10-1

Introduction

This chapter discusses the factors that affect aircraft 

performance, which include the aircraft weight, atmospheric 

conditions, runway environment, and the fundamental 

physical laws governing the forces acting on an aircraft.

Importance of Performance Data

The performance or operational information section of the 

Aircraft Flight Manual/Pilot’s Operating Handbook (AFM/

POH) contains the operating data for the aircraft; that is, the 

data pertaining to takeoff, climb, range, endurance, descent, 

and landing. The use of this data in flying operations is 

mandatory for safe and efficient operation. Considerable 

knowledge and familiarity of the aircraft can be gained 

through study of this material.

Aircraft 

Performance

Chapter 10
10-2

30

25

20

15

10

 5

 0

Inches of 

Mercury

Millibars

1016

847

677

508

339

 170

 0

29.92"

Standard

Sea Level

Pressure

Hg 1013

Standard

Sea Level

Pressure

mb

A t m o s p h e r i c P r e s s u r e

Figure 10-1. Standard sea level pressure.

It must be emphasized that the manufacturers’ information 

and data furnished in the AFM/POH is not standardized. 

Some provide the data in tabular form, while others use 

graphs. In addition, the performance data may be presented 

on the basis of standard atmospheric conditions, pressure 

altitude, or density altitude. The performance information in 

the AFM/POH has little or no value unless the user recognizes 

those variations and makes the necessary adjustments. 

To be able to make practical use of the aircraft’s capabilities 

and limitations, it is essential to understand the significance 

of the operational data. The pilot must be cognizant of the 

basis for the performance data, as well as the meanings of 

the various terms used in expressing performance capabilities 

and limitations.

Since the characteristics of the atmosphere have a major 

effect on performance, it is necessary to review two dominant 

factors—pressure and temperature.

Structure of the Atmosphere

The atmosphere is an envelope of air that surrounds the Earth 

and rests upon its surface. It is as much a part of the Earth as 

its land and water. However, air differs from land and water 

inasmuch as it is a mixture of gases. It has mass, weight, and 

indefinite shape.

Air, like any other fluid, is able to flow and change its shape 

when subjected to even minute pressures because of the lack of 

strong molecular cohesion. For example, gas will completely 

fill any container into which it is placed, expanding or 

contracting to adjust its shape to the limits of the container. 

The atmosphere is composed of 78 percent nitrogen, 21 percent 

oxygen, and 1 percent other gases, such as argon or helium. 

Most of the oxygen is contained below 35,000 feet altitude.

Atmospheric Pressure

Though there are various kinds of pressure, pilots are mainly 

concerned with atmospheric pressure. It is one of the basic 

factors in weather changes, helps to lift the aircraft, and 

actuates some of the most important flight instruments in 

the aircraft. These instruments often include the altimeter, 

the airspeed indicator (ASI), the vertical speed indicator, and 

the manifold pressure gauge.

Though air is very light, it has mass and is affected by the 

attraction of gravity. Therefore, like any other substance, it 

has weight; because it has weight, it has force. Since it is a 

fluid substance, this force is exerted equally in all directions, 

and its effect on bodies within the air is called pressure. 

Under standard conditions at sea level, the average pressure 

exerted by the weight of the atmosphere is approximately 14.7 

pounds per square inch (psi). The density of air has significant 

effects on the aircraft’s performance. As air becomes less 

dense, it reduces:

• Power, because the engine takes in less air.

• Thrust, because the propeller is less efficient in thin 

air.

• Lift, because the thin air exerts less force on the 

airfoils.

The pressure of the atmosphere varies with time and altitude. 

Due to the changing atmospheric pressure, a standard 

reference was developed. The standard atmosphere at sea level 

is a surface temperature of 59 degrees Fahrenheit (°F) or 15 

degrees Celsius (°C) and a surface pressure of 29.92 inches of 

mercury ("Hg) or 1013.2 millibars (mb). [Figure 10-1]

A standard temperature lapse rate is one in which the 

temperature decreases at the rate of approximately 3.5 °F or 

2 °C per thousand feet up to 36,000 feet. Above this point, 

the temperature is considered constant up to 80,000 feet. A 

standard pressure lapse rate is one in which pressure decreases 

at a rate of approximately 1 "Hg per 1,000 feet of altitude gain 

to 10,000 feet. [Figure 10-2] The International Civil Aviation 

Organization (ICAO) has established this as a worldwide 

standard, and it is often referred to as International Standard 

Atmosphere (ISA) or ICAO Standard Atmosphere. Any 

temperature or pressure that differs from the standard lapse 

rates is considered nonstandard temperature and pressure. 

Adjustments for nonstandard temperatures and pressures are 

provided on the manufacturer’s performance charts. 
10-3

Altitude (ft) Pressure

("Hg)

Temperature

(°C) (°F)

Standard Atmosphere

0

1,000

2,000

3,000

4,000

5,000

6,000

7,000

8,000

9,000

10,000

11,000

12,000

13,000

14,000

15,000

16,000

17,000

18,000

19,000

20,000

29.92

28.86

27.82

26.82

25.84

24.89

23.98

23.09

22.22

21.38

20.57

19.79

19.02

18.29

17.57

16.88

16.21

15.56

14.94

14.33

13.74

15.0

13.0

11.0

9.1

7.1

5.1

3.1

1.1

−0.9

−2.8

−4.8

−6.8

−8.8

−10.8

−12.7

−14.7

−16.7

−18.7

−20.7

−22.6

−24.6

59.0

55.4

51.9

48.3

44.7

41.2

37.6

34.0

30.5

26.9

23.3

19.8

16.2

12.6

9.1

5.5

1.9

−1.6

−5.2

−8.8

−12.3

Figure 10-2. Properties of standard atmosphere.

Since all aircraft performance is compared and evaluated with 

respect to the standard atmosphere, all aircraft instruments 

are calibrated for the standard atmosphere. Thus, certain 

corrections must apply to the instrumentation, as well as the 

aircraft performance, if the actual operating conditions do 

not fit the standard atmosphere. In order to account properly 

for the nonstandard atmosphere, certain related terms must 

be defined.

Pressure Altitude

Pressure altitude is the height above the standard datum 

plane (SDP). The aircraft altimeter is essentially a sensitive 

barometer calibrated to indicate altitude in the standard 

atmosphere. If the altimeter is set for 29.92 "Hg SDP, the 

altitude indicated is the pressure altitude—the altitude in the 

standard atmosphere corresponding to the sensed pressure.

The SDP is a theoretical level where the pressure of the 

atmosphere is 29.92 "Hg and the weight of air is 14.7 psi. As 

atmospheric pressure changes, the SDP may be below, at, or 

above sea level. Pressure altitude is important as a basis for 

determining aircraft performance, as well as for assigning 

flight levels to aircraft operating at above 18,000 feet.

The pressure altitude can be determined by either of two 

methods:

1. By setting the barometric scale of the altimeter to 

29.92 "Hg and reading the indicated altitude, or

2. By applying a correction factor to the indicated altitude 

according to the reported “altimeter setting.”

Density Altitude

The more appropriate term for correlating aerodynamic 

performance in the nonstandard atmosphere is density 

altitude—the altitude in the standard atmosphere 

corresponding to a particular value of air density.

Density altitude is pressure altitude corrected for nonstandard 

temperature. As the density of the air increases (lower 

density altitude), aircraft performance increases. Conversely, 

as air density decreases (higher density altitude), aircraft 

performance decreases. A decrease in air density means a 

high density altitude; an increase in air density means a lower 

density altitude. Density altitude is used in calculating aircraft 

performance. Under standard atmospheric condition, air at 

each level in the atmosphere has a specific density; under 

standard conditions, pressure altitude and density altitude 

identify the same level. Density altitude, then, is the vertical 

distance above sea level in the standard atmosphere at which 

a given density is to be found.

The computation of density altitude must involve consideration 

of pressure (pressure altitude) and temperature. Since aircraft 

performance data at any level is based upon air density under 

standard day conditions, such performance data apply to air 

density levels that may not be identical to altimeter indications. 

Under conditions higher or lower than standard, these levels 

cannot be determined directly from the altimeter.

Density altitude is determined by first finding pressure 

altitude, and then correcting this altitude for nonstandard 

temperature variations. Since density varies directly with 

pressure, and inversely with temperature, a given pressure 

altitude may exist for a wide range of temperature by allowing 

the density to vary. However, a known density occurs for any 

one temperature and pressure altitude. The density of the air, 

of course, has a pronounced effect on aircraft and engine 

performance. Regardless of the actual altitude at which the 

aircraft is operating, it will perform as though it were operating 

at an altitude equal to the existing density altitude.

For example, when set at 29.92 "Hg, the altimeter may 

indicate a pressure altitude of 5,000 feet. According to the 

AFM/POH, the ground run on takeoff may require a distance 

of 790 feet under standard temperature conditions.

However, if the temperature is 20 °C above standard, the 

expansion of air raises the density level. Using temperature 

correction data from tables or graphs, or by deriving the 
10-4

Altimeter

setting

Altitude

correction

28.0

28.1

28.2

28.3

28.4

28.5

28.6

28.7

28.8

28.9

29.0

29.1

29.2

29.3

29.4

29.5

29.6

29.7

29.8

29.9

29.92

30.0

30.1

30.2

30.3

30.4

30.5

30.6

30.7

30.8

30.9

31.0

1,825

1,725

1,630

1,535

1,435

1,340

1,245

1,150

1,050

955

865

770

675

580

485

390

300

205

110

20

0

−75

−165

−255

−350

−440

−530

−620

−710

−805

−895

−965

Subtract Add

To field

elevation

From field

elevation

To get

pressure altitude

Alternate Method for Determining 

Pressure Altitude

Method for Determining 

Pressure Altitude

Field elevation is sea level

Figure 10-3. Field elevation versus pressure. The aircraft is located 

on a field which happens to be at sea level. Set the altimeter to the 

current altimeter setting (29.7). The difference of 205 feet is added 

to the elevation or a PA of 205 feet.

Outside air temperature (OAT)

Density altitude (feet)

15,000

14,000

13,000

12,000

11,000

10,000

9,000

8,000

7,000

6,000

5,000

4,000

3,000

2,000

1,000

Sea level

14,000

13,000

12,000

11,000

10,000

9,000 Pressure altitude (feet)

8,000

7,000

6,000

5,000

4,000

3,000

2,000

1,000

-1,000

Sea level

-2,000

-20° -10° 0° 10° 20° 30° 40°

0° 10° 20° 30° 40° 50° 60° 70° 80° 90° 100°

Standard temperature

C

F

Figure 10-4. Density altitude chart.

density altitude with a computer, it may be found that the 

density level is above 7,000 feet, and the ground run may be 

closer to 1,000 feet.

Air density is affected by changes in altitude, temperature, 

and humidity. High density altitude refers to thin air while 

low density altitude refers to dense air. The conditions that 

result in a high density altitude are high elevations, low 

atmospheric pressures, high temperatures, high humidity, or 

some combination of these factors. Lower elevations, high 

atmospheric pressure, low temperatures, and low humidity 

are more indicative of low density altitude. 

Using a flight computer, density altitude can be computed 

by inputting the pressure altitude and outside air temperature 

at flight level. Density altitude can also be determined by 

referring to the table and chart in Figures 10-3 and 10-4.

Effects of Pressure on Density

Since air is a gas, it can be compressed or expanded. When 

air is compressed, a greater amount of air can occupy a 

given volume. Conversely, when pressure on a given volume 

of air is decreased, the air expands and occupies a greater 

space. That is, the original column of air at a lower pressure 

contains a smaller mass of air. In other words, the density is 

decreased. In fact, density is directly proportional to pressure. 

If the pressure is doubled, the density is doubled, and if the 

pressure is lowered, so is the density. This statement is true 

only at a constant temperature.
10-5

Drag

Speed

Total drag

Parasite drag

Stall

Induced drag

L/DMAX

Figure 10-5. Drag versus speed.

Effects of Temperature on Density

Increasing the temperature of a substance decreases its density. 

Conversely, decreasing the temperature increases the density. 

Thus, the density of air varies inversely with temperature. 

This statement is true only at a constant pressure. 

In the atmosphere, both temperature and pressure decrease 

with altitude, and have conflicting effects upon density. 

However, the fairly rapid drop in pressure as altitude is 

increased usually has the dominant effect. Hence, pilots can 

expect the density to decrease with altitude.

Effects of Humidity (Moisture) on Density

The preceding paragraphs are based on the presupposition of 

perfectly dry air. In reality, it is never completely dry. The 

small amount of water vapor suspended in the atmosphere 

may be negligible under certain conditions, but in other 

conditions humidity may become an important factor in the 

performance of an aircraft. Water vapor is lighter than air; 

consequently, moist air is lighter than dry air. Therefore, as the 

water content of the air increases, the air becomes less dense, 

increasing density altitude and decreasing performance. It is 

lightest or least dense when, in a given set of conditions, it 

contains the maximum amount of water vapor. 

Humidity, also called relative humidity, refers to the amount 

of water vapor contained in the atmosphere, and is expressed 

as a percentage of the maximum amount of water vapor 

the air can hold. This amount varies with the temperature; 

warm air can hold more water vapor, while colder air can 

hold less. Perfectly dry air that contains no water vapor has 

a relative humidity of zero percent, while saturated air that 

cannot hold any more water vapor has a relative humidity 

of 100 percent. Humidity alone is usually not considered an 

essential factor in calculating density altitude and aircraft 

performance; however, it does contribute. 

The higher the temperature, the greater amount of water 

vapor that the air can hold. When comparing two separate air 

masses, the first warm and moist (both qualities making air 

lighter) and the second cold and dry (both qualities making 

it heavier), the first must be less dense than the second. 

Pressure, temperature, and humidity have a great influence 

on aircraft performance because of their effect upon density. 

There is no rule-of-thumb or chart used to compute the effects 

of humidity on density altitude, but it must be taken into 

consideration. Expect a decrease in overall performance in 

high humidity conditions.

Performance

Performance is a term used to describe the ability of an aircraft 

to accomplish certain things that make it useful for certain 

purposes. For example, the ability of an aircraft to land and 

take off in a very short distance is an important factor to the 

pilot who operates in and out of short, unimproved airfields. 

The ability to carry heavy loads, fly at high altitudes at fast 

speeds, or travel long distances is essential performance for 

operators of airline and executive type aircraft.

The primary factors most affected by performance are the 

takeoff and landing distance, rate of climb, ceiling, payload, 

range, speed, maneuverability, stability, and fuel economy. 

Some of these factors are often directly opposed: for example, 

high speed versus short landing distance, long range versus 

great payload, and high rate of climb versus fuel economy. It 

is the preeminence of one or more of these factors that dictates 

differences between aircraft and explains the high degree of 

specialization found in modern aircraft.

The various items of aircraft performance result from the 

combination of aircraft and powerplant characteristics. The 

aerodynamic characteristics of the aircraft generally define 

the power and thrust requirements at various conditions of 

flight, while powerplant characteristics generally define the 

power and thrust available at various conditions of flight. 

