PilotHandbook

Prior to the unit’s intended use, an operational check Special Flight Permits
must be performed in accordance with the applicable A special flight permit is a Special Airworthiness Certificate
sections of 14 CFR part 91 on checking, removing, authorizing operation of an aircraft that does not currently
and replacing magnetic chip detectors. meet applicable airworthiness requirements but is safe
for a specific flight. Before the permit is issued, an FAA
• Inspection and maintenance tasks prescribed and inspector may personally inspect the aircraft, or require it
specifically identified as preventive maintenance to be inspected by an FAA-certificated A&P mechanic or an
in a primary category aircraft type certificate or appropriately certificated repair station to determine its safety
supplemental type certificate holder’s approved special for the intended flight. The inspection shall be recorded in
inspection and preventive maintenance program when the aircraft records.
accomplished on a primary category aircraft.
The special flight permit is issued to allow the aircraft to be
• Updating self-contained, front instrument panel- flown to a base where repairs, alterations, or maintenance
mounted air traffic control (ATC) navigational can be performed; for delivering or exporting the aircraft; or
software databases (excluding those of automatic for evacuating an aircraft from an area of impending danger.
flight control systems, transponders, and microwave A special flight permit may be issued to allow the operation
frequency DME) only if no disassembly of the unit is of an overweight aircraft for flight beyond its normal range
required and pertinent instructions are provided; prior over water or land areas where adequate landing facilities
to the unit’s intended use, an operational check must or fuel is not available.
be performed in accordance with applicable sections
of 14 CFR part 91. If a special flight permit is needed, assistance and the necessary
forms may be obtained from the local FSDO or Designated
Certificated pilots, excluding student pilots, sport pilots, and Airworthiness Representative (DAR). [Figure 8-10]
recreational pilots, may perform preventive maintenance
on any aircraft that is owned or operated by them provided Airworthiness Directives (ADs)
that aircraft is not used in air carrier service or 14 CFR part
121, 129, or 135. A pilot holding a sport pilot certificate A primary safety function of the FAA is to require correction
may perform preventive maintenance on an aircraft owned of unsafe conditions found in an aircraft, aircraft engine,
or operated by that pilot if that aircraft is issued a special propeller, or appliance when such conditions exist and
airworthiness certificate in the LSA category. (Sport pilots are likely to exist or develop in other products of the same
operating LSA should refer to 14 CFR part 65 for maintenance design. The unsafe condition may exist because of a design
privileges.) 14 CFR part 43, appendix A, contains a list of the defect, maintenance, or other causes. 14 CFR part 39 and
operations that are considered to be preventive maintenance. Airworthiness Directives (ADs) define the authority and
responsibility of the Administrator for requiring the necessary
Repairs and Alterations corrective action. ADs are used to notify aircraft owners and
Repairs and alterations are classified as either major or minor. other interested persons of unsafe conditions and to specify
14 CFR part 43, appendix A, describes the alterations and the conditions under which the product may continue to be
repairs considered major. Major repairs or alterations shall operated. ADs are divided into two categories:
be approved for return to service on FAA Form 337, Major
Repair and Alteration, by an appropriately rated certificated 1. Those of an emergency nature requiring immediate
repair station, an FAA-certificated A&P mechanic holding compliance prior to further flight
an IA, or a representative of the Administrator. Minor repairs
and minor alterations may be approved for return to service 2. Those of a less urgent nature requiring compliance
with a proper entry in the maintenance records by an FAA- within a specified period of time
certificated A&P mechanic or an appropriately certificated
repair station. ADs are regulatory and shall be complied with unless a
specific exemption is granted. It is the responsibility of the
For modifications of experimental aircraft, refer to the aircraft owner or operator to ensure compliance with all
operating limitations issued to that aircraft. Modifications pertinent ADs, including those ADs that require recurrent
in accordance with FAA Order 8130.2, Airworthiness or continuing action. For example, an AD may require a
Certification of Aircraft and Related Products, may require repetitive inspection each 50 hours of operation, meaning
the notification of the issuing authority. the particular inspection shall be accomplished and recorded

8-12

Figure 8-10. FAA Form 8130-7, Special Airworthiness Certificate.

every 50 hours of time in service. Owners/operators are aircraft and helicopter books contain all ADs applicable to
reminded there is no provision to overfly the maximum hour small aircraft (12,500 pounds or less maximum certificated
requirement of an AD unless it is specifically written into the takeoff weight) and ADs applicable to all helicopters. The
AD. To help determine if an AD applies to an amateur-built large aircraft books contain all ADs applicable to large
aircraft, contact the local FSDO. aircraft.

