in the area 

immediately surrounding the aircraft

• Ambient pressure—the pressure in the area immediately 

surrounding the aircraft

• Cabin altitude—cabin pressure in terms of equivalent 

altitude above sea level

• Differential pressure—the difference in pressure 

between the pressure acting on one side of a wall and 

the pressure acting on the other side of the wall. In 

aircraft air-conditioning and pressurizing systems, it is 

the difference between cabin pressure and atmospheric 

pressure.

The cabin pressure control system provides cabin pressure 

regulation, pressure relief, vacuum relief, and the means 

for selecting the desired cabin altitude in the isobaric and 

differential range. In addition, dumping of the cabin pressure 

is a function of the pressure control system. A cabin pressure 

regulator, an outflow valve, and a safety valve are used to 

accomplish these functions.

The cabin pressure regulator controls cabin pressure to a 

selected value in the isobaric range and limits cabin pressure 

to a preset differential value in the differential range. When an 

aircraft reaches the altitude at which the difference between 

the pressure inside and outside the cabin is equal to the 

highest differential pressure for which the fuselage structure 

is designed, a further increase in aircraft altitude will result in 

a corresponding increase in cabin altitude. Differential control 

is used to prevent the maximum differential pressure, for 

which the fuselage was designed, from being exceeded. This 

differential pressure is determined by the structural strength of 

the cabin and often by the relationship of the cabin size to the 

probable areas of rupture, such as window areas and doors.

The cabin air pressure safety valve is a combination 

pressure relief, vacuum relief, and dump valve. The pressure 

relief valve prevents cabin pressure from exceeding a 

predetermined differential pressure above ambient pressure. 

The vacuum relief prevents ambient pressure from exceeding 

cabin pressure by allowing external air to enter the cabin 

when ambient pressure exceeds cabin pressure. The flight 

deck control switch actuates the dump valve. When this 

switch is positioned to ram, a solenoid valve opens, causing 

the valve to dump cabin air to atmosphere.

The degree of pressurization and the operating altitude of 

the aircraft are limited by several critical design factors. 

Primarily, the fuselage is designed to withstand a particular 

maximum cabin differential pressure. 

Several instruments are used in conjunction with the 

pressurization controller. The cabin differential pressure gauge 

indicates the difference between inside and outside pressure. 

This gauge should be monitored to assure that the cabin does 

not exceed the maximum allowable differential pressure. A 

cabin altimeter is also provided as a check on the performance 

of the system. In some cases, these two instruments are 

combined into one. A third instrument indicates the cabin rate 

of climb or descent. A cabin rate-of-climb instrument and a 

cabin altimeter are illustrated in Figure 6-42. 

Decompression is defined as the inability of the aircraft’s 

pressurization system to maintain its designed pressure 

differential. This can be caused by a malfunction in the 

pressurization system or structural damage to the aircraft. 

Physiologically, decompressions fall into two categories:

• Explosive decompression—a change in cabin pressure 

faster than the lungs can decompress, possibly causing 

lung damage. Normally, the time required to release 

air from the lungs without restrictions, such as 

masks, is 0.2 seconds. Most authorities consider any 
6-34

Figure 6-42. Cabin pressurization instruments.

Cabin pressurization instruments

Cabin rate-of-climb indicator Cabin/differential pressure indicator

4

6

I

0

2

2 I

4

CABIN CLIMB THOUSAND FT PER MIN .5

.5

0

0

5

10

15

20

25

30

35 0

1

2

3 4

5

6

PSI

CABIN ALT

DIFF PRESS

1000 Feet

Cabin pressure 

altitude indicator

(thousands of feet)

Maximum cabin 

differential pressure 

limit

Cabin differential 

pressure indicator

(pounds per square

inch differential)

decompression that occurs in less than 0.5 seconds to 

be explosive and potentially dangerous.

• Rapid decompression—a change in cabin pressure 

in which the lungs decompress faster than the cabin, 

resulting in no likelihood of lung damage.