The matching of the aerodynamic configuration with the 

powerplant is accomplished by the manufacturer to provide 

maximum performance at the specific design condition (e.g., 

range, endurance, and climb).

Straight-and-Level Flight

All of the principal components of flight performance involve 

steady-state flight conditions and equilibrium of the aircraft. 

For the aircraft to remain in steady, level flight, equilibrium 

must be obtained by a lift equal to the aircraft weight and a 

powerplant thrust equal to the aircraft drag. Thus, the aircraft 

drag defines the thrust required to maintain steady, level 

flight. As presented in Chapter 4, Aerodynamics of Flight, 

all parts of an aircraft contribute to the drag, either induced 

(from lifting surfaces) or parasite drag.

While the parasite drag predominates at high speed, induced 

drag predominates at low speed. [Figure 10-5] For example, 
10-6Power required

Speed

Low cruise speed

Min. speed

High cruise speed

Maximum available power Maximum level flight speed

Figure 10-6. Power versus speed.

Power required

Speed

Speed for max. angle of climb

Stall

Thrust available

Reserve thrust

Thrust required or drag

Figure 10-7. Thrust versus climb angle.

if an aircraft in a steady flight condition at 100 knots is then 

accelerated to 200 knots, the parasite drag becomes four 

times as great, but the power required to overcome that 

drag is eight times the original value. Conversely, when the 

aircraft is operated in steady, level flight at twice as great a 

speed, the induced drag is one-fourth the original value, and 

the power required to overcome that drag is only one-half 

the original value. 

When an aircraft is in steady, level flight, the condition of 

equilibrium must prevail. The unaccelerated condition of flight 

is achieved with the aircraft trimmed for lift equal to weight 

and the powerplant set for a thrust to equal the aircraft drag.

The maximum level flight speed for the aircraft will be obtained 

when the power or thrust required equals the maximum power 

or thrust available from the powerplant. [Figure 10-6] The 

minimum level flight airspeed is not usually defined by thrust 

or power requirement since conditions of stall or stability and 

control problems generally predominate.

Climb Performance

Climb performance is a result of using the aircrafts potential 

energy provided by one, or a combination of two factors. The 

first is the use of excess power above that required for level 

flight. An aircraft equipped with an engine capable of 200 

horsepower (at a given altitude) but using 130 horsepower 

to sustain level flight (at a given airspeed) has 70 excess 

horsepower available for climbing. A second factor is that 

the aircraft can tradeoff its kinetic energy and increase its 

potential energy by reducing its airspeed. The reduction in 

airspeed will increase the aircraft’s potential energy thereby 

also making the aircraft climb. Both terms, power and thrust 

are often used in aircraft performance however, they should 

not be confused.

Although the terms “power” and “thrust” are sometimes 

used interchangeably, erroneously implying that they are 

synonymous, it is important to distinguish between the two 

when discussing climb performance. Work is the product of 

a force moving through a distance and is usually independent 

of time. Work is measured by several standards; the most 

common unit is called a foot-pound. If a one pound mass 

is raised one foot, a work unit of one foot-pound has been 

performed. The common unit of mechanical power is 

horsepower; one horsepower is work equivalent to lifting 

33,000 pounds a vertical distance of one foot in one minute. 

The term power implies work rate or units of work per unit 

of time, and as such is a function of the speed at which the 

force is developed. Thrust, also a function of work, means 

the force that imparts a change in the velocity of a mass. This 

force is measured in pounds but has no element of time or 

rate. It can be said then, that during a steady climb, the rate 

of climb is a function of excess thrust. 

This relationship means that, for a given weight of an aircraft, 

the angle of climb depends on the difference between thrust 

and drag, or the excess power. [Figure 10-7] Of course, when 

the excess thrust is zero, the inclination of the flightpath is 

zero, and the aircraft will be in steady, level flight. When the 

thrust is greater than the drag, the excess thrust will allow 

a climb angle depending on the value of excess thrust. On 

the other hand, when the thrust is less than the drag, the 

deficiency of thrust will allow an angle of descent.

The most immediate interest in the climb angle performance 

involves obstacle clearance. The most obvious purpose for 

which it might be used is to clear obstacles when climbing 

out of short or confined airports.
10-7

Power available and power required

Speed

Speed for max. rate of climb

Power available propeller

Reserve

power

Power required

Power available jet

Figure 10-8. Power versus climb rate.

The maximum angle of climb would occur where there 

exists the greatest difference between thrust available and 

thrust required; i.e., for the propeller-powered airplane, the 

maximum excess thrust and angle of climb will occur at some 

speed just above the stall speed. Thus, if it is necessary to 

clear an obstacle after takeoff, the propeller-powered airplane 

will attain maximum angle of climb at an airspeed close to—if 

not at—the takeoff speed.

Of greater interest in climb performance are the factors that 

affect the rate of climb. The vertical velocity of an aircraft 

depends on the flight speed and the inclination of the 

flightpath. In fact, the rate of climb is the vertical component 

of the flightpath velocity.

For rate of climb, the maximum rate would occur where 

there exists the greatest difference between power available 

and power required. [Figure 10-8] The above relationship 

means that, for a given weight of an aircraft, the rate of climb 

depends on the difference between the power available and 

the power required, or the excess power. Of course, when the 

excess power is zero, the rate of climb is zero and the aircraft 

is in steady, level flight. When power available is greater 

than the power required, the excess power will allow a rate 

of climb specific to the magnitude of excess power. 

During a steady climb, the rate of climb will depend on excess 

power while the angle of climb is a function of excess thrust.

The climb performance of an aircraft is affected by certain 

variables. The conditions of the aircraft’s maximum climb 

angle or maximum climb rate occur at specific speeds, 

and variations in speed will produce variations in climb 

performance. There is sufficient latitude in most aircraft that 

small variations in speed from the optimum do not produce 

large changes in climb performance, and certain operational 

considerations may require speeds slightly different from 

the optimum. Of course, climb performance would be most 

critical with high gross weight, at high altitude, in obstructed 

takeoff areas, or during malfunction of a powerplant. Then, 

optimum climb speeds are necessary.

Weight has a very pronounced effect on aircraft performance. 

If weight is added to an aircraft, it must fly at a higher angle 

of attack (AOA) to maintain a given altitude and speed. This 

increases the induced drag of the wings, as well as the parasite 

drag of the aircraft. Increased drag means that additional 

thrust is needed to overcome it, which in turn means that less 

reserve thrust is available for climbing. Aircraft designers 

go to great effort to minimize the weight since it has such a 

marked effect on the factors pertaining to performance.

A change in an aircraft’s weight produces a twofold effect on 

climb performance. First, a change in weight will change the 

drag and the power required. This alters the reserve power 

available, which in turn, affects both the climb angle and the 

climb rate. Secondly, an increase in weight will reduce the 

maximum rate of climb, but the aircraft must be operated at a 

higher climb speed to achieve the smaller peak climb rate.

An increase in altitude also will increase the power required 

and decrease the power available. Therefore, the climb 

performance of an aircraft diminishes with altitude. The 

speeds for maximum rate of climb, maximum angle of climb, 

and maximum and minimum level flight airspeeds vary with 

altitude. As altitude is increased, these various speeds finally 

converge at the absolute ceiling of the aircraft. At the absolute 

ceiling, there is no excess of power and only one speed will 

allow steady, level flight. Consequently, the absolute ceiling 

of an aircraft produces zero rate of climb. The service ceiling 

is the altitude at which the aircraft is unable to climb at a rate 

greater than 100 feet per minute (fpm). Usually, these specific 

performance reference points are provided for the aircraft at 

a specific design configuration. [Figure 10-9]

In discussing performance, it frequently is convenient to use 

the terms power loading, wing loading, blade loading, and disk 

loading. Power loading is expressed in pounds per horsepower 

and is obtained by dividing the total weight of the aircraft by 

the rated horsepower of the engine. It is a significant factor 

in an aircraft’s takeoff and climb capabilities. Wing loading 

is expressed in pounds per square foot and is obtained by 

dividing the total weight of an airplane in pounds by the wing 

area (including ailerons) in square feet. It is the airplane’s 

wing loading that determines the landing speed. Blade loading 

is expressed in pounds per square foot and is obtained by 

dividing the total weight of a helicopter by the area of the 

rotor blades. Blade loading is not to be confused with disk 

loading, which is the total weight of a helicopter divided by 

the area of the disk swept by the rotor blades. 
10-8 Fuel flow/Power required (HP)

Speed

Reference line

At altitude

A

B Maximum range at L/DMAX

Maximum endurance at

minimum power required Applicable for a particular

 Weight

Altitude

Configuration

Figure 10-10. Airspeed for maximum endurance.

Indicated airspeed (knots)

Standard altitude (feet)

Service ceiling

Absolute ceiling

Best angle

of climb (Vx)

Best rate

of climb (Vy)

24,000

22,000

20,000

18,000

16,000

14,000

12,000

10,000

8,000

6,000

4,000

2,000

Sea level

70 80 90 100 110 120

Figure 10-9. Absolute and service ceiling.

Range Performance

The ability of an aircraft to convert fuel energy into flying 

distance is one of the most important items of aircraft 

performance. In flying operations, the problem of efficient 

range operation of an aircraft appears in two general forms:

1. To extract the maximum flying distance from a given 

fuel load

2. To fly a specified distance with a minimum expenditure 

of fuel

A common element for each of these operating problems is the 

specific range; that is, nautical miles (NM) of flying distance 

versus the amount of fuel consumed. Range must be clearly 

distinguished from the item of endurance. Range involves 

consideration of flying distance, while endurance involves 

consideration of flying time. Thus, it is appropriate to define 

a separate term, specific endurance.

specific endurance = flight hours 

pounds of fuel

or

specific endurance = flight hours/hour 

 pounds of fuel/hour

or

specific endurance = 1 

fuel flow

Fuel flow can be defined in either pounds or gallons. If 

maximum endurance is desired, the flight condition must 

provide a minimum fuel flow. In Figure 10-10 at point A 

the airspeed is low and fuel flow is high. This would occur 

during ground operations or when taking off and climbing. 

As airspeed is increased, power requirements decrease due 

to aerodynamic factors and fuel flow decreases to point B. 

This is the point of maximum endurance. Beyond this point 

increases in airspeed come at a cost. Airspeed increases 

require additional power and fuel flow increases with 

additional power. 

Cruise flight operations for maximum range should be 

conducted so that the aircraft obtains maximum specific range 

throughout the flight. The specific range can be defined by 

the following relationship.

specific range = NM 

pounds of fuel

or

specific range = NM/hour 

 pounds of fuel/hour

or
10-9

Power required

Speed

L/DMAX

Constant altitude

Higher weight

Lower weight

Basic weight

Figure 10-11. Effect of weight.

flight and on performance charts power can be substituted 

for fuel flow. This fact allows for the determination of range 

through analysis of power required versus speed.

The maximum endurance condition would be obtained at the 

point of minimum power required since this would require the 

lowest fuel flow to keep the airplane in steady, level flight. 

Maximum range condition would occur where the ratio of 

speed to power required is greatest. [Figure 10-10]

The maximum range condition is obtained at maximum lift/

drag ratio (L/DMAX), and it is important to note that for a given 

aircraft configuration, the L/DMAX occurs at a particular AOA 

and lift coefficient, and is unaffected by weight or altitude. 

A variation in weight will alter the values of airspeed and 

power required to obtain the L/DMAX. [Figure 10-11]

The variations of speed and power required must be monitored 

by the pilot as part of the cruise control procedure to maintain 

the L/DMAX. When the aircraft’s fuel weight is a small part of 

the gross weight and the aircraft’s range is small, the cruise 

control procedure can be simplified to essentially maintaining 

a constant speed and power setting throughout the time of 

cruise flight. However, a long-range aircraft has a fuel weight 

that is a considerable part of the gross weight, and cruise 

control procedures must employ scheduled airspeed and 

power changes to maintain optimum range conditions.

The variations of speed and power required must be 

monitored by the pilot as part of the cruise control procedure 

to maintain the L/DMAX. When the aircraft’s fuel weight is a 

small part of the gross weight and the aircraft’s range is small, 

the cruise control procedure can be simplified to essentially 

maintaining a constant speed and power setting throughout 

the time of cruise flight. However, a long-range aircraft has a 

fuel weight that is a considerable part of the gross weight, and 

cruise control procedures must employ scheduled airspeed 

and power changes to maintain optimum range conditions.

specific range = knots 

 fuel flow

If maximum specific range is desired, the flight condition must 

provide a maximum of speed per fuel flow. While the peak value 

of specific range would provide maximum range operation, 

long-range cruise operation is generally recommended at some 

slightly higher airspeed. Most long-range cruise operations are 

conducted at the flight condition that provides 99 percent of 

the absolute maximum specific range. The advantage of such 

operation is that one percent of range is traded for three to five 

percent higher cruise speed. Since the higher cruise speed has 

a great number of advantages, the small sacrifice of range is 

a fair bargain. The values of specific range versus speed are 

affected by three principal variables:

1. Aircraft gross weight

2. Altitude

3. The external aerodynamic configuration of the 

aircraft.

These are the source of range and endurance operating data 

included in the performance section of the AFM/POH.

Cruise control of an aircraft implies that the aircraft is operated 

to maintain the recommended long-range cruise condition 

throughout the flight. Since fuel is consumed during cruise, the 

gross weight of the aircraft will vary and optimum airspeed, 

altitude, and power setting can also vary. Cruise control means 

the control of the optimum airspeed, altitude, and power setting 

to maintain the 99 percent maximum specific range condition. 

At the beginning of cruise flight, the relatively high initial 

weight of the aircraft will require specific values of airspeed, 

altitude, and power setting to produce the recommended cruise 

condition. As fuel is consumed and the aircraft’s gross weight 

decreases, the optimum airspeed and power setting may 

decrease, or, the optimum altitude may increase. In addition, 

the optimum specific range will increase. Therefore, the pilot 

must provide the proper cruise control procedure to ensure 

that optimum conditions are maintained.

Total range is dependent on both fuel available and specific 

range. When range and economy of operation are the principal 

goals, the pilot must ensure that the aircraft is operated at the 

recommended long-range cruise condition. By this procedure, 

the aircraft will be capable of its maximum design-operating 

radius, or can achieve flight distances less than the maximum 

with a maximum of fuel reserve at the destination.

A propeller-driven aircraft combines the propeller with the 

reciprocating engine for propulsive power. Fuel flow is 

determined mainly by the shaft power put into the propeller 

rather than thrust. Thus, the fuel flow can be related directly 

to the power required to maintain the aircraft in steady, level 
10-10Power required

Speed

L/DMAX

Constant weight

Sea level

At altitude

Figure 10-12. Effect of altitude on range.

The effect of altitude on the range of a propeller-driven aircraft 

is illustrated in Figure 10-12. A flight conducted at high altitude 

has a greater true airspeed (TAS), and the power required is 

proportionately greater than when conducted at sea level. The 

drag of the aircraft at altitude is the same as the drag at sea level, 

but the higher TAS causes a proportionately greater power 

required. NOTE: The straight line that is tangent to the sea level 

power curve is also tangent to the altitude power curve.