14 CFR section 91.417 requires a record to be maintained For current information on how to order paper copies of AD
that shows the current status of applicable ADs, including books and the AD Biweekly visit the FAA online regulatory
the method of compliance; the AD number and revision date, and guidance library at: http://rgl.faa.gov.
if recurring; next due date and time; the signature; kind of
certificate; and certificate number of the repair station or Aircraft Owner/Operator Responsibilities
mechanic who performed the work. For ready reference,
many aircraft owners have a chronological listing of the The registered owner/operator of an aircraft is responsible
pertinent ADs in the back of their aircraft, engine, and for:
propeller maintenance records.
• Having a current Airworthiness Certificate and a
All ADs and the AD Biweekly are free on the Internet Certificate of Aircraft Registration in the aircraft.
at http://rgl.faa.gov. In July of 2007, the FAA made ADs
available through e-mail. Individuals can enroll for the e- • Maintaining the aircraft in an airworthy condition,
mail service at the link above. Mailing paper copies of ADs including compliance with all applicable ADs, and
will be discontinued when the e-mail system is proven to assuring that maintenance is properly recorded.
be effective.
• Keeping abreast of current regulations concerning the
Paper copies of the Summary of Airworthiness Directives and operation and maintenance of the aircraft.
the AD Biweekly may be purchased from the Superintendent
of Documents. The Summary contains all the valid ADs • Notifying the FAA Aircraft Registry immediately
previously published and is divided into two areas. The small of any change of permanent mailing address, or of
the sale or export of the aircraft, or of the loss of the
eligibility to register an aircraft. (Refer to 14 CFR
section 47.41.)

8-13

• Having a current Federal Communications Commission
(FCC) radio station license if equipped with radios,
including emergency locator transmitter (ELT), if
operated outside of the United States.

Chapter Summary

Knowledge of an aircraft’s AFM/POH and documents such as
ADs help a pilot to have ready access to pertinent information
needed to safely fly a particular aircraft. By understanding
the operations, limitations, and performance characteristics
of the aircraft, the pilot can make good flight decisions.
By learning what preventive maintenance is allowed on
the aircraft, a pilot can maintain his or her aircraft in an
airworthy condition. The goal of every pilot is a safe flight;
flight manuals and aircraft documentation are essential tools
used to reach that goal.

8-14

WeightChapter 9 and
Balance

Introduction

Compliance with the weight and balance limits of any aircraft
is critical to flight safety. Operating above the maximum
weight limitation compromises the structural integrity of
an aircraft and adversely affects its performance. Operation
with the center of gravity (CG) outside the approved limits
results in control difficulty.

Weight Control

As discussed in Chapter 4, Aerodynamics of Flight, weight
is the force with which gravity attracts a body toward the
center of the Earth. It is a product of the mass of a body and
the acceleration acting on the body. Weight is a major factor
in aircraft construction and operation, and demands respect
from all pilots.

The force of gravity continuously attempts to pull an aircraft
down toward Earth. The force of lift is the only force that
counteracts weight and sustains an aircraft in flight. The
amount of lift produced by an airfoil is limited by the airfoil
design, angle of attack (AOA), airspeed, and air density. To
assure that the lift generated is sufficient to counteract weight,
loading an aircraft beyond the manufacturer’s recommended
weight must be avoided. If the weight is greater than the lift
generated, the aircraft may be incapable of flight.

Effects of Weight
Any item aboard the aircraft that increases the total weight
is undesirable for performance. Manufacturers attempt to
make an aircraft as light as possible without sacrificing
strength or safety.

9-1

The pilot should always be aware of the consequences of Balance, Stability, and Center of Gravity
overloading. An overloaded aircraft may not be able to leave
the ground, or if it does become airborne, it may exhibit Balance refers to the location of the CG of an aircraft, and is
unexpected and unusually poor flight characteristics. If not important to stability and safety in flight. The CG is a point
properly loaded, the initial indication of poor performance at which the aircraft would balance if it were suspended at
usually takes place during takeoff. that point.