During an explosive decompression, there may be noise, and 

one may feel dazed for a moment. The cabin air fills with 

fog, dust, or flying debris. Fog occurs due to the rapid drop in 

temperature and the change of relative humidity. Normally, 

the ears clear automatically. Air rushes from the mouth and 

nose due to the escape of air from the lungs, and may be 

noticed by some individuals.

Rapid decompression decreases the period of useful 

consciousness because oxygen in the lungs is exhaled rapidly, 

reducing pressure on the body. This decreases the partial 

pressure of oxygen in the blood and reduces the pilot’s 

effective performance time by one-third to one-fourth its 

normal time. For this reason, an oxygen mask should be worn 

when flying at very high altitudes (35,000 feet or higher). It 

is recommended that the crewmembers select the 100 percent 

oxygen setting on the oxygen regulator at high altitude if 

the aircraft is equipped with a demand or pressure demand 

oxygen system.

The primary danger of decompression is hypoxia. Quick, 

proper utilization of oxygen equipment is necessary to avoid 

unconsciousness. Another potential danger that pilots, crew, 

and passengers face during high altitude decompressions is 

evolved gas decompression sickness. This occurs when the 

pressure on the body drops sufficiently, nitrogen comes out 

of solution, and forms bubbles that can have adverse effects 

on some body tissues. 

Decompression caused by structural damage to the aircraft 

presents another type of danger to pilots, crew, and 

passengers––being tossed or blown out of the aircraft if 

they are located near openings. Individuals near openings 

should wear safety harnesses or seatbelts at all times when 

the aircraft is pressurized and they are seated. Structural 

damage also has the potential to expose them to wind blasts 

and extremely cold temperatures.

Rapid descent from altitude is necessary if these problems are 

to be minimized. Automatic visual and aural warning systems 

are included in the equipment of all pressurized aircraft.

Oxygen Systems

Most high altitude aircraft come equipped with some type 

of fixed oxygen installation. If the aircraft does not have 

a fixed installation, portable oxygen equipment must be 

readily accessible during flight. The portable equipment 

usually consists of a container, regulator, mask outlet, 

and pressure gauge. Aircraft oxygen is usually stored in 

high pressure system containers of 1,800–2,200 psi. When 

the ambient temperature surrounding an oxygen cylinder 

decreases, pressure within that cylinder decreases because 

pressure varies directly with temperature if the volume of 

a gas remains constant. If a drop in indicated pressure on a 

supplemental oxygen cylinder is noted, there is no reason to 

suspect depletion of the oxygen supply, which has simply 

been compacted due to storage of the containers in an 

unheated area of the aircraft. High pressure oxygen containers 

should be marked with the psi tolerance (i.e., 1,800 psi) before 

filling the container to that pressure. The containers should 

be supplied with aviation oxygen only, which is 100 percent 

pure oxygen. Industrial oxygen is not intended for breathing 
6-35

Figure 6-43. Oxygen system regulator.

Figure 6-44. Cannula with green flow detector.

and may contain impurities, and medical oxygen contains 

water vapor that can freeze in the regulator when exposed to 

cold temperatures. To assure safety, periodic inspection and 

servicing of the oxygen system should be done.

An oxygen system consists of a mask or cannula and a 

regulator that supplies a flow of oxygen dependent upon 

cabin altitude. Cannulas are not approved for flights above 

18,000 feet. Regulators approved for use up to 40,000 feet 

are designed to provide zero percent cylinder oxygen and 

100 percent cabin air at cabin altitudes of 8,000 feet or 

less, with the ratio changing to 100 percent oxygen and 

zero percent cabin air at approximately 34,000 feet cabin 

altitude. [Figure 6-43] Regulators approved up to 45,000 feet 

are designed to provide 40 percent cylinder oxygen and 60 

percent cabin air at lower altitudes, with the ratio changing to 

100 percent at the higher altitude. Pilots should avoid flying 

above 10,000 feet without oxygen during the day and above 

8,000 feet at night. 