The effect of altitude on specific range also can be appreciated 

from the previous relationships. If a change in altitude causes 

identical changes in speed and power required, the proportion 

of speed to power required would be unchanged. The fact 

implies that the specific range of a propeller-driven aircraft 

would be unaffected by altitude. Actually, this is true to the 

extent that specific fuel consumption and propeller efficiency 

are the principal factors that could cause a variation of 

specific range with altitude. If compressibility effects are 

negligible, any variation of specific range with altitude is 

strictly a function of engine/propeller performance.

An aircraft equipped with a reciprocating engine will 

experience very little, if any, variation of specific range up 

to its absolute altitude. There is negligible variation of brake 

specific fuel consumption for values of brake horsepower 

below the maximum cruise power rating of the engine that 

is the lean range of engine operation. Thus, an increase in 

altitude will produce a decrease in specific range only when 

the increased power requirement exceeds the maximum cruise 

power rating of the engine. One advantage of supercharging 

is that the cruise power may be maintained at high altitude, 

and the aircraft may achieve the range at high altitude with 

the corresponding increase in TAS. The principal differences 

in the high altitude cruise and low altitude cruise are the TAS 

and climb fuel requirements.

Region of Reversed Command

The aerodynamic properties of an aircraft generally determine 

the power requirements at various conditions of flight, while 

the powerplant capabilities generally determine the power 

available at various conditions of flight. When an aircraft 

is in steady, level flight, a condition of equilibrium must 

prevail. An unaccelerated condition of flight is achieved 

when lift equals weight, and the powerplant is set for thrust 

equal to drag. The power required to achieve equilibrium in 

constant-altitude flight at various airspeeds is depicted on a 

power required curve. The power required curve illustrates 

the fact that at low airspeeds near the stall or minimum 

controllable airspeed, the power setting required for steady, 

level flight is quite high.

Flight in the region of normal command means that while 

holding a constant altitude, a higher airspeed requires a higher 

power setting and a lower airspeed requires a lower power 

setting. The majority of aircraft flying (climb, cruise, and 

maneuvers) is conducted in the region of normal command.

Flight in the region of reversed command means flight in 

which a higher airspeed requires a lower power setting and 

a lower airspeed requires a higher power setting to hold 

altitude. It does not imply that a decrease in power will 

produce lower airspeed. The region of reversed command is 

encountered in the low speed phases of flight. Flight speeds 

below the speed for maximum endurance (lowest point 

on the power curve) require higher power settings with a 

decrease in airspeed. Since the need to increase the required 

power setting with decreased speed is contrary to the normal 

command of flight, the regime of flight speeds between the 

speed for minimum required power setting and the stall speed 

(or minimum control speed) is termed the region of reversed 

command. In the region of reversed command, a decrease in 

airspeed must be accompanied by an increased power setting 

in order to maintain steady flight.

Figure 10-13 shows the maximum power available as a 

curved line. Lower power settings, such as cruise power, 

would also appear in a similar curve. The lowest point on 

the power required curve represents the speed at which the 

lowest brake horsepower will sustain level flight. This is 

termed the best endurance airspeed.

An airplane performing a low airspeed, high pitch attitude 

power approach for a short-field landing is an example 

of operating in the region of reversed command. If an 

unacceptably high sink rate should develop, it may be 

possible for the pilot to reduce or stop the descent by applying 

power. But without further use of power, the airplane would 

probably stall or be incapable of flaring for the landing. 
10-11

Power setting

Speed

Best endurance speed

Power required

Region of

reversed

command

Maximum power available

Excess power

Figure 10-13. Power required curve.

Figure 10-14. Takeoff distance chart.

Merely lowering the nose of the airplane to regain flying 

speed in this situation, without the use of power, would result 

in a rapid sink rate and corresponding loss of altitude. 

If during a soft-field takeoff and climb, for example, the pilot 

attempts to climb out of ground effect without first attaining 

normal climb pitch attitude and airspeed, the airplane may 

inadvertently enter the region of reversed command at a 

dangerously low altitude. Even with full power, the airplane 

may be incapable of climbing or even maintaining altitude. 

The pilot’s only recourse in this situation is to lower the pitch 

attitude in order to increase airspeed, which will inevitably 

result in a loss of altitude.

Airplane pilots must give particular attention to precise 

control of airspeed when operating in the low flight speeds 

of the region of reversed command.

Takeoff and Landing Performance

The majority of pilot-caused aircraft accidents occur during 

the takeoff and landing phase of flight. Because of this fact, 

the pilot must be familiar with all the variables that influence 

the takeoff and landing performance of an aircraft and must 

strive for exacting, professional procedures of operation 

during these phases of flight.

Takeoff and landing performance is a condition of accelerated 

and decelerated motion. For instance, during takeoff, an 

aircraft starts at zero speed and accelerates to the takeoff 

speed to become airborne. During landing, the aircraft touches 

down at the landing speed and decelerates to zero speed. The 

important factors of takeoff or landing performance are:

• The takeoff or landing speed is generally a function 

of the stall speed or minimum flying speed.

• The rate of acceleration/deceleration during the 

takeoff or landing roll. The speed (acceleration and 

deceleration) experienced by any object varies directly 

with the imbalance of force and inversely with the 

mass of the object. An airplane on the runway moving 

at 75 knots has four times the energy it has traveling 

at 37 knots. Thus, an airplane requires four times as 

much distance to stop as required at half the speed.

• The takeoff or landing roll distance is a function of 

both acceleration/deceleration and speed.

Runway Surface and Gradient

Runway conditions affect takeoff and landing performance. 

Typically, performance chart information assumes paved, level, 

smooth, and dry runway surfaces. Since no two runways are 

alike, the runway surface differs from one runway to another, 

as does the runway gradient or slope. [Figure 10-14]
10-12

Figure 10-16. Airport/facility directory (A/FD) information.

Airport name

Runway slope and direction of slope

Runway surface

Runway

Figure 10-15. An aircraft’s performance depends greatly on the runway surface.

Runway surfaces vary widely from one airport to another. 

The runway surface encountered may be concrete, asphalt, 

gravel, dirt, or grass. The runway surface for a specific airport 

is noted in the Airport/Facility Directory (A/FD). Any surface 

that is not hard and smooth will increase the ground roll 

during takeoff. This is due to the inability of the tires to roll 

smoothly along the runway. Tires can sink into soft, grassy, 

or muddy runways. Potholes or other ruts in the pavement 

can be the cause of poor tire movement along the runway. 

Obstructions such as mud, snow, or standing water reduce the 

airplane’s acceleration down the runway. Although muddy 

and wet surface conditions can reduce friction between 

the runway and the tires, they can also act as obstructions 

and reduce the landing distance. [Figure 10-15] Braking 

effectiveness is another consideration when dealing with 

various runway types. The condition of the surface affects 

the braking ability of the airplane.

The amount of power that is applied to the brakes without 

skidding the tires is referred to as braking effectiveness. 

Ensure that runways are adequate in length for takeoff 

acceleration and landing deceleration when less than ideal 

surface conditions are being reported.

The gradient or slope of the runway is the amount of change 

in runway height over the length of the runway. The gradient 

is expressed as a percentage such as a 3 percent gradient. This 

means that for every 100 feet of runway length, the runway 

height changes by 3 feet. A positive gradient indicates the 

runway height increases, and a negative gradient indicates the 

runway decreases in height. An upsloping runway impedes 

acceleration and results in a longer ground run during takeoff. 

However, landing on an upsloping runway typically reduces 

the landing roll. A downsloping runway aids in acceleration 

on takeoff resulting in shorter takeoff distances. The opposite 

is true when landing, as landing on a downsloping runway 

increases landing distances. Runway slope information is 

contained in the A/FD. [Figure 10-16]
10-13

Figure 10-17. Tire pressure.

Tire Pressure

Minimum dynamic hydroplaning speed (rounded off) =

9 x Tire pressure (in psi)

9 x 6 = 54 knots

36 = 6

Water on the Runway and Dynamic Hydroplaning

Water on the runways reduces the friction between the tires 

and the ground, and can reduce braking effectiveness. The 

ability to brake can be completely lost when the tires are 

hydroplaning because a layer of water separates the tires from 

the runway surface. This is also true of braking effectiveness 

when runways are covered in ice.

When the runway is wet, the pilot may be confronted with 

dynamic hydroplaning. Dynamic hydroplaning is a condition 

in which the aircraft tires ride on a thin sheet of water rather 

than on the runway’s surface. Because hydroplaning wheels 

are not touching the runway, braking and directional control 

are almost nil. To help minimize dynamic hydroplaning, 

some runways are grooved to help drain off water; most 

runways are not.

Tire pressure is a factor in dynamic hydroplaning. Using 

the simple formula in Figure 10-17, a pilot can calculate 

the minimum speed, in knots, at which hydroplaning will 

begin. In plain language, the minimum hydroplaning speed 

is determined by multiplying the square root of the main gear 

tire pressure in psi by nine. For example, if the main gear tire 

pressure is at 36 psi, the aircraft would begin hydroplaning 

at 54 knots.

Landing at higher than recommended touchdown speeds will 

expose the aircraft to a greater potential for hydroplaning. 

And once hydroplaning starts, it can continue well below the 

minimum initial hydroplaning speed.

On wet runways, directional control can be maximized 

by landing into the wind. Abrupt control inputs should be 

avoided. When the runway is wet, anticipate braking problems 

well before landing and be prepared for hydroplaning. Opt for 

a suitable runway most aligned with the wind. Mechanical 

braking may be ineffective, so aerodynamic braking should 

be used to its fullest advantage.

Takeoff Performance

The minimum takeoff distance is of primary interest in 

the operation of any aircraft because it defines the runway 

requirements. The minimum takeoff distance is obtained by 

taking off at some minimum safe speed that allows sufficient 

margin above stall and provides satisfactory control and 

initial rate of climb. Generally, the lift-off speed is some fixed 

percentage of the stall speed or minimum control speed for the 

aircraft in the takeoff configuration. As such, the lift-off will 

be accomplished at some particular value of lift coefficient 

and AOA. Depending on the aircraft characteristics, the liftoff speed will be anywhere from 1.05 to 1.25 times the stall 

speed or minimum control speed.

To obtain minimum takeoff distance at the specific lift-off 

speed, the forces that act on the aircraft must provide the 

maximum acceleration during the takeoff roll. The various 

forces acting on the aircraft may or may not be under the 

control of the pilot, and various procedures may be necessary 

in certain aircraft to maintain takeoff acceleration at the 

highest value.

The powerplant thrust is the principal force to provide the 

acceleration and, for minimum takeoff distance, the output 

thrust should be at a maximum. Lift and drag are produced 

as soon as the aircraft has speed, and the values of lift and 

drag depend on the AOA and dynamic pressure.

In addition to the important factors of proper procedures, 

many other variables affect the takeoff performance of an 

aircraft. Any item that alters the takeoff speed or acceleration 

rate during the takeoff roll will affect the takeoff distance.

For example, the effect of gross weight on takeoff distance 

is significant and proper consideration of this item must be 

made in predicting the aircraft’s takeoff distance. Increased 

gross weight can be considered to produce a threefold effect 

on takeoff performance:

1. Higher lift-off speed

2. Greater mass to accelerate

3. Increased retarding force (drag and ground friction) 

If the gross weight increases, a greater speed is necessary to 

produce the greater lift necessary to get the aircraft airborne 

at the takeoff lift coefficient. As an example of the effect of 

a change in gross weight, a 21 percent increase in takeoff 

weight will require a 10 percent increase in lift-off speed to 

support the greater weight.
10-14

Figure 10-18. Effect of wind on takeoff and landing.

Percent increase 

in takeoff or 

landing distance

Percent decrease 

in takeoff or 

landing distance

Ratio of wind

velocity to takeoff

or landing speed

Headwind

Tailwind

Reference line

80

70

60

50

40

30

20

10

30% 20% 10%

 10% 20% 30%

10

20

30

40

50

60

Ratio of wind

velocity to takeoff

or landing speed

A change in gross weight will change the net accelerating 

force and change the mass that is being accelerated. If the 

aircraft has a relatively high thrust-to-weight ratio, the change 

in the net accelerating force is slight and the principal effect 

on acceleration is due to the change in mass.

For example, a 10 percent increase in takeoff gross weight 

would cause:

• A 5 percent increase in takeoff velocity.

• At least a 9 percent decrease in rate of acceleration.

• At least a 21 percent increase in takeoff distance.

With ISA conditions, increasing the takeoff weight of the 

average Cessna 182 from 2,400 pounds to 2,700 pounds (11 

percent increase) results in an increased takeoff distance from 

440 feet to 575 feet (23 percent increase).

For the aircraft with a high thrust-to-weight ratio, the 

increase in takeoff distance might be approximately 21 to 

22 percent, but for the aircraft with a relatively low thrustto-weight ratio, the increase in takeoff distance would be 

approximately 25 to 30 percent. Such a powerful effect 

requires proper consideration of gross weight in predicting 

takeoff distance.

The effect of wind on takeoff distance is large, and proper 

consideration also must be provided when predicting takeoff 

distance. The effect of a headwind is to allow the aircraft 

to reach the lift-off speed at a lower groundspeed while the 

effect of a tailwind is to require the aircraft to achieve a 

greater groundspeed to attain the lift-off speed.

A headwind that is 10 percent of the takeoff airspeed will 

reduce the takeoff distance approximately 19 percent. 

However, a tailwind that is 10 percent of the takeoff airspeed 

will increase the takeoff distance approximately 21 percent. In 

the case where the headwind speed is 50 percent of the takeoff 

speed, the takeoff distance would be approximately 25 percent 

of the zero wind takeoff distance (75 percent reduction).

The effect of wind on landing distance is identical to its 

effect on takeoff distance. Figure 10-18 illustrates the general 

effect of wind by the percent change in takeoff or landing 

distance as a function of the ratio of wind velocity to takeoff 

or landing speed.

The effect of proper takeoff speed is especially important 

when runway lengths and takeoff distances are critical. The 

takeoff speeds specified in the AFM/POH are generally 

the minimum safe speeds at which the aircraft can become 

airborne. Any attempt to take off below the recommended 

speed means that the aircraft could stall, be difficult to 

control, or have a very low initial rate of climb. In some cases, 

an excessive AOA may not allow the aircraft to climb out 

of ground effect. On the other hand, an excessive airspeed 

at takeoff may improve the initial rate of climb and “feel” 

of the aircraft, but will produce an undesirable increase in 

takeoff distance. Assuming that the acceleration is essentially 

unaffected, the takeoff distance varies with the square of the 

takeoff velocity. 

Thus, ten percent excess airspeed would increase the takeoff 

distance 21 percent. In most critical takeoff conditions, such 

an increase in takeoff distance would be prohibitive, and the 

pilot must adhere to the recommended takeoff speeds.