Excessive weight reduces the flight performance in almost The primary concern in balancing an aircraft is the fore and
every respect. For example, the most important performance aft location of the CG along the longitudinal axis. The CG
deficiencies of an overloaded aircraft are: is not necessarily a fixed point; its location depends on the
distribution of weight in the aircraft. As variable load items are
• Higher takeoff speed shifted or expended, there is a resultant shift in CG location.
The distance between the forward and back limits for the
• Longer takeoff run position of the center for gravity or CG range is certified for
an aircraft by the manufacturer. The pilot should realize that
• Reduced rate and angle of climb if the CG is displaced too far forward on the longitudinal
axis, a nose-heavy condition will result. Conversely, if the
• Lower maximum altitude CG is displaced too far aft on the longitudinal axis, a tail
heavy condition results. It is possible that the pilot could not
• Shorter range control the aircraft if the CG location produced an unstable
condition. [Figure 9-1]
• Reduced cruising speed
Location of the CG with reference to the lateral axis is also
• Reduced maneuverability important. For each item of weight existing to the left of the
fuselage centerline, there is an equal weight existing at a
• Higher stalling speed corresponding location on the right. This may be upset by
unbalanced lateral loading. The position of the lateral CG
• Higher approach and landing speed is not computed in all aircraft, but the pilot must be aware
that adverse effects arise as a result of a laterally unbalanced
• Longer landing roll condition. In an airplane, lateral unbalance occurs if the fuel
load is mismanaged by supplying the engine(s) unevenly from
• Excessive weight on the nose wheel or tail wheel tanks on one side of the airplane. The pilot can compensate
for the resulting wing-heavy condition by adjusting the
The pilot must be knowledgeable about the effect of weight
on the performance of the particular aircraft being flown. Lateral or longitudinal unbalance
Preflight planning should include a check of performance
charts to determine if the aircraft’s weight may contribute Empty Full
to hazardous flight operations. Excessive weight in itself
reduces the safety margins available to the pilot, and becomes
even more hazardous when other performance-reducing
factors are combined with excess weight. The pilot must
also consider the consequences of an overweight aircraft if
an emergency condition arises. If an engine fails on takeoff
or airframe ice forms at low altitude, it is usually too late to
reduce an aircraft’s weight to keep it in the air.

Weight Changes Lateral unbalance will cause wing heaviness.
The operating weight of an aircraft can be changed by Excess baggage
simply altering the fuel load. Gasoline has considerable
weight—6 pounds per gallon. Thirty gallons of fuel may
weigh more than one passenger. If a pilot lowers airplane
weight by reducing fuel, the resulting decrease in the range
of the airplane must be taken into consideration during flight
planning. During flight, fuel burn is normally the only weight
change that takes place. As fuel is used, an aircraft becomes
lighter and performance is improved.

Changes of fixed equipment have a major effect upon the Longitudinal unbalance will cause
weight of an aircraft. The installation of extra radios or either nose or tail heaviness.
instruments, as well as repairs or modifications may also Figure 9-1. Lateral and longitudinal unbalance.
affect the weight of an aircraft.