Pilots should be aware of the danger of fire when using 

oxygen. Materials that are nearly fireproof in ordinary air may 

be susceptible to combustion in oxygen. Oils and greases may 

ignite if exposed to oxygen, and cannot be used for sealing 

the valves and fittings of oxygen equipment. Smoking during 

any kind of oxygen equipment use is prohibited. Before 

each flight, the pilot should thoroughly inspect and test all 

oxygen equipment. The inspection should include a thorough 

examination of the aircraft oxygen equipment, including 

available supply, an operational check of the system, and 

assurance that the supplemental oxygen is readily accessible. 

The inspection should be accomplished with clean hands and 

should include a visual inspection of the mask and tubing 

for tears, cracks, or deterioration; the regulator for valve 

and lever condition and positions; oxygen quantity; and the 

location and functioning of oxygen pressure gauges, flow 

indicators and connections. The mask should be donned and 

the system should be tested. After any oxygen use, verify that 

all components and valves are shut off.

Oxygen Masks

There are numerous types and designs of oxygen masks in 

use. The most important factor in oxygen mask use is to insure 

the masks and oxygen system are compatible. Crew masks 

are fitted to the user’s face with a minimum of leakage and 

usually contain a microphone. Most masks are the oronasal 

type, which covers only the mouth and nose.

A passenger mask may be a simple, cup-shaped rubber 

molding sufficiently flexible to obviate individual fitting. It 

may have a simple elastic head strap or the passenger may 

hold it to his or her face.

All oxygen masks should be kept clean to reduce the danger 

of infection and prolong the life of the mask. To clean the 

mask, wash it with a mild soap and water solution and rinse 

it with clear water. If a microphone is installed, use a clean 

swab, instead of running water, to wipe off the soapy solution. 

The mask should also be disinfected. A gauze pad that has 

been soaked in a water solution of Merthiolate can be used 

to swab out the mask. This solution used should contain 

one-fifth teaspoon of Merthiolate per quart of water. Wipe 

the mask with a clean cloth and air dry.

Cannula

A cannula is an ergonomic piece of plastic tubing which 

runs under the nose and is often used to administer oxygen 

in non-pressurized aircraft. [Figure 6-44] Cannulas are 

typically more comfortable then masks and can be used up 

to 18,000 feet. Altitudes greater than 18,000 feet require the 
6-36

Figure 6-45. Continuous flow mask and rebreather bag.

Figure 6-46. EDS-011 portable pulse-demand oxygen system.

use of an oxygen mask. Many cannulas have a flow meter 

in the line. If equipped, a periodic check of the green flow 

detector should be part of a pilot’s regular scan. 

Diluter-Demand Oxygen Systems

Diluter-demand oxygen systems supply oxygen only when 

the user inhales through the mask. An automix lever allows 

the regulators to automatically mix cabin air and oxygen or 

supply 100 percent oxygen, depending on the altitude. The 

demand mask provides a tight seal over the face to prevent 

dilution with outside air and can be used safely up to 40,000 

feet. A pilot who has a beard or mustache should be sure it is 

trimmed in a manner that will not interfere with the sealing 

of the oxygen mask. The fit of the mask around the beard 

or mustache should be checked on the ground for proper 

sealing.

Pressure-Demand Oxygen Systems

Pressure-demand oxygen systems are similar to diluter 

demand oxygen equipment, except that oxygen is supplied to 

the mask under pressure at cabin altitudes above 34,000 feet. 

Pressure-demand regulators create airtight and oxygen-tight 

seals, but they also provide a positive pressure application of 

oxygen to the mask face piece that allows the user’s lungs 

to be pressurized with oxygen. This feature makes pressure 

demand regulators safe at altitudes above 40,000 feet. Some 

systems may have a pressure demand mask with the regulator 

attached directly to the mask, rather than mounted on the 

instrument panel or other area within the flight deck. The 

mask-mounted regulator eliminates the problem of a long 

hose that must be purged of air before 100 percent oxygen 

begins flowing into the mask.

Continuous-Flow Oxygen System

Continuous-flow oxygen systems are usually provided for 

passengers. The passenger mask typically has a reservoir 

bag, which collects oxygen from the continuous-flow oxygen 

system during the time when the mask user is exhaling. 