The effect of pressure altitude and ambient temperature 

is to define the density altitude and its effect on takeoff 

performance. While subsequent corrections are appropriate 

for the effect of temperature on certain items of powerplant 

performance, density altitude defines specific effects on 

takeoff performance. An increase in density altitude can 

produce a twofold effect on takeoff performance:

1. Greater takeoff speed

2. Decreased thrust and reduced net accelerating force

If an aircraft of given weight and configuration is operated at 

greater heights above standard sea level, the aircraft requires 

the same dynamic pressure to become airborne at the takeoff 

lift coefficient. Thus, the aircraft at altitude will take off at the 

same indicated airspeed (IAS) as at sea level, but because of 

the reduced air density, the TAS will be greater. 
10-15

The effect of density altitude on powerplant thrust depends 

much on the type of powerplant. An increase in altitude 

above standard sea level will bring an immediate decrease in 

power output for the unsupercharged reciprocating engine. 

However, an increase in altitude above standard sea level will 

not cause a decrease in power output for the supercharged 

reciprocating engine until the altitude exceeds the critical 

operating altitude. For those powerplants that experience 

a decay in thrust with an increase in altitude, the effect 

on the net accelerating force and acceleration rate can be 

approximated by assuming a direct variation with density. 

Actually, this assumed variation would closely approximate 

the effect on aircraft with high thrust-to-weight ratios.

Proper accounting of pressure altitude and temperature is 

mandatory for accurate prediction of takeoff roll distance. 

The most critical conditions of takeoff performance are the 

result of some combination of high gross weight, altitude, 

temperature, and unfavorable wind. In all cases, the pilot 

must make an accurate prediction of takeoff distance from 

the performance data of the AFM/POH, regardless of the 

runway available, and strive for a polished, professional 

takeoff procedure.

In the prediction of takeoff distance from the AFM/POH data, 

the following primary considerations must be given:

• Pressure altitude and temperature—to define the effect 

of density altitude on distance 

• Gross weight—a large effect on distance 

• Wind—a large effect due to the wind or wind 

component along the runway 

• Runway slope and condition—the effect of an incline 

and retarding effect of factors such as snow or ice 

Landing Performance

In many cases, the landing distance of an aircraft will define 

the runway requirements for flight operations. The minimum 

landing distance is obtained by landing at some minimum 

safe speed, which allows sufficient margin above stall and 

provides satisfactory control and capability for a go-around. 

Generally, the landing speed is some fixed percentage of 

the stall speed or minimum control speed for the aircraft 

in the landing configuration. As such, the landing will be 

accomplished at some particular value of lift coefficient 

and AOA. The exact values will depend on the aircraft 

characteristics but, once defined, the values are independent 

of weight, altitude, and wind.

To obtain minimum landing distance at the specified landing 

speed, the forces that act on the aircraft must provide maximum 

deceleration during the landing roll. The forces acting on the 

aircraft during the landing roll may require various procedures 

to maintain landing deceleration at the peak value. 

A distinction should be made between the procedures for 

minimum landing distance and an ordinary landing roll with 

considerable excess runway available. Minimum landing 

distance will be obtained by creating a continuous peak 

deceleration of the aircraft; that is, extensive use of the brakes 

for maximum deceleration. On the other hand, an ordinary 

landing roll with considerable excess runway may allow 

extensive use of aerodynamic drag to minimize wear and tear 

on the tires and brakes. If aerodynamic drag is sufficient to 

cause deceleration, it can be used in deference to the brakes 

in the early stages of the landing roll; i.e., brakes and tires 

suffer from continuous hard use, but aircraft aerodynamic 

drag is free and does not wear out with use. The use of 

aerodynamic drag is applicable only for deceleration to 60 

or 70 percent of the touchdown speed. At speeds less than 

60 to 70 percent of the touchdown speed, aerodynamic drag 

is so slight as to be of little use, and braking must be utilized 

to produce continued deceleration. Since the objective during 

the landing roll is to decelerate, the powerplant thrust should 

be the smallest possible positive value (or largest possible 

negative value in the case of thrust reversers). 

In addition to the important factors of proper procedures, 

many other variables affect the landing performance. Any 

item that alters the landing speed or deceleration rate during 

the landing roll will affect the landing distance. 

The effect of gross weight on landing distance is one of the 

principal items determining the landing distance. One effect 

of an increased gross weight is that a greater speed will be 

required to support the aircraft at the landing AOA and lift 

coefficient. For an example of the effect of a change in gross 

weight, a 21 percent increase in landing weight will require 

a ten percent increase in landing speed to support the greater 

weight. 

When minimum landing distances are considered, braking 

friction forces predominate during the landing roll and, for 

the majority of aircraft configurations, braking friction is the 

main source of deceleration.

The minimum landing distance will vary in direct proportion 

to the gross weight. For example, a ten percent increase in 

gross weight at landing would cause a:

• Five percent increase in landing velocity

• Ten percent increase in landing distance

A contingency of this is the relationship between weight and 

braking friction force.
10-16

The effect of wind on landing distance is large and deserves 

proper consideration when predicting landing distance. Since 

the aircraft will land at a particular airspeed independent of 

the wind, the principal effect of wind on landing distance is 

the change in the groundspeed at which the aircraft touches 

down. The effect of wind on deceleration during the landing 

is identical to the effect on acceleration during the takeoff. 

The effect of pressure altitude and ambient temperature is to 

define density altitude and its effect on landing performance. 

An increase in density altitude increases the landing speed 

but does not alter the net retarding force. Thus, the aircraft 

at altitude lands at the same IAS as at sea level but, because 

of the reduced density, the TAS is greater. Since the aircraft 

lands at altitude with the same weight and dynamic pressure, 

the drag and braking friction throughout the landing roll have 

the same values as at sea level. As long as the condition is 

within the capability of the brakes, the net retarding force 

is unchanged, and the deceleration is the same as with the 

landing at sea level. Since an increase in altitude does not alter 

deceleration, the effect of density altitude on landing distance 

is due to the greater TAS.

The minimum landing distance at 5,000 feet is 16 percent 

greater than the minimum landing distance at sea level. The 

approximate increase in landing distance with altitude is 

approximately three and one-half percent for each 1,000 feet 

of altitude. Proper accounting of density altitude is necessary 

to accurately predict landing distance.

The effect of proper landing speed is important when runway 

lengths and landing distances are critical. The landing speeds 

specified in the AFM/POH are generally the minimum safe 

speeds at which the aircraft can be landed. Any attempt to land 

at below the specified speed may mean that the aircraft may 

stall, be difficult to control, or develop high rates of descent. 

On the other hand, an excessive speed at landing may improve 

the controllability slightly (especially in crosswinds), but 

causes an undesirable increase in landing distance. 

A ten percent excess landing speed causes at least a 21 percent 

increase in landing distance. The excess speed places a greater 

working load on the brakes because of the additional kinetic 

energy to be dissipated. Also, the additional speed causes 

increased drag and lift in the normal ground attitude, and the 

increased lift reduces the normal force on the braking surfaces. 

The deceleration during this range of speed immediately after 

touchdown may suffer, and it is more probable for a tire to be 

blown out from braking at this point.

The most critical conditions of landing performance are 

combinations of high gross weight, high density altitude, 

and unfavorable wind. These conditions produce the 

greatest required landing distances and critical levels of 

energy dissipation required of the brakes. In all cases, it is 

necessary to make an accurate prediction of minimum landing 

distance to compare with the available runway. A polished, 

professional landing procedure is necessary because the 

landing phase of flight accounts for more pilot-caused aircraft 

accidents than any other single phase of flight.

In the prediction of minimum landing distance from the AFM/

POH data, the following considerations must be given:

• Pressure altitude and temperature—to define the effect 

of density altitude

• Gross weight—which defines the CAS for landing.

• Wind—a large effect due to wind or wind component 

along the runway

• Runway slope and condition—relatively small 

correction for ordinary values of runway slope, but a 

significant effect of snow, ice, or soft ground

A tail wind of ten knots increases the landing distance by 

about 21 percent. An increase of landing speed by ten percent 

increases the landing distance by 20 percent. Hydroplaning 

makes braking ineffective until a decrease of speed to that 

determined using Figure 10-17. 

For instance, a pilot is downwind for runway 18, and the 

tower asks if runway 27 could be accepted. There is a light 

rain and the winds are out of the east at ten knots. The pilot 

accepts because he or she is approaching the extended 

centerline of runway 27. The turn is tight and the pilot must 

descend (dive) to get to runway 27. After becoming aligned 

with the runway and at 50 feet AGL, the pilot is already 1,000 

feet down the 3,500 feet runway. The airspeed is still high by 

about ten percent (should be at 70 knots and is at about 80 

knots). The wind of ten knots is blowing from behind.

First, the airspeed being high by about ten percent (80 knots 

versus 70 knots), as presented in the performance chapter, 

results in a 20 percent increase in the landing distance. 

In performance planning, the pilot determined that at 70 

knots the distance would be 1,600 feet. However, now it 

is increased by 20 percent and the required distance is now 

1,920 feet. 

The newly revised landing distance of 1,920 feet is also 

affected by the wind. In looking at Figure 10-18, the affect 

of the wind is an additional 20 percent for every ten miles 

per hour (mph) in wind. This is computed not on the original 

estimate but on the estimate based upon the increased 

airspeed. Now the landing distance is increased by another 
10-17

320 feet for a total requirement of 2,240 feet to land the 

airplane after reaching 50 feet AGL. 

That is the original estimate of 1,600 under planned 

conditions plus the additional 640 feet for excess speed 

and the tailwind. Given the pilot overshot the threshhold 

by 1,000 feet, the total length required is 3, 240 on a 3,500 

foot runway; 260 feet to spare. But this is in a perfect 

environment. Most pilots become fearful as the end of the 

runway is facing them just ahead. A typical pilot reaction 

is to brake—and brake hard. Because the aircraft does not 

have antilock braking features like a car, the brakes lock, 

and the aircraft hydroplanes on the wet surface of the runway 

until decreasing to a speed of about 54 knots (the square 

root of the tire pressure (√36) x 9). Braking is ineffective 

when hydroplaning.

The 260 feet that a pilot might feel is left over has long since 

evaporated as the aircraft hydroplaned the first 300–500 feet 

when the brakes locked. This is an example of a true story, 

but one which only changes from year to year because of new 

participants and aircraft with different N-numbers. 

In this example, the pilot actually made many bad decisions. 

Bad decisions, when combined, have a synergy greater 

than the individual errors. Therefore, the corrective actions 

become larger and larger until correction is almost impossible. 

Aeronautical decision-making will be discussed more fully in 

Chapter 17, Aeronautical Decision-Making (ADM).

Performance Speeds

True Airspeed (TAS)—the speed of the aircraft in relation to 

the air mass in which it is flying.

Indicated Airspeed (IAS)—the speed of the aircraft as 

observed on the ASI. It is the airspeed without correction 

for indicator, position (or installation), or compressibility 

errors.

Calibrated Airspeed (CAS)—the ASI reading corrected for 

position (or installation), and instrument errors. (CAS is 

equal to TAS at sea level in standard atmosphere.) The color 

coding for various design speeds marked on ASIs may be 

IAS or CAS.

Equivalent Airspeed (EAS)—the ASI reading corrected 

for position (or installation), or instrument error, and for 

adiabatic compressible flow for the particular altitude. (EAS 

is equal to CAS at sea level in standard atmosphere.)

VS0—the calibrated power-off stalling speed or the minimum 

steady flight speed at which the aircraft is controllable in the 

landing configuration.

VS1—the calibrated power-off stalling speed or the minimum 

steady flight speed at which the aircraft is controllable in a 

specified configuration.

VY—the speed at which the aircraft will obtain the maximum 

increase in altitude per unit of time. This best rate-of-climb 

speed normally decreases slightly with altitude.

VX—the speed at which the aircraft will obtain the highest 

altitude in a given horizontal distance. This best angle-ofclimb speed normally increases slightly with altitude.

VLE—the maximum speed at which the aircraft can be safely 

flown with the landing gear extended. This is a problem 

involving stability and controllability. 

VLO—the maximum speed at which the landing gear can 

be safely extended or retracted. This is a problem involving 

the air loads imposed on the operating mechanism during 

extension or retraction of the gear.

VFE—the highest speed permissible with the wing flaps in a 

prescribed extended position. This is because of the air loads 

imposed on the structure of the flaps.

VA—the calibrated design maneuvering airspeed. This is 

the maximum speed at which the limit load can be imposed 

(either by gusts or full deflection of the control surfaces) 

without causing structural damage. Operating at or below 

manuevering speed does not provide structural protection 

against multiple full control inputs in one axis or full control 

inputs in more than one axis at the same time. 

VNO—the maximum speed for normal operation or the 

maximum structural cruising speed. This is the speed at 

which exceeding the limit load factor may cause permanent 

deformation of the aircraft structure.

VNE—the speed which should never be exceeded. If flight is 

attempted above this speed, structural damage or structural 

failure may result.

Performance Charts

Performance charts allow a pilot to predict the takeoff, climb, 

cruise, and landing performance of an aircraft. These charts, 

provided by the manufacturer, are included in the AFM/POH. 

Information the manufacturer provides on these charts has 

been gathered from test flights conducted in a new aircraft, 

under normal operating conditions while using average 

piloting skills, and with the aircraft and engine in good 

working order. Engineers record the flight data and create 

performance charts based on the behavior of the aircraft 

during the test flights. By using these performance charts, 
10-18

Figure 10-19. Conditions notes chart.

,

, ,

a pilot can determine the runway length needed to take off 

and land, the amount of fuel to be used during flight, and the 

time required to arrive at the destination. It is important to 

remember that the data from the charts will not be accurate 

if the aircraft is not in good working order or when operating 

under adverse conditions. Always consider the necessity to 

compensate for the performance numbers if the aircraft is not 

in good working order or piloting skills are below average. 

Each aircraft performs differently and, therefore, has different 

performance numbers. Compute the performance of the 

aircraft prior to every flight, as every flight is different. (See 

appendix for examples of performance charts for a Cessna 

Model 172R and Challenger 605.) 

Every chart is based on certain conditions and contains 

notes on how to adapt the information for flight conditions. 

It is important to read every chart and understand how to 

use it. Read the instructions provided by the manufacturer. 

For an explanation on how to use the charts, refer to the 

example provided by the manufacturer for that specific chart. 

[Figure 10-19] 

The information manufacturers furnish is not standardized. 

Information may be contained in a table format, and other 

information may be contained in a graph format. Sometimes 

combined graphs incorporate two or more graphs into one chart 

to compensate for multiple conditions of flight. Combined 

graphs allow the pilot to predict aircraft performance for 

variations in density altitude, weight, and winds all on one 

chart. Because of the vast amount of information that can be 

extracted from this type of chart, it is important to be very 

accurate in reading the chart. A small error in the beginning 

can lead to a large error at the end.

The remainder of this section covers performance information 

for aircraft in general and discusses what information the 

charts contain and how to extract information from the charts 

by direct reading and interpolation methods. Every chart 

contains a wealth of information that should be used when 

flight planning. Examples of the table, graph, and combined 

graph formats for all aspects of flight will be discussed.