9-2

trim or by holding a constant control pressure. This action pilot’s operating handbook (POH). If the CG is not within the
places the aircraft controls in an out-of-streamline condition, allowable limits after loading, it will be necessary to relocate
increases drag, and results in decreased operating efficiency. some items before flight is attempted.
Since lateral balance is addressed when needed in the aircraft
flight manual (AFM) and longitudinal balance is more The forward CG limit is often established at a location that
critical, further reference to balance in this handbook means is determined by the landing characteristics of an aircraft.
longitudinal location of the CG. A single pilot operating a During landing, one of the most critical phases of flight,
small rotorcraft, may require additional weight to keep the exceeding the forward CG limit may result in excessive loads
aircraft laterally balanced. on the nosewheel, a tendency to nose over on tailwheel type
airplanes, decreased performance, higher stalling speeds, and
Flying an aircraft that is out of balance can produce increased higher control forces.
pilot fatigue with obvious effects on the safety and efficiency
of flight. The pilot’s natural correction for longitudinal Control
unbalance is a change of trim to remove the excessive control In extreme cases, a CG location that is beyond the forward
pressure. Excessive trim, however, has the effect of reducing limit may result in nose heaviness, making it difficult or
not only aerodynamic efficiency but also primary control impossible to flare for landing. Manufacturers purposely
travel distance in the direction the trim is applied. place the forward CG limit as far rearward as possible to
aid pilots in avoiding damage when landing. In addition to
Effects of Adverse Balance decreased static and dynamic longitudinal stability, other
Adverse balance conditions affect flight characteristics in undesirable effects caused by a CG location aft of the
much the same manner as those mentioned for an excess allowable range may include extreme control difficulty,
weight condition. It is vital to comply with weight and violent stall characteristics, and very light control forces
balance limits established for all aircraft, especially rotorcraft. which make it easy to overstress an aircraft inadvertently.
Operating above the maximum weight limitation compromises
the structural integrity of the rotorcraft and adversely affects A restricted forward CG limit is also specified to assure
performance. Balance is also critical because on some fully that sufficient elevator/control deflection is available at
loaded rotorcraft, CG deviations as small as three inches can minimum airspeed. When structural limitations do not limit
dramatically change handling characteristics. Stability and the forward CG position, it is located at the position where
control are also affected by improper balance. full-up elevator/control deflection is required to obtain a high
AOA for landing.
Stability
Loading in a nose-heavy condition causes problems in The aft CG limit is the most rearward position at which the
controlling and raising the nose, especially during takeoff CG can be located for the most critical maneuver or operation.
and landing. Loading in a tail heavy condition has a serious As the CG moves aft, a less stable condition occurs, which
effect upon longitudinal stability, and reduces the capability decreases the ability of the aircraft to right itself after
to recover from stalls and spins. Tail heavy loading also maneuvering or turbulence.
produces very light control forces, another undesirable
characteristic. This makes it easy for the pilot to inadvertently For some aircraft, both fore and aft CG limits may be
overstress an aircraft. specified to vary as gross weight changes. They may also
be changed for certain operations, such as acrobatic flight,
It is important to reevaluate the balance in a rotorcraft retraction of the landing gear, or the installation of special
whenever loading changes. In most aircraft, off-loading a loads and devices that change the flight characteristics.
passenger is unlikely to adversely affect the CG, but off-
loading a passenger from a rotorcraft can create an unsafe The actual location of the CG can be altered by many variable
flight condition. An out-of-balance loading condition also factors and is usually controlled by the pilot. Placement of
decreases maneuverability since cyclic control is less baggage and cargo items determines the CG location. The
effective in the direction opposite to the CG location. assignment of seats to passengers can also be used as a means
of obtaining a favorable balance. If an aircraft is tail heavy, it
Limits for the location of the CG are established by the is only logical to place heavy passengers in forward seats.
manufacturer. These are the fore and aft limits beyond
which the CG should not be located for flight. These limits Fuel burn can also affect the CG based on the location of the
are published for each aircraft in the Type Certificate Data fuel tanks. For example, most small aircraft carry fuel in the
Sheet (TCDS), or aircraft specification and the AFM or

9-3

wings very near the CG and burning off fuel has little effect criteria for negligible weight change is outlined in Advisory
on the loaded CG. On rotorcraft, the fuel tanks are often Circular (AC) 43.13-1 (as revised), Methods Techniques and
located behind the CG and fuel consumption from a tank Practices—Aircraft Inspection and Repair:
aft of the rotor mast causes the loaded CG to move forward.
A rotorcraft in this condition has a nose-low attitude when • One pound or less for an aircraft whose weight empty
coming to a hover following a vertical takeoff. Excessive is less than 5,000 pounds;
rearward displacement of the cyclic control is needed to
maintain a hover in a no-wind condition. Flight should not • Two pounds or less for aircraft with an empty weight
be continued since rearward cyclic control fades as fuel is of more than 5,000 pounds to 50,000 pounds;
consumed. Deceleration to a stop may also be impossible. In
the event of engine failure and autorotation, there may not be • Five pounds or less for aircraft with an empty weight
enough cyclic control to flare properly for a landing. of more than 50,000 pounds.