The oxygen collected in the reservoir bag allows a higher 

aspiratory flow rate during the inhalation cycle, which 

reduces the amount of air dilution. Ambient air is added to 

the supplied oxygen during inhalation after the reservoir bag 

oxygen supply is depleted. The exhaled air is released to the 

cabin. [Figure 6-45] 

Electrical Pulse-Demand Oxygen System

Portable electrical pulse-demand oxygen systems deliver 

oxygen by detecting an individual’s inhalation effort and 

provide oxygen flow during the initial portion of inhalation. 

Pulse demand systems do not waste oxygen during the 

breathing cycle because oxygen is only delivered during 

inhalation. Compared to continuous-flow systems, the pulsedemand method of oxygen delivery can reduce the amount 

of oxygen needed by 50–85 percent. Most pulse-demand 

oxygen systems also incorporate an internal barometer 

that automatically compensates for changes in altitude by 

increasing the amount of oxygen delivered for each pulse as 

altitude is increased. [Figure 6-46]
6-37

Figure 6-47. Onyx pulse oximeter.

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

Tubes deflated

Tubes inflated

De-icing boots on the leading edge of the wing

Pulse Oximeters

A pulse oximeter is a device that measures the amount of 

oxygen in an individual’s blood, in addition to heart rate. 

This non-invasive device measures the color changes that 

red blood cells undergo when they become saturated with 

oxygen. By transmitting a special light beam through a 

fingertip to evaluate the color of the red cells, a pulse oximeter 

can calculate the degree of oxygen saturation within one 

percent of directly measured blood oxygen. Because of their 

portability and speed, pulse oximeters are very useful for pilots 

operating in nonpressurized aircraft above 12,500 feet where 

supplemental oxygen is required. A pulse oximeter permits 

crewmembers and passengers of an aircraft to evaluate their 

actual need for supplemental oxygen. [Figure 6-47]

Servicing of Oxygen Systems

Before servicing any aircraft with oxygen, consult the 

specific aircraft service manual to determine the type of 

equipment required and procedures to be used. Certain 

precautions should be observed whenever aircraft oxygen 

systems are to be serviced. Oxygen system servicing should 

be accomplished only when the aircraft is located outside 

of the hangars. Personal cleanliness and good housekeeping 

are imperative when working with oxygen. Oxygen under 

pressure and petroleum products create spontaneous results 

when they are brought in contact with each other. Service 

people should be certain to wash dirt, oil, and grease 

(including lip salves and hair oil) from their hands before 

working around oxygen equipment. It is also essential that 

clothing and tools are free of oil, grease, and dirt. Aircraft 

with permanently installed oxygen tanks usually require two 

persons to accomplish servicing of the system. One should 

be stationed at the service equipment control valves, and 

the other stationed where he or she can observe the aircraft 

system pressure gauges. Oxygen system servicing is not 

recommended during aircraft fueling operations or while 

other work is performed that could provide a source of 

ignition. Oxygen system servicing while passengers are on 

board the aircraft is not recommended.

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] 
6-38

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

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]

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] 
6-39

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

Prop ammeter and anti-ice boots

The boot is divided into two sections: inboard and outboard.

When the anti-ice is operating, the inboard section heats on 

each blade, and then cycles to the outboard section. If a 

boot fails to heat properly on one blade, unequal ice loading 

may result, causing severe vibration.

PROP ANTI-ICE BOOT

0 20

10

PROP DEICER

AMPS

When the system is operating, 

the prop ammeter indicates 

normal operating range. As 

each boot section cycles, the 

ammeter fluctuates.

PROP ANTI-ICE AMMETER

Outboard section

Inboard section

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.

Chapter Summary

All aircraft have a requirement for essential systems such 

as the engine, propeller, induction, ignition systems as 

well as the fuel, lubrication, cooling, electrical, landing 

gear, and environmental control systems to support flight. 

Understanding the aircraft systems of the aircraft being 

flown is critical to its safe operation and proper maintenance. 

Consult the AFM/POH for specific information pertaining to 

the aircraft being flown. Various manufacturer and owners 

group websites can also be a valuable source of additional 

information.
6-40