Interpolation

Not all of the information on the charts is easily extracted. 

Some charts require interpolation to find the information for 

specific flight conditions. Interpolating information means 

that by taking the known information, a pilot can compute 

intermediate information. However, pilots sometimes round 

off values from charts to a more conservative figure. 

Using values that reflect slightly more adverse conditions 

provides a reasonable estimate of performance information 

and gives a slight margin of safety. The following illustration 

is an example of interpolating information from a takeoff 

distance chart. [Figure 10-20]

Density Altitude Charts

Use a density altitude chart to figure the density altitude at the 

departing airport. Using Figure 10-21, determine the density 

altitude based on the given information.

Sample Problem 1

Airport Elevation...............................................5,883 feet

OAT...........................................................................70 °F

Altimeter...........................................................30.10" Hg
10-19

Figure 10-20. Interpolating charts.

TAKEOFF DISTANCE

MAXIMUM WEIGHT 2,400 LB

Conditions

Weight

(lb)

2,400

Press

ALT

(feet)

S.L.

1,000

2,000

3,000

4,000

5,000

6,000

7,000

8,000

Grnd

Roll 

(feet)

795

875

960

1,055

1,165

1,285

1,425

1,580

1,755

Total feet

to clear

50 ft OBS

1,460

1,605

1,770

1,960

2,185

2,445

2,755

3,140

3,615

Grnd

Roll 

(feet)

860

940

1,035

1,140

1,260

1,390

1,540

1,710

1,905

Total feet

to clear

50 ft OBS

1,570

1,725

1,910

2,120

2,365

2,660

3,015

3,450

4,015

Grnd

Roll 

(feet)

995

1,090

1,200

1,325

1,465

1,620

1,800

2,000

- - - 

Total feet

to clear

50 ft OBS

1,810

2,000

2,220

2,480

2,790

3,160

3,620

4,220

- - -

Grnd

Roll 

(feet)

1,065

1,170

1,290

1,425

1,575

1,745

1,940

- - - 

- - - 

Total feet

to clear

50 ft OBS

1,945

2,155

2,395

2,685

3,030

3,455

3,990

- - - 

- - - 

Lift

off

51

AT

50 ft

56 

Takeoff

speed KIAS

0 °C 10 °C 20 °C 30 °C

Grnd

Roll 

(feet)

925

1,015

1,115

1,230

1,355

1,500

1,665

1,850

2,060

Total feet

to clear

50 ft OBS

1,685

1,860

2,060

2,295

2,570

2,895

3,300

3,805

4,480

40 °C

Flaps 10°

Full throttle prior to brake release

Paved level runway

Zero wind

To find the takeoff distance for a pressure altitude of 2,500 feet 

at 20 °C, average the ground roll for 2,000 feet and 3,000 feet.

1,115 + 1,230

2

= 1,173 feet

Figure 10-21. Density altitude chart.

Density Altitude Chart

Outside air temperature

Approximate density altitude (thousand feet)

15

14

13

12

11

10

9

8

7

6

5

4

3

2

1

S.L.

14,000

13,000

12,000

11,000

10,000

9,000 Pressure altitude (feet)

8,000

7,000

6,000

5,000

4,000

3,000

2,000

1,000

-1,000

Sea level

-18 -12° -7° -1° 4° 10° 16° 21° 27° 32° 38°

0° 10° 20° 30° 40° 50° 60° 70° 80° 90° 100°

Standard temperature

F

C

28.0 1,824

28.1 1,727

28.2 1,630

28.3 1,533

28.4 1,436

28.5 1,340

28.6 1,244

28.7 1,148

28.8 1,053

28.9 957

29.0 863

29.1 768

29.2 673

29.3 579

29.4 485

29.5 392

29.6 298

29.7 205

29.8 112

29.9 20

29.92 0

30.0 −73

30.1 −165

30.2 −257

30.3 −348

30.4 −440

30.5 −531

30.6 −622

30.7 −712

30.8 −803 Altimeter setting ("Hg) Pressure altitude conversion factor

First, compute the pressure altitude conversion. Find 30.10 

under the altimeter heading. Read across to the second 

column. It reads “–165.” Therefore, it is necessary to subtract 

165 from the airport elevation giving a pressure altitude of 

5,718 feet. Next, locate the outside air temperature on the 

scale along the bottom of the graph. From 70°, draw a line up 

to the 5,718 feet pressure altitude line, which is about twothirds of the way up between the 5,000 and 6,000 foot lines. 

Draw a line straight across to the far left side of the graph 

and read the approximate density altitude. The approximate 

density altitude in thousands of feet is 7,700 feet. 

Takeoff Charts

Takeoff charts are typically provided in several forms and allow 

a pilot to compute the takeoff distance of the aircraft with no 

flaps or with a specific flap configuration. A pilot can also 

compute distances for a no flap takeoff over a 50 foot obstacle 

scenario, as well as with flaps over a 50 foot obstacle. The 

takeoff distance chart provides for various aircraft weights, 

altitudes, temperatures, winds, and obstacle heights.

Sample Problem 2

Pressure Altitude...............................................2,000 feet

OAT..........................................................................22 °C

Takeoff Weight.............................................2,600 pounds

Headwind...............................................................6 knots

Obstacle Height.......................................50 foot obstacle

Refer to Figure 10-22. This chart is an example of a combined 

takeoff distance graph. It takes into consideration pressure 

altitude, temperature, weight, wind, and obstacles all on one 

chart. First, find the correct temperature on the bottom left-
10-20

Figure 10-22. Takeoff distance graph.

 C -40° -30° -20° -10° 0° 10° 20° 30° 40° 50° 2,800 2,600 2,400 2,200 0 10 20 30 0 50

 Outside air temperature Weight Wind component Obstacle

 (pounds) (knots) height (feet)

 Outside air temperature Weight 

 (pounds) 

 Wind component 

 (knots) 

 Obstacle

height (feet) F -40° -20° 0° 20° 40° 60° 80° 100° 120°

6,000

5,000

4,000

3,000

2,000

1,000

0

10,000

8,000

6,000

4,000

2,000

S.L.

Guide lines not applicable for

Intermediate

Tailwind

Headwind

Reference lineReference lineReference line

Pressure altitude - feet ISA

Obstacle heights

Weight

pounds

 kts MPH

Takeoff speed

Lift-off 50 ft

 kts MPH

2,950

2,800

2,600

2,400

2,200

66

64

63

61

58

76

74

72

70

67

72

70

68

66

63

83

81

78

76

73

Associated conditions

Power Full throttle 2,600 rpm

Mixture Lean to appropriate fuel 

pressure

Flaps Up

Landing Retract after positive 

gear climb established

Cowl Open

flaps

Pressure altitude - feet ISA de - feet altitud ssure Pres

et ISA

ure a s Pre sure alt tud eet e f - ISA

hand side of the graph. Follow the line from 22° C straight 

up until it intersects the 2,000 foot altitude line. From that 

point, draw a line straight across to the first dark reference 

line. Continue to draw the line from the reference point in 

a diagonal direction following the surrounding lines until it 

intersects the corresponding weight line. From the intersection 

of 2,600 pounds, draw a line straight across until it reaches the 

second reference line. Once again, follow the lines in a diagonal 

manner until it reaches the six knot headwind mark. Follow 

straight across to the third reference line and from here, draw 

a line in two directions. First, draw a line straight across to 

figure the ground roll distance. Next, follow the diagonal lines 

again until it reaches the corresponding obstacle height. In this 

case, it is a 50 foot obstacle. Therefore, draw the diagonal line 

to the far edge of the chart. This results in a 600 foot ground 

roll distance and a total distance of 1,200 feet over a 50 foot 

obstacle. To find the corresponding takeoff speeds at lift-off 

and over the 50 foot obstacle, refer to the table on the top of 

the chart. In this case, the lift-off speed at 2,600 pounds would 

be 63 knots and over the 50 foot obstacle would be 68 knots.

Sample Problem 3

Pressure Altitude...............................................3,000 feet

OAT.........................................................................30 °C

Takeoff Weight............................................2,400 pounds

Headwind............................................................18 knots

Refer to Figure 10-23. This chart is an example of a takeoff 

distance table for short-field takeoffs. For this table, first find 

the takeoff weight. Once at 2,400 pounds, begin reading from 

left to right across the table. The takeoff speed is in the second 

column and, in the third column under pressure altitude, find 

the pressure altitude of 3,000 feet. Carefully follow that line 

to the right until it is under the correct temperature column 

of 30 °C. The ground roll total reads 1,325 feet and the total 

required to clear a 50 foot obstacle is 2,480 feet. At this point, 

there is an 18 knot headwind. According to the notes section 

under point number two, decrease the distances by ten percent 

for each 9 knots of headwind. With an 18 knot headwind, it 

is necessary to decrease the distance by 20 percent. Multiply 

1,325 feet by 20 percent (1,325 x .20 = 265), subtract the 

product from the total distance (1,325 – 265 = 1,060). Repeat 

this process for the total distance over a 50 foot obstacle. The 

ground roll distance is 1,060 feet and the total distance over 

a 50 foot obstacle is 1,984 feet.

Climb and Cruise Charts

Climb and cruise chart information is based on actual 

flight tests conducted in an aircraft of the same type. This 

information is extremely useful when planning a crosscountry to predict the performance and fuel consumption of 

the aircraft. Manufacturers produce several different charts 

for climb and cruise performance. These charts include 

everything from fuel, time, and distance to climb, to best 

power setting during cruise, to cruise range performance.

The first chart to check for climb performance is a fuel, 

time, and distance-to-climb chart. This chart will give the 

fuel amount used during the climb, the time it will take to 

accomplish the climb, and the ground distance that will 

be covered during the climb. To use this chart, obtain the 
10-21

Figure 10-23. Takeoff distance short field charts.

TAKEOFF DISTANCE

MAXIMUM WEIGHT 2,400 LB

Notes Conditions

Weight

(lb)

2,400

2,200

2,000

Press

ALT

(FT)

S.L.

1,000

2,000

3,000

4,000

5,000

6,000

7,000

8,000

S.L.

1,000

2,000

3,000

4,000

5,000

6,000

7,000

8,000

S.L.

1,000

2,000

3,000

4,000

5,000

6,000

7,000

8,000

Grnd

Roll 

(FT) 

795

875

960

1,055

1,165

1,285

1,425

1,580

1,755

650

710

780

855

945

1,040

1,150

1,270

1,410

525

570

625

690

755

830

920

1,015

1,125

Total feet

to clear

50 ft OBS

1,460

1,605

1,770

1,960

2,185

2,445

2,755

3,140

3,615

1,195

1,310

1,440

1,585

1,750

1,945

2,170

2,440

2,760

 970

1,060

1,160

1,270

1,400

1,545

1,710

1,900

2,125

Grnd

Roll 

(FT) 

860

940

1,035

1,140

1,260

1,390

1,540

1,710

1,905

700

765

840

925

1,020

1,125

1,240

1,375

1,525

565

615

675

740

815

900

990

1,095

1,215

Total feet

to clear

50 ft OBS

1,570

1,725

1,910

2,120

2,365

2,660

3,015

3,450

4,015

1,280

1,405

1,545

1,705

1,890

2,105

2,355

2,655

3,015

1,035

1,135

1,240

1,365

1,500

1,660

1,845

2,055

2,305

Grnd

Roll 

(FT) 

995

1,090

1,200

1,325

1,465

1,620

1,800

2,000

� � � 

805

885

975

1,070

1,180

1,305

1,445

1,605

1,785

650

710

780

860

945

2,145

2,405

2,715

1,410

Total feet

to clear

50 ft OBS

1,810

2,000

2,220

2,480

2,790

3,160

3,620

4,220

- - -

1,470

1,615

1,785

1,975

2,200

2,465

2,775

3,155

3,630

1,185

1,295

1,425

1,570

1,735

1,925

2,145

2,405

2,715

Grnd

Roll 

(FT) 

1,065

1,170

1,290

1,425

1,575

1,745

1,940

- - - 

- - - 

865

 950

1,045

1,150

1,270

1,405

1,555

1,730

1,925

695

765

840

920

1,015

1,120

1,235

1,370

1,520

Total feet

to clear

50 ft OBS

1,945

2,155

2,395

2,685

3,030

3,455

3,990

- - - 

- - - 

1,575

1,735

1,915

2,130

2,375

2,665

3,020

3,450

4,005

1,265

1,385

1,525

1,685

1,865

2,070

2,315

2,605

2,950

Lift

off

51

49

46

AT

50 ft

56

54

51 

Takeoff

speed KIAS

0 °C 10 °C 20 °C 30 °C

Grnd

Roll 

(FT) 

925

1,015

1,115

1,230

1,355

1,500

1,665

1,850

2,060

750

825

905

995

1,100

1,210

1,340

1,485

1,650

605

665

725

800

880

970

1,070

1,180

1,310

Total feet

to clear

50 ft OBS

1,685

1,860

2,060

2,295

2,570

2,895

3,300

3,805

4,480

1,375

1,510

1,660

1,835

2,040

2,275

2,555

2,890

3,305

1,110

1,215

1,330

1,465

1,615

1,790

1,990

2,225

2,500

40 °C

SHORT FIELD

Flaps 10°

Full throttle prior to brake release

Paved level runway

Zero wind

1. Prior to takeoff from fields above 3,000 feet elevation, the mixture should be leaned to give maximum rpm in a full throttle, static runup.

2. Decrease distances 10% for each 9 knots headwind. For operation with tailwind up to 10 knots, increase distances by 10% for each 2 knots.

3. For operation on a dry, grass runway, increase distances by 15% of the “ground roll” figure.

Figure 10-24. Fuel time distance climb chart.

20,000

18,000

16,000

14,000

12,000

10,000

8,000

6,000

4,000

2,000

Sea level

-40° -30° -20° -10° 0° 10° 20° 30° 40°C 0 10 20 30 40 50

 Outside air temperature Fuel, time and distance to climb

Associated conditions

Maximum continuous power*

3,600 lb gross weight

Flaps up

90 KIAS

No wind

* 2,700 rpm & 36 in M.P. (3-blade prop)

2,575 rpm & 36 in M.P. (2-blade prop)

Pressure ALT - feet

Fuel - gallons

Time - minutes

Distance - nautical miles

Time - minutes

Distance - nautical miles

Cruise

Departure

information for the departing airport and for the cruise 

altitude. Using Figure 10-24, calculate the fuel, time, and 

distance to climb based on the information provided.
10-22

Figure 10-25. Fuel time distance climb. Notes Conditions

Weight

(pounds)

Rate of

climb

FPM

Press

ALT

(feet) Time

(minutes)

Fuel used

(pounds)

Distance

(nautical

miles)

From sea level

S.L.

4,000

8,000

12,000

16,000

20,000

S.L.

4,000

8,000

12,000

16,000

20,000

S.L.