Management of Weight and Balance Control Negligible CG change is any change of less than 0.05 percent
Title 14 of the Code of Federal Regulations (14 CFR) section Mean Aerodynamic Chord (MAC) for fixed-wing aircraft, 0.2
23.23 requires establishment of the ranges of weights and percent of the maximum allowable CG range for rotorcraft.
CGs within which an aircraft may be operated safely. The Exceeding these limits would require a weight and balance
manufacturer provides this information, which is included in check.
the approved AFM, TCDS, or aircraft specifications.
Before any flight, the pilot should determine the weight
While there are no specified requirements for a pilot operating and balance condition of the aircraft. Simple and orderly
under 14 CFR part 91 to conduct weight and balance procedures based on sound principles have been devised
calculations prior to each flight, 14 CFR section 91.9 requires by the manufacturer for the determination of loading
the pilot in command (PIC) to comply with the operating conditions. The pilot uses these procedures and exercises
limits in the approved AFM. These limits include the weight good judgment when determining weight and balance. In
and balance of the aircraft. To enable pilots to make weight many modern aircraft, it is not possible to fill all seats,
and balance computations, charts and graphs are provided baggage compartments, and fuel tanks, and still remain within
in the approved AFM. the approved weight and balance limits. If the maximum
passenger load is carried, the pilot must often reduce the fuel
load or reduce the amount of baggage.

Weight and balance control should be a matter of concern to 14 CFR part 125 requires aircraft with 20 or more seats or
all pilots. The pilot controls loading and fuel management weighing 6,000 pounds or more to be weighed every 36
(the two variable factors that can change both total weight calendar months. Multi-engine aircraft operated under a
and CG location) of a particular aircraft. The aircraft owner 14 CFR part 135 are also required to be weighed every 36
or operator should make certain that up-to-date information months. Aircraft operated under 14 CFR part 135 are exempt
is available for pilot use, and should ensure that appropriate from the 36 month requirement if operated under a weight
entries are made in the records when repairs or modifications and balance system approved in the operations specifications
have been accomplished. The removal or addition of of the certificate holder. AC 43.13-1, Acceptable Methods,
equipment results in changes to the CG. Techniques and Practices—Aircraft Inspection and Repair
also requires that the aircraft mechanic must ensure the
Weight changes must be accounted for and the proper weight and balance data in the aircraft records is current and
notations made in weight and balance records. The equipment accurate after a 100-hour or annual inspection.
list must be updated, if appropriate. Without such information,
the pilot has no foundation upon which to base the necessary Terms and Definitions
calculations and decisions. The pilot should be familiar with terms used in working
problems related to weight and balance. The following list
Standard parts with negligible weight or the addition of of terms and their definitions is standardized, and knowledge
minor items of equipment such as nuts, bolts, washers, of these terms aids the pilot to better understand weight and
rivets, and similar standard parts of negligible weight on balance calculations of any aircraft. Terms defined by the
fixed-wing aircraft do not require a weight and balance General Aviation Manufacturers Association (GAMA) as
check. Rotorcraft are, in general, more critical with respect industry standard are marked in the titles with GAMA.
to control with changes in the CG position. The following