4,000

8,000

12,000

16,000

20,000

4,000

3,700

3,400

605

570

530

485

430

365

700

665

625

580

525

460

810

775

735

690

635

565

0

7

14

22

31

41

0

6

12

19

26

34

0

5

10

16

22

29

0

14

28

44

62

82

0

12

24

37

52

68

0

10

21

32

44

57

0

13

27

43

63

87

0

11

23

37

53

72

0

9

20

31

45

61

NORMAL CLIMB

 110 KIAS

Flaps up

Gear up

2,500 RPM

30" Hg

120 PPH fuel flow

Cowl flaps open

Standard temperature

1. Add 16 pounds of fuel for engine start, taxi, and takeoff allowance.

2. Increase time, fuel, and distance by 10% for each 7°C above standard 

temperature.

3. Distances shown are based on zero wind.

Sample Problem 4

Departing Airport Pressure Altitude.................6,000 feet

Departing Airport OAT............................................25 °C 

Cruise Pressure Altitude..................................10,000 feet 

Cruise OAT..............................................................10 °C 

First, find the information for the departing airport. Find the 

OAT for the departing airport along the bottom, left-hand side 

of the graph. Follow the line from 25 °C straight up until it 

intersects the line corresponding to the pressure altitude of 

6,000 feet. Continue this line straight across until it intersects 

all three lines for fuel, time, and distance. Draw a line straight 

down from the intersection of altitude and fuel, altitude and 

time, and a third line at altitude and distance. It should read 

three and one-half gallons of fuel, 6.5 minutes of time, and 

nine NM. Next, repeat the steps to find the information for 

the cruise altitude. It should read six and one-half gallons 

of fuel, 11.5 minutes of time, and 15 NM. Take each set of 

numbers for fuel, time, and distance and subtract them from 

one another (6.5 – 3.5 = 3 gallons of fuel). It will take three 

gallons of fuel and 5 minutes of time to climb to 10,000 

feet. During that climb, the distance covered is six NM. 

Remember, according to the notes at the top of the chart, these 

numbers do not take into account wind, and it is assumed 

maximum continuous power is being used.

The next example is a fuel, time, and distance-to-climb table. 

For this table, use the same basic criteria as for the previous 

chart. However, it is necessary to figure the information in a 

different manner. Refer to Figure 10-25 to work the following 

sample problem.

Sample Problem 5

Departing Airport Pressure Altitude..................Sea level

Departing Airport OAT............................................22 °C

Cruise Pressure Altitude....................................8,000 feet

Takeoff Weight.............................................3,400 pounds

To begin, find the given weight of 3,400 in the first column of 

the chart. Move across to the pressure altitude column to find 

the sea level altitude numbers. At sea level, the numbers read 

zero. Next, read the line that corresponds with the cruising 

altitude of 8,000 feet. Normally, a pilot would subtract these 

two sets of number from one another, but given the fact that 

the numbers read zero at sea level, it is known that the time 

to climb from sea level to 8,000 feet is 10 minutes. It is also 

known that 21 pounds of fuel will be used and 20 NM will be 

covered during the climb. However, the temperature is 22 °C, 

which is 7° above the standard temperature of 15 °C. The notes 

section of this chart indicate that the findings must be increased 

by ten percent for each 7° above standard. Multiply the findings 

by ten percent or .10 (10 x .10 = 1, 1 + 10 = 11 minutes). After 

accounting for the additional ten percent, the findings should 

read 11 minutes, 23.1 pounds of fuel, and 22 NM. Notice that 

the fuel is reported in pounds of fuel, not gallons. Aviation fuel 

weighs six pounds per gallon, so 23.1 pounds of fuel is equal 

to 3.85 gallons of fuel (23.1 ÷ 6 = 3.85).

The next example is a cruise and range performance chart. 

This type of table is designed to give TAS, fuel consumption, 

endurance in hours, and range in miles at specific cruise 

configurations. Use Figure 10-26 to determine the cruise and 

range performance under the given conditions.

Sample Problem 6

Pressure Altitude...............................................5,000 feet

RPM..................................................................2,400 rpm

Fuel Carrying Capacity..................38 gallons, no reserve

Find 5,000 feet pressure altitude in the first column on the 

left-hand side of the table. Next, find the correct rpm of 2,400 

in the second column. Follow that line straight across and 

read the TAS of 116 mph, and a fuel burn rate of 6.9 gallons 

per hour. As per the example, the aircraft is equipped with 

a fuel carrying capacity of 38 gallons. Under this column, 
10-23

Figure 10-26. Cruise and range performance. Notes Conditions

ALT 

2,500

5,000

7,500

10,000

GAL/

Hour

9.7

8.6

7.8

7.2

6.7

6.3

9.0

8.1

7.4

6.9

6.5

6.0

8.4

7.7

7.1

6.7

6.2

7.6

7.3

6.9

6.4

6.0

Endr.

hours

3.9

4.4

4.9

5.3

5.7

6.1

4.2

4.7

5.1

5.5

5.9

6.3

4.5

4.9

5.3

5.7

6.1

5.0

5.2

5.5

5.9

6.3

Range

miles

525

570

600

620

630

625

565

600

625

635

635

630

600

625

645

645

640

640

650

655

650

635

Endr.

hours

4.9

5.6

6.2

6.7

7.2

7.7

5.3

5.9

6.4

6.9

7.4

7.9

5.7

6.2

6.7

7.2

7.7

6.3

6.5

7.0

7.5

8.0

Range

miles

660

720

760

780

795

790

710

760

790

805

805

795

755

790

810

820

810

810

820

830

825

800

%

BHP

86

79

72

65

58

52

82

75

68

61

55

49

78

71

64

58

52

70

67

61

55

49

TAS

MPH

134

129

123

117

111

103

134

128

122

116

108

100

133

127

121

113

105

129

125

118

110

100

38 gal (no reserve) 48 gal RPM (no reserve)

2,700

2,600

2,500

2,400

2,300

2,200

2,700

2,600

2,500

2,400

2,300

2,200

2,700

2,600

2,500

2,400

2,300

2,650

2,600

2,500

2,400

2,300

Gross weight—2,300 lb.

Standard conditions

Zero wind 

Lean mixture

Maximum cruise is normally limited to 75% power.

Figure 10-27. Cruise power setting.

CRUISE POWER SETTING

65% MAXIMUM CONTINUOUS POWER (OR FULL THROTTLE)

2,800 POUNDS

Press

ALT

27

19

12

5

-2

-8

-15

-22

-29

-3

-7

-11

-15

-19

-22

-26

-30

-34

°F °C RPM " HG PSI GPH kts MPH

IOAT TAS Engine

speed

Man.

press

Fuel

flow per

engine

S.L.

2,000

4,000

6,000

8,000

10,000

12,000

14,000

16,000

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

20.7

20.4

20.1

19.8

19.5

19.2

18.8

17.4

16.1

6.6

6.6

6.6

6.6

6.6

6.6

6.4

5.8

5.3

11.5

11.5

11.5

11.5

11.5

11.5

11.3

10.5

9.7

147

149

152

155

157

160

162

159

156

169

171

175

178

181

184

186

183

180

°F °C RPM " HG PSI GPH kts MPH

IOAT TAS Engine

speed

Man.

press

Fuel

flow per

engine

°F °C RPM PSI GPH kts MPH

"

" HG

IOAT TAS Engine

speed

Man.

press

Fuel

flow per

engine

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

2,450

6.6

6.6

6.6

6.6

6.6

6.6

6.1

5.6

5.1

6.6

6.6

6.6

6.6

6.6

6.5

5.9

5.4

4.9

11.5

11.5

11.5

11.5

11.5

11.5

10.9

10.1

9.4

11.5

11.5

11.5

11.5

11.5

11.4

10.6

9.8

9.1

63

55

48

41

36

28

21

14

7

17

13

9

5

2

-2

-6

-10

-14

21.2

21.0

20.7

20.4

20.2

19.9

18.8

17.4

16.1

150

153

156

158

161

163

163

160

156

173

176

180

182

185

188

188

184

180

99

91

84

79

72

64

57

50

43

37

33

29

26

22

18

14

10

6

21.8

21.5

21.3

21.0

20.8

20.3

18.8

17.4

16.1

153

156

159

161

164

166

163

160

155

176

180

183

185

189

191

188

184

178

ISA -20° (-36 °F) Standard day (ISA) ISA +20° (+36 °F)

1. Full throttle manifold pressure settings are approximate.

2. Shaded area represents operation with full throttle.

Notes

read that the endurance in hours is 5.5 hours and the range 

in miles is 635 miles. 

Cruise power setting tables are useful when planning crosscountry flights. The table gives the correct cruise power 

settings, as well as the fuel flow and airspeed performance 

numbers at that altitude and airspeed.

Sample Problem 7

Pressure Altitude at Cruise................................6,000 feet 

OAT..................................................36 °F above standard 

Refer to Figure 10-27 for this sample problem. First, locate 

the pressure altitude of 6,000 feet on the far left side of the 

table. Follow that line across to the far right side of the table 

under the 20 °C (or 36 °F) column. At 6,000 feet, the rpm 

setting of 2,450 will maintain 65 percent continuous power 

at 21.0 "Hg with a fuel flow rate of 11.5 gallons per hour and 

airspeed of 161 knots.

Another type of cruise chart is a best power mixture range 

graph. This graph gives the best range based on power 

setting and altitude. Using Figure 10-28, find the range at 

65 percent power with and without a reserve based on the 

provided conditions.

Sample Problem 8

OAT....................................................................Standard

Pressure Altitude...............................................5,000 feet

First, move up the left side of the graph to 5,000 feet and 

standard temperature. Follow the line straight across the 

graph until it intersects the 65 percent line under both the 

reserve and no reserve categories. Draw a line straight down 

from both intersections to the bottom of the graph. At 65 

percent power with a reserve, the range is approximately 

522 miles. At 65 percent power with no reserve, the range 

should be 581 miles.

The last cruise chart referenced is a cruise performance graph. 

This graph is designed to tell the TAS performance of the 

airplane depending on the altitude, temperature, and power 
10-24

Figure 10-28. Best power mixture range.

Best Power Mixture Range

Range (nautical miles)

(Includes distance to climb and descend)

14

12

10

8

6

4

2

S.L.

-13°

-9°

-5°

-1°

3°

7°

11°

15°

450 500 550 600 500 550 600 650

Pressure ALT (1,000 feet)

Standard Temperature °C

Associated conditions

Mixture Leaned per section 4

Weight 2,300 lb.

Wings No 

Fuel 48 gal usable

Wheel Fairings installed

Cruise Mid cruise

Power 75%

Power

65%

Power 55%

Power 75%

Power 65%

Power 55%

Notes

Range may be reduced

by up to 7% if wheel

fairings are not installed

Notes

Add 0.6 NM for each

degree Celsius above

standard temperature

and subtract 1 NM for

each degree Celsius

below standard

temperature.

45 minutes reserve at 55%

power best economy mixture No reserve

Figure 10-29. Cruise performance graph.

20,000

18,000

16,000

14,000

12,000

10,000

8,000

6,000

4,000

2,000

Sea level

-40° -30° -20° -10° 0° 10° 20° 30° 40° 100 120 140 160 180 200

 Outside air temperature (°C) True airspeed (knots)

Pressure ALT (feet)

Associated conditions

Weight 3,600 lb. gross weight

Flaps Up

Best power Mixture leaned to 100° 

rich of peak EGT

Best economy Mixture leaned to peak EGT

1,650° Max allowable EGT

Wheel Fairings installed

Power 55%Power 65%

Power 75%

Notes

Subtract 8 knots if wheel

fairings are not installed.

Best power

Best economy

Standard temperature

 2,575 rpm at 36 IN. M.P. – 2-blade prop

 2,700 rpm at 36 IN. M.P. – 3-blade prop

setting. Using Figure 10-29, find the TAS performance based 

on the given information.

Sample Problem 9

OAT.........................................................................16 °C

Pressure Altitude...............................................6,000 feet

Power Setting................................65 percent, best power

Wheel Fairings..............................................Not installed

Begin by finding the correct OAT on the bottom, left side of 

the graph. Move up that line until it intersects the pressure 

altitude of 6,000 feet. Draw a line straight across to the 

65 percent, best power line. This is the solid line, which 

represents best economy. Draw a line straight down from 

this intersection to the bottom of the graph. The TAS at 65 

percent best power is 140 knots. However, it is necessary 

to subtract 8 knots from the speed since there are no wheel 

fairings. This note is listed under the title and conditions. 

The TAS will be 132 knots.

Crosswind and Headwind Component Chart

Every aircraft is tested according to Federal Aviation 

Administration (FAA) regulations prior to certification. The 

aircraft is tested by a pilot with average piloting skills in 

90° crosswinds with a velocity up to 0.2 VSO or two-tenths 

of the aircraft’s stalling speed with power off, gear down, 

and flaps down. This means that if the stalling speed of the 

aircraft is 45 knots, it must be capable of landing in a 9-knot, 

90° crosswind. The maximum demonstrated crosswind 

component is published in the AFM/POH. The crosswind and 

headwind component chart allows for figuring the headwind 

and crosswind component for any given wind direction and 

velocity.

Sample Problem 10

Runway..........................................................................17

Wind........................................................140° at 25 knots

Refer to Figure 10-30 to solve this problem. First, determine 

how many degrees difference there is between the runway 

and the wind direction. It is known that runway 17 means 

a direction of 170°; from that subtract the wind direction of 

140°. This gives a 30° angular difference, or wind angle. Next, 

locate the 30° mark and draw a line from there until it intersects 
10-25

Figure 10-30. Crosswind component chart.

Crosswind component

Headwind component

70

60

50

40

30

20

10

10 20 30 40 50 60 70 0

Wind velocityWind veloccity

10° 0°

20°

30°

40°

50°

60°

70°

80°

90°

Figure 10-31. Landing distance table. Note

LANDING DISTANCE

Conditions

Gross

weight

lb Ground roll Total to clear

50 ft OBS

Approach speed

IAS, MPH

At sea level & 59 °F

Flaps lowered to 40°

Power off

Hard surface runway

Zero wind

1,600 60 445 1,075 470 1,135 495 1,195 520 1,255 

Ground roll Total to clear

50 ft OBS

At 2,500 ft & 59 °F

Ground roll Total to clear

50 ft OBS

At 5,000 ft & 41 °F

Ground roll Total to clear

50 ft OBS

At 7,500 ft & 32 °F

1. Decrease the distances shown by 10% for each 4 knots of headwind. 2. Increase the distance by 10% for each 60 °F temperature increase above standard. 3. For operation on a dry, grass runway, increase distances (both “ground roll” and “total to clear 50 ft obstacle”) by 20% of the “total to clear 50 ft obstacle” figure. 

the correct wind velocity of 25 knots. From there, draw a 

line straight down and a line straight across. The headwind 

component is 22 knots and the crosswind component is 13 

knots. This information is important when taking off and 

landing so that, first of all, the appropriate runway can be 

picked if more than one exists at a particular airport, but also 

so that the aircraft is not pushed beyond its tested limits.

Landing Charts

Landing performance is affected by variables similar to those 

affecting takeoff performance. It is necessary to compensate 

for differences in density altitude, weight of the airplane, and 

headwinds. Like takeoff performance charts, landing distance 

information is available as normal landing information, as 

well as landing distance over a 50 foot obstacle. As usual, read 

the associated conditions and notes in order to ascertain the 

basis of the chart information. Remember, when calculating 

landing distance that the landing weight will not be the same 

as the takeoff weight. The weight must be recalculated to 

compensate for the fuel that was used during the flight.

Sample Problem 11

Pressure Altitude...............................................1,250 feet

Temperature.........................................................Standard

Refer to Figure 10-31. This example makes use of a landing 

distance table. Notice that the altitude of 1,250 feet is not 

on this table. It is, therefore, necessary to interpolate to find 

the correct landing distance. The pressure altitude of 1,250 

is halfway between sea level and 2,500 feet. First, find the 

column for sea level and the column for 2,500 feet. Take the 

total distance of 1,075 for sea level and the total distance of 

1,135 for 2,500 and add them together. Divide the total by 

two to obtain the distance for 1,250 feet. The distance is 1,105 

feet total landing distance to clear a 50 foot obstacle. Repeat 

this process to obtain the ground roll distance for the pressure 

altitude. The ground roll should be 457.5 feet.

Sample Problem 12

OAT.......................................................................... 57 °F

Pressure Altitude.............................................. 4,000 feet

Landing Weight...........................................2,400 pounds

Headwind.............................................................. 6 knots

Obstacle Height..................................................... 50 feet

Using the given conditions and Figure 10-32, determine the 

landing distance for the aircraft. This graph is an example of 

a combined landing distance graph and allows compensation 

for temperature, weight, headwinds, tailwinds, and varying 

obstacle height. Begin by finding the correct OAT on the 

scale on the left side of the chart. Move up in a straight 

line to the correct pressure altitude of 4,000 feet. From this 

intersection, move straight across to the first dark reference 

line. Follow the lines in the same diagonal fashion until the 

correct landing weight is reached. At 2,400 pounds, continue 

in a straight line across to the second dark reference line. 

Once again, draw a line in a diagonal manner to the correct 

wind component and then straight across to the third dark 
10-26

Figure 10-32. Landing distance graph.

 C -40° -30° -20° -10° 0° 10° 20° 30° 40° 50° 2,800 2,600 2,400 2,200 0 10 20 30 0 50

 Outside air temperature Weight Wind component Obstacle

 (pounds) (knots) height (feet)

 Outside air temperature Weight 

 (pounds) 

 Wind component 

 (knots) 

 Obstacle

height (feet) F -40° -20° 0° 20° 40° 60° 80° 100° 120°

3,500

3,000

2,500

2,000

1,500

1,000

500

 Guide lines not

applicable for Intermediate

Tailwind

Headwind

Reference lineReference lineReference line

Pressure altitude (feet)

Obstacle heights

Weight

(pounds) kts MPH

Speed

at 50 feet

2,950

2,800

2,600

2,400

2,200

70

68

65

63

60

80

78

75

72

69

Associated conditions

Power Retarded to maintain

900 feet/on final approach

Flaps Down

Landing gear Down

Runway Paved, level, dry surface

Approach speed IAS as tabulated

Braking Maximum

10,000

8,000

6,000

2,00

4,000

0

S.L.

ISA

Figure 10-33. Stall speed table.

Stall Speed

Gross weight

2,750 lb

MPH 62 67 74 88

knots 54 58 64 76

MPH 75 81 89 106

knots 65 70 77 92

MPH 54 58 64 76 

knots 47 50 56 66

MPH 66 71 78 93

knots 57 62 68 81

Angle of bank

Power Power

Gear and flaps up

Gear and flaps down

Level 30° 45° 60°

On

Off

On

Off

reference line. From this point, draw a line in two separate 

directions: one straight across to figure the ground roll and 

one in a diagonal manner to the correct obstacle height. This 

should be 900 feet for the total ground roll and 1,300 feet for 

the total distance over a 50 foot obstacle.

Stall Speed Performance Charts

Stall speed performance charts are designed to give an 

understanding of the speed at which the aircraft will stall 

in a given configuration. This type of chart will typically 

take into account the angle of bank, the position of the gear 

and flaps, and the throttle position. Use Figure 10-33 and 

the accompanying conditions to find the speed at which the 

airplane will stall.

Sample Problem 13

Power........................................................................ OFF

Flaps....................................................................... Down 

Gear........................................................................ Down

Angle of Bank............................................................. 45° 

First, locate the correct flap and gear configuration. The 

bottom half of the chart should be used since the gear and 

flaps are down. Next, choose the row corresponding to a 

power-off situation. Now, find the correct angle of bank 

column, which is 45°. The stall speed is 78 mph, and the 

stall speed in knots would be 68 knots.

Performance charts provide valuable information to the 

pilot. Take advantage of these charts. A pilot can predict the 

performance of the aircraft under most flying conditions, and 

this enables a better plan for every flight. The Code of Federal 

Regulations (CFR) requires that a pilot be familiar with all 

information available prior to any flight. Pilots should use 

the information to their advantage as it can only contribute 

to safety in flight.

Transport Category Airplane 

Performance

Transport category aircraft are certificated under Title 14 

of the CFR (14 CFR) parts 25 and 29. The airworthiness 

certification standards of part 25 and 29 require proven 

levels of performance and guarantee safety margins for these 

aircraft, regardless of the specific operating regulations under 

which they are employed.
10-27

Major Differences in Transport Category 

Versus Non-Transport Category Performance 

Requirements

• Full temperature accountability—all of the performance 

charts for the transport category aircraft require that 

takeoff and climb performance be computed with the 

full effects of temperature considered. 

• Climb performance expressed as percent gradient 

of climb—the transport category aircraft’s climb 

performance is expressed as a percent gradient of 

climb rather than a figure calculated in fpm of climb. 

This percent gradient of climb is a much more practical 

expression of performance since it is the aircraft’s 

angle of climb that is critical in an obstacle clearance 

situation.

• Change in lift-off technique—lift-off technique in 

transport category aircraft allows the reaching of V2

(takeoff safety speed) after the aircraft is airborne. 

This is possible because of the excellent acceleration 

and reliability characteristics of the engines on these 

aircraft and due to the larger surplus of power.

• Performance requirements applicable to all segments 

of aviation—all aircraft certificated by the FAA in the 

transport category, whatever the size, must be operated 

in accordance with the same performance criteria. 

This applies to both commercial and non-commercial 

operations.

Performance Requirements

The performance requirements that the transport category 

aircraft must meet are:

Takeoff

• Takeoff speeds

• Takeoff runway required

• Takeoff climb required

• Obstacle clearance requirements

Landing

• Landing speeds

• Landing runway required

• Landing climb required

Takeoff Planning

Listed below are the speeds that affect the transport category 

aircraft’s takeoff performance. The flight crew must be 

thoroughly familiar with each of these speeds and how they 

are used in takeoff planning. 

• VS—stalling speed or the minimum steady flight speed 

at which the aircraft is controllable. 

• VMCG—minimum control speed on the ground, with 

one engine inoperative, (critical engine on two-engine 

airplanes) takeoff power on other engine(s), using 

aerodynamic controls only for directional control 

(must be less than V1). 

• VMCA—minimum control speed in the air, with one 

engine inoperative, (critical engine on two-engine 

aircraft) operating engine(s) at takeoff power, 

maximum of 5° bank into the good engine(s). 

• V1—critical engine failure speed or decision speed. 

Engine failure below this speed shall result in an 

aborted takeoff; above this speed the takeoff run 

should be continued.

• VR—speed at which the rotation of the aircraft is 

initiated to takeoff attitude. The speed cannot be less 

than V1 or less than 1.05 times VMC. With an engine 

failure, it must also allow for the acceleration to V2

at the 35-foot height at the end of the runway.

• VLOF—lift-off speed. The speed at which the aircraft 

first becomes airborne.

• V2—the takeoff safety speed which must be attained 

at the 35-foot height at the end of the required runway 

distance. This is essentially the best one-engine 

operative angle of climb speed for the aircraft and 

should be held until clearing obstacles after takeoff, 

or until at least 400 feet above the ground.

• VFS—final segment climb speed, which is based upon 

one-engine inooerative climb, clean configuration, and 

mximum continuos power setting.

All of the V speeds should be considered during every 

takeoff. The V1, VR, V2, and VFS speeds should be visibly 

posted in the flightdeck for reference during the takeoff.

Takeoff speeds vary with aircraft weight. Before takeoff 

speeds can be computed, the pilot must first determine the 

maximum allowable takeoff weight. The three items that can 

limit takeoff weight are runway requirements, takeoff climb 

requirements, and obstacle clearance requirements.

Runway Requirements

The runway requirements for takeoff are affected by:

• Pressure altitude

• Temperature

• Headwind component

• Runway gradient or slope

• Aircraft weight
10-28

Figure 10-34. Minimum required takeoff. 25

V1 VR VLO V2

All engine acceleration One engine acceleration

35'

25

V1 VR VLO

All engine takeoff distance to 35' 15% margin

All engine takeoff field length

35'

25

V1

All engine acceleration Stop distance

Takeoff distance with

an engine failure

Accelerate-stop

distance

115% All engine takeoff

distance

The runway required for takeoff must be based upon the 

possible loss of an engine at the most critical point, which is 

at V1 (decision speed). By regulation, the aircraft’s takeoff 

weight has to accommodate the longest of three distances:

1. Accelerate-go distance—the distance required to 

accelerate to V1 with all engines at takeoff power, 

experience an engine failure at V1 and continue the 

takeoff on the remaining engine(s). The runway 

required includes the distance required to climb to 35 

feet by which time V2 speed must be attained.

2. Accelerate-stop distance—the distance required to 

accelerate to V1 with all engines at takeoff power, 

experience an engine failure at V1, and abort the takeoff 

and bring the aircraft to a stop using braking action 

only (use of thrust reversing is not considered).

3. Takeoff distance—the distance required to complete 

an all-engines operative takeoff to the 35-foot height. 

It must be at least 15 percent less than the distance 

required for a one-engine inoperative engine takeoff. 

This distance is not normally a limiting factor as it is 

usually less than the one-engine inoperative takeoff 

distance. 

These three required takeoff runway considerations are 

shown in Figure 10-34.

Balanced Field Length

In most cases, the pilot will be working with a performance 

chart for takeoff runway required, which will give “balanced 

field length” information. This means that the distance 
10-29

Figure 10-35. Normal takeoff runway required.

TAKEOFF RUNWAY REQUIREMENTS

Standard ISA conditions

Notes Conditions

Takeoff

gross weight

at brake release Sea level (V1)

47 (121)

48 (121)

53 (122)

58 (123)

43 (121)

43 (121)

48 (122)

53 (123)

45 (118)

46 (118)

50 (118)

55 (120)

40 (118)

42 (118)

45 (118)

50 (120)

40 (114)

41 (115)

45 (114)

50 (115)

36 (114)

37 (115)

41 (114)

46 (115)

36 (108)

37 (110)

40 (108)

45 (111)

32 (108)

34 (110)

36 (108)

41 (111)

32 (104)

34 (105)

36 (104)

41 (106)

29 (104)

31 (105)

32 (104)

37 (106)

 28 (98)

30 (100)

 32 (99)

36 (101)

 25 (98)

 27 (100)

 29 (99)

 32 (101)

 1,000 (V1)

48 (121)

51 (121)

56 (122)

62 (124)

43 (121)

46 (121)

51 (122)

57 (124)

45 (118)

48 (118)

53 (119)

59 (121)

41 (118)

44 (118)

48 (119)

54 (121)

41 (114)

43 (114)

48 (115)

53 (116)

37 (114)

39 (114)

44 (115)

48 (116)

37 (108)

39 (108)

43 (110)

46 (112)

33 (108)

35 (108)

39 (110)

44 (112)

33 (103)

35 (103)

38 (105)

43 (107)

30 (103)

32 (103)

34 (105)

39 (107)

 30 (98)

 31 (98)

 34 (100)

 38 (102)

 27 (98)

 29 (98)

 31 (100)

 34 (102)

 2,000 (V1)

50 (120)

55 (121)

60 (123)

68 (124)

45 (120)

50 (122)

55 (123)

62 (124)

47 (117)

51 (118)

57 (120)

64 (121)

43 (117)

46 (118)

52 (120)

58 (121)

42 (113)

46 (114)

51 (115)

58 (116)

38 (113)

42 (114)

46 (115)

53 (116)

38 (107)

41 (109)

46 (111)

52 (112)

34 (107)

37 (109)

42 (111)

47 (112)

34 (103)

37 (104)

41 (105)

46 (107)

31 (103)

33 (104)

37 (105)

42 (107)

 30 (98)

 33 (99)

 37 (101)

 41 (102)

 27 (98)

 30 (99)

 33 (101)

 37 (102)

 3,000 (V1)

53 (121)

60 (122)

65 (124)

73 (125)

48 (121)

55 (122)

59 (124)

66 (125)

50 (118)

56 (119)

66 (121)

73 (122)

45 (118)

51 (119)

56 (121)

63 (122)

45 (113)

50 (115)

56 (116)

62 (117)

41 (113)

46 (115)

51 (116)

56 (117)

40 (108)

45 (110)

50 (111)

56 (113)

36 (108)

41 (110)

45 (111)

51 (113)

36 (103)

41 (105)

45 (106)

50 (108)

32 (103)

37 (105)

41 (106)

45 (108)

 32 (98)

36 (100)

40 (102)

44 (103)

 29 (98)

32 (100)

36 (102)

40 (103)

 4,000 (V1)

57 (122)

63 (123)

70 (125)

78 (126)

52 (122)

57 (123)

63 (125)

71 (126)

54 (119)

59 (120)

66 (121)

73 (123)

49 (119)

54 (120)

60 (121)

66 (123)

49 (114)

53 (115)

59 (116)

66 (118)

45 (114)

48 (115)

56 (116)

60 (118)

44 (109)

48 (110)

53 (112)

59 (114)

40 (109)

44 (110)

48 (112)

54 (114)

39 (105)

43 (106)

48 (107)

53 (108)

35 (105)

39 (106)

44 (107)

48 (108)

 35 (99)

38 (101)

42 (102)

47 (104)

 32 (99)

34 (101)

38 (102)

43 (104)

 5,000 (V1)

62 (123)

69 (124)

77 (125)

85 (127)

56 (123)

63 (124)

70 (125)

77 (127)

59 (120)

65 (121)

72 (122)

80 (124)

54 (120)

59 (121)

65 (122)

73 (124)

53 (115)

59 (116)

65 (116)

73 (118)

48 (115)

54 (116)

59 (116)

66 (118)

48 (110)

53 (111)

58 (111)

65 (114)

44 (110)

48 (111)

53 (111)

59 (114)

43 (106)

47 (107)

52 (107)

58 (109)

39 (106)

43 (107)

47 (107)

53 (109)

38 (101)

42 (102)

46 (102)

51 (104)

34 (101)

38 (102)

42 (102)

46 (104)

 6,000 (V1)

70 (123)

77 (125)

85 (126)

95 (129)

64 (123)

70 (125)

77 (126)

85 (129)

66 (120)

73 (121)

80 (123)

90 (124)

60 (120)

66 (121)

72 (123)

81 (124)

60 (115)

66 (117)

72 (117)

80 (119)

54 (115)

60 (117)

65 (117)

73 (119)

53 (111)

59 (112)

65 (113)

72 (114)

48 (111)

54 (112)

59 (113)

65 (114)

48 (106)

53 (107)

58 (108)

64 (110)

44 (106)

48 (107)

53 (108)

58 (110)

42 (101)

46 (102)

51 (103)

56 (105)

38 (101)

42 (102)

46 (103)

51 (105)

Headwind

(knots) °F

30

50

70

90

30

50

70

90

30

50

70

90

30

50

70

90

30

50

70

90

30

50

70

90

30

50

70

90

30

50

70

90

30

50

70

90

30

50

70

90

30

50

70

90

30

50

70

90

°C

-1.1

10

21

32

-1.1

10

21

32

-1.1

10

21

32

-1.1

10

21

32

-1.1

10

21

32

-1.1

10

21

32

-1.1

10

21

32

-1.1

10

21

32

-1.1

10

21

32

-1.1

10

21

32

-1.1

10

21

32

-1.1

10

21

32

Temp. Pressure altitude (feet)

0

20

0

20

0

20

0

20

0

20

0

20

19,612

VR = 126

V2 = 134

19,000

VR = 124

V2 = 131

18,000

VR = 119

V2 = 127

17,000

VR = 115

V2 = 124

16,000

VR = 111

V2 = 120

15,000

VR = 106

V2 = 116

Cabin pressurization On

Zero slope runway

No flaps—Anti-ice RAM air inlets Off

Anti-skid On

Distances—100 feet (V1 – KIAS)

 Shaded area indicates conditions that do not meet second segment climb requirements.

Refer to F.M. for takeoff limitations.

shown for the takeoff will include both the accelerate-go and 

accelerate-stop distances. One effective means of presenting 

the normal takeoff data is shown in the tabulated chart in 

Figure 10-35.

The chart in Figure 10-35 shows the runway distance required 

under normal conditions and is useful as a quick reference 

chart for the standard takeoff. The V speeds for the various 

weights and conditions are also shown. 

For other than normal takeoff conditions, such as with 

engine anti-ice, anti-skid brakes inoperative, or extremes in 

temperature or runway slope, the pilot should consult the 

appropriate takeoff performance charts in the performance 

section of the AFM.

There are other occasions of very high weight and 

temperature where the runway requirement may be dictated 

by the maximum brake kinetic energy limits that affect 
10-30

Figure 10-36. One engine inoperative takeoff.

Ground roll 1st segment—climb 2nd segment—climb 3rd segment— 

acceleration

4th 

segment—

climb

Takeoff

Down

Variable

Down Retraction

All operating One inoperative

Retracted

V2 Variable

Retracted

M.C.

VFS or 1.25 Vs min.

Power

Flaps

Engine

Airspeed

Landing gear

400'

1,500'

35'

*

M.C. = Maximum continuous

V1 = Critical-engine-failure speed

V2 = Takeoff safety speed

VS = Calibrated stalling speed, or 

 minimum steady flight speed at 

 which the aircraft is controllable

VR = Speed at which aircraft can 

 start safely raising nose wheel 

 off surface (Rotational Speed)

VLOF = Speed at point where airplane

 lifts off

VR VLOF

V2

V1

2 Engine Positive 2.4% Positive 1.2%

3 Engine 3.0% 2.7% Positive 1.5%

4 Engine 5.0% 3.0% Positive 1.7%

Wing flaps T.O. T.O. T.O. Up

Landing gear Down Up Up Up

Engines 1 Out 1 Out 1 Out 1 Out

Power T.O. T.O. T.O. M.C.

Air speed VLOF V2 V2 V2 1.25 VS(Min) 1.25 VS(Min)

* Required Absolute Minimum Gradient of Flightpath

1st T/O

segment

2nd T/O

segment

Final T/O

segment

Transition (acceleration) Items

the aircraft’s ability to stop. Under these conditions, the 

accelerate-stop distance may be greater than the acceleratego. The procedure to bring performance back to a balanced 

field takeoff condition is to limit the V1 speed so that it 

does not exceed the maximum brake kinetic energy speed 

(sometimes called VBE). This procedure also results in a 

reduction in allowable takeoff weight. 

Climb Requirements

After the aircraft has reached the 35 foot height with one engine 

inoperative, there is a requirement that it be able to climb 

at a specified climb gradient. This is known as the takeoff 

flightpath requirement. The aircraft’s performance must be 

considered based upon a one-engine inoperative climb up to 

1,500 feet above the ground. The takeoff flightpath profile 

with required gradients of climb for the various segments 

and configurations is shown in Figure 10-36.

NOTE: Climb gradient can best be described as being a 

specific gain of vertical height for a given distance covered 

horizontally. For instance, a 2.4 percent gradient means that 

24 feet of altitude would be gained for each 1,000 feet of 

distance covered horizontally across the ground.

The following brief explanation of the one-engine inoperative 

climb profile may be helpful in understanding the chart in 

Figure 10-36.

First Segment

This segment is included in the takeoff runway required 

charts, and is measured from the point at which the aircraft 

becomes airborne until it reaches the 35-foot height at the 

end of the runway distance required. Speed initially is VLO

and must be V2 at the 35 foot height.
10-31

Figure 10-37. Takeoff obstacle clearance.

35'

35'

35'

400'

Obstacle clearance pat

Net tak

h

eoff flightpath

Minimum actual takeoff flightpath

Second Segment

This is the most critical segment of the profile. The second 

segment is the climb from the 35 foot height to 400 feet 

above the ground. The climb is done at full takeoff power 

on the operating engine(s), at V2 speed, and with the flaps 

in the takeoff configuration. The required climb gradient 

in this segment is 2.4 percent for two-engine aircraft, 2.7 

percent for three-engine aircraft, and 3.0 percent for fourengine aircraft.

Third or Acceleration Segment

During this segment, the airplane is considered to be 

maintaining the 400 feet above the ground and accelerating 

from the V2 speed to the VFS speed before the climb profile 

is continued. The flaps are raised at the beginning of the 

acceleration segment and power is maintained at the takeoff 

setting as long as possible (5 minutes maximum).

Fourth or Final Segment

This segment is from the 400 to 1,500 foot AGL altitude 

with power set at maximum continuous. The required climb 

in this segment is a gradient of 1.2 percent for two-engine 

airplanes, 1.55 for three-engine airplanes, and 1.7 percent 

for four-engine airplanes.

Second Segment Climb Limitations

The second segment climb requirements, from 35 to 400 

feet, are the most restrictive (or hardest to meet) of the 

climb segments. The pilot must determine that the second 

segment climb is met for each takeoff. In order to achieve this 

performance at the higher density altitude conditions, it may 

be necessary to limit the takeoff weight of the aircraft. 

It must be realized that, regardless of the actual available 

length of the takeoff runway, takeoff weight must be 

adjusted so that the second segment climb requirements can 

be met. The aircraft may well be capable of lifting off with 

one engine inoperative, but it must then be able to climb 

and clear obstacles. Although second segment climb may 

not present much of a problem at the lower altitudes, at the 

higher altitude airports and higher temperatures, the second 

segment climb chart should be consulted to determine the 

effects on maximum takeoff weights before figuring takeoff 

runway distance required.

Air Carrier Obstacle Clearance 

Requirements

Regulations require that large transport category turbine 

powered aircraft certificated after September 30, 1958, be 

taken off at a weight that allows a net takeoff flightpath 

(one engine inoperative) that clears all obstacles either by 

a height of at least 35 feet vertically, or by at least 200 feet 

horizontally within the airport boundaries and by at least 300 

feet horizontally after passing the boundaries. The takeoff 

flightpath is considered to begin 35 feet above the takeoff 

surface at the end of the takeoff distance, and extends to a 

point in the takeoff at which the aircraft is 1,500 feet above 

the takeoff surface, or at which the transition from the takeoff 

to the en route configuration is completed. The net takeoff 

flightpath is the actual takeoff flightpath reduced at each point 

by 0.8 percent for two engine aircraft, 0.9 percent for threeengine aircraft, and 1.0 percent for four-engine aircraft.

Air carrier pilots therefore are responsible not only for 

determining that there is enough runway available for an 

engine inoperative takeoff (balanced field length), and 

the ability to meet required climb gradients; but they must 

also assure that the aircraft will safely be able to clear any 

obstacles that may be in the takeoff flightpath. The net 

takeoff flightpath and obstacle clearance required are shown 

in Figure 10-37.
10-32

The usual method of computing net takeoff flightpath 

performance is to add up the total ground distances required 

for each of the climb segments and/or use obstacle clearance 

performance charts in the AFM. Although this obstacle 

clearance requirement is seldom a limitation at the normally 

used airports, it is quite often an important consideration 

under critical conditions such as high takeoff weight and/or 

high density altitude. Consider that at a 2.4 percent climb 

gradient (2.4 feet up for every 100 feet forward) a 1,500 

foot altitude gain would take a horizontal distance of 10.4 

NM to achieve.

Summary of Takeoff Requirements

In order to establish the allowable takeoff weight for a 

transport category aircraft, at any airfield, the following must 

be considered:

• Airfield pressure altitude

• Temperature

• Headwind component

• Runway length

• Runway gradient or slope

• Obstacles in the flightpath

Once the above details are known and applied to the 

appropriate performance charts, it is possible to determine 

the maximum allowable takeoff weight. This weight would 

be the lower of the maximum weights as allowed by:

• Balanced field length required

• Engine inoperative climb ability (second segment 

limited)

• Obstacle clearance requirement

In practice, restrictions to takeoff weight at low altitude 

airports are usually due to runway length limitations; engine 

inoperative climb limitations are most common at the higher 

altitude airports. All limitations to weight must be observed. 

Since the combined weight of fuel and payload in the aircraft 

may amount to nearly half the maximum takeoff weight, 

it is usually possible to reduce fuel weight to meet takeoff 

limitations. If this is done, however, flight planning must be 

recalculated in light of reduced fuel and range.

Landing Performance

As in the takeoff planning, certain speeds must be considered 

during landing. These speeds are shown below.

• VSO—stalling speed or the minimum steady flight 

speed in the landing configuration. 

• VREF—1.3 times the stalling speed in the landing 

configuration. This is the required speed at the 50-foot 

height above the threshold end of the runway. 

• Approach climb—the speed which gives the best climb 

performance in the approach confguration, with one 

engine inoperative, and with maximum takeoff power 

on the operating engine(s). The required gradient of 

climb in this configuration is 2.1 percent for twoengine aircraft, 2.4 percent for three-emgine aircraft, 

and 2.7 percent for four-engine aircraft.

• Landing climb—the speed giving the best performance 

in the full landing configuration with maximum 

takeoff power on all engines. The gradient of climb 

required in this configuration is 3.2 percent.

Planning the Landing

As in the takeoff, the landing speeds shown above should be 

precomputed and visible to both pilots prior to the landing. 

The VREF speed, or threshold speed, is used as a reference 

speed throughout the traffic pattern or instrument approach 

as in the following example:

VREF plus 30K D o w n w i n d o r p r o c e d u r e t u r n 

VREF plus 20K Base leg or final course inbound to final 

 fix 

VREF plus 10K Final or final course inbound from fix 

 (ILS final)

VREF Speed at the 50 foot height above the 

 threshold

Landing Requirements

The maximum landing weight of an aircraft can be restricted 

by either the approach climb requirements or by the landing 

runway available.

Approach Climb Requirements

The approach climb is usually more limiting (or more difficult 

to meet) than the landing climb, primarily because it is based 

upon the ability to execute a missed approach with one engine 

inoperative. The required climb gradient can be affected 

by pressure altitude and temperature and, as in the second 

segment climb in the takeoff, aircraft weight must be limited 

as needed in order to comply with this climb requirement.
10-33

Figure 10-38. Landing runway requirements. 25

Required landing distance per 14 CFR part 91 40% of runway length

Required landing runway length per 14 CFR part 121

50' Complete stop

VSO Approach

1.3 VSO

Landing Runway Required

The runway distance needed for landing can be affected by 

the following:

• Pressure altitude

• Temperature

• Headwind component

• Runway gradient or slope

• Aircraft weight 

In computing the landing distance required, some manufacturers 

do not include all of the above items in their charts, since the 

regulations state that only pressure altitude, wind, and aircraft 

weight must be considered. Charts are provided for anti-skid 

on and anti-skid off conditions, but the use of reverse thrust is 

not used in computing required landing distances.

The landing distance, as required by the regulations, is 

that distance needed to land and come to a complete stop 

from a point 50 feet above the threshold end of the runway. 

It includes the air distance required to travel from the 50 

foot height to touchdown (which can consume 1,000 feet 

of runway distance), plus the stopping distance, with no 

margin left over. This is all that is required for 14 CFR part 

91 operators (non-air carrier), and all that is shown on some 

landing distance required charts.

For air carriers and other commercial operators subject to 

14 CFR part 121, a different set of rules applies stating that 

the required landing distance from the 50 foot height cannot 

exceed 60 percent of the actual runway length available. 

In all cases, the minimum airspeed allowed at the 50 foot 

height must be no less than 1.3 times the aircraft’s stalling 

speed in the landing configuration. This speed is commonly 

called the aircraft’s VREF speed and varies with landing 

weight. Figure 10-38 is a diagram of these landing runway 

requirements.

Summary of Landing Requirements

In order to establish the allowable landing weight for a 

transport category aircraft, the following details must be 

considered:

• Airfield pressure altitude

• Temperature

• Headwind component

• Runway length

• Runway gradient or slope

• Runway surface condition

With these details, it is possible to establish the maximum 

allowable landing weight, which will be the lower of the 

weights as dictated by:

• Landing runway requirements

• Approach climb requirements

In practice, the approach climb limitations (ability to climb 

in approach configuration with one engine inoperative) 

are seldom encountered because the landing weights upon 

arrival at the destination airport are usually low. However, 

as in the second segment climb requirement for takeoff, this 

approach climb gradient must be met and landing weights 

must be restricted if necessary. The most likely conditions 

that would make the approach climb critical would be the 

landings at high weights and high pressure altitudes and 

temperatures, which might be encountered if a landing were 

required shortly after takeoff.

Landing field requirements can more frequently limit an 

aircraft’s allowable landing weight than the approach climb 

limitations. Again, however, unless the runway is particularly 

short, this is seldom problematical as the average landing 

weight at the destination rarely approaches the maximum 

design landing weight due to fuel burn off. 
10-34

Chapter Summary

Performance characteristics and capabilities vary greatly 

among aircraft. Moreover, aircraft weight, atmospheric 

conditions, and external environmental factors can 

significantly affect aircraft performance. It is essential that 

a pilot become intimately familiar with the performance 

characteristics and capabilities of the aircraft being flown. 

The primary source of this information is the AFM/POH.