9-4

• Arm (moment arm)—the horizontal distance in inches • Maximum takeoff weight—the maximum allowable
from the reference datum line to the CG of an item. The weight for takeoff.
algebraic sign is plus (+) if measured aft of the datum,
and minus (–) if measured forward of the datum. • Maximum weight—the maximum authorized weight
of the aircraft and all of its equipment as specified in
• Basic empty weight (GAMA)—the standard empty the TCDS for the aircraft.
weight plus the weight of optional and special
equipment that have been installed. • Maximum zero fuel weight (GAMA)—the maximum
weight, exclusive of usable fuel.
• Center of gravity (CG)—the point about which an
aircraft would balance if it were possible to suspend • Mean aerodynamic chord (MAC)—the average
it at that point. It is the mass center of the aircraft, distance from the leading edge to the trailing edge of
or the theoretical point at which the entire weight of the wing.
the aircraft is assumed to be concentrated. It may be
expressed in inches from the reference datum, or in • Moment—the product of the weight of an item
percent of MAC. The CG is a three-dimensional point multiplied by its arm. Moments are expressed in
with longitudinal, lateral, and vertical positioning in pound-inches (in-lb). Total moment is the weight of
the aircraft. the airplane multiplied by the distance between the
datum and the CG.
• CG limits—the specified forward and aft points
within which the CG must be located during flight. • Moment index (or index)—a moment divided by a
These limits are indicated on pertinent aircraft constant such as 100, 1,000, or 10,000. The purpose of
specifications. using a moment index is to simplify weight and balance
computations of aircraft where heavy items and long
• CG range—the distance between the forward and aft CG arms result in large, unmanageable numbers.
limits indicated on pertinent aircraft specifications.
• Payload (GAMA)—the weight of occupants, cargo,
• Datum (reference datum)—an imaginary vertical plane and baggage.
or line from which all measurements of arm are taken.
The datum is established by the manufacturer. Once • Standard empty weight (GAMA)—aircraft weight
the datum has been selected, all moment arms and the that consists of the airframe, engines, and all items of
location of CG range are measured from this point. operating equipment that have fixed locations and are
permanently installed in the aircraft, including fixed
• Delta—a Greek letter expressed by the symbol  to ballast, hydraulic fluid, unusable fuel, and full engine
indicate a change of values. As an example, CG oil.
indicates a change (or movement) of the CG.
• Standard weights—established weights for numerous
• Floor load limit—the maximum weight the floor items involved in weight and balance computations.
can sustain per square inch/foot as provided by the These weights should not be used if actual weights are
manufacturer. available. Some of the standard weights are:

• Fuel load—the expendable part of the load of the Gasoline ............................................... 6 lb/US gal
aircraft. It includes only usable fuel, not fuel required
to fill the lines or that which remains trapped in the Jet A, Jet A-1 .................................... 6.8 lb/US gal
tank sumps.
Jet B ...................................................6.5 lb/US gal
• Licensed empty weight—the empty weight that
consists of the airframe, engine(s), unusable fuel, and Oil ......................................................7.5 lb/US gal
undrainable oil plus standard and optional equipment
as specified in the equipment list. Some manufacturers Water .....................................................8.35 lb/US gal
used this term prior to GAMA standardization.
• Station—a location in the aircraft that is identified by
• Maximum landing weight—the greatest weight that a number designating its distance in inches from the
an aircraft normally is allowed to have at landing. datum. The datum is, therefore, identified as station
zero. An item located at station +50 would have an
• Maximum ramp weight—the total weight of a loaded arm of 50 inches.
aircraft, and includes all fuel. It is greater than the
takeoff weight due to the fuel that will be burned during • Useful load—the weight of the pilot, copilot,
the taxi and runup operations. Ramp weight may also passengers, baggage, usable fuel, and drainable oil. It is
be referred to as taxi weight. the basic empty weight subtracted from the maximum
allowable gross weight. This term applies to general
aviation (GA) aircraft only.

9-5

Principles of Weight and Balance Computations of the object or part is often referred to as the station. If
It might be advantageous at this point to review and the weight of any object or component is multiplied by the
discuss some of the basic principles of weight and balance distance from the datum (arm), the product is the moment.
determination. The following method of computation can be The moment is the measurement of the gravitational force
applied to any object or vehicle for which weight and balance that causes a tendency of the weight to rotate about a point
information is essential. or axis and is expressed in inch-pounds (in-lb).

By determining the weight of the empty aircraft and adding To illustrate, assume a weight of 50 pounds is placed on
the weight of everything loaded on the aircraft, a total weight the board at a station or point 100 inches from the datum.
can be determined—a simple concept. A greater problem, The downward force of the weight can be determined by
particularly if the basic principles of weight and balance are multiplying 50 pounds by 100 inches, which produces a
not understood, is distributing this weight in such a manner moment of 5,000 in-lb. [Figure 9-3]
that the entire mass of the loaded aircraft is balanced around a
point (CG) that must be located within specified limits. Datum

The point at which an aircraft balances can be determined by 100"
locating the CG, which is, as stated in the definitions of terms,
the imaginary point at which all the weight is concentrated. To 50
provide the necessary balance between longitudinal stability lb
and elevator control, the CG is usually located slightly
forward of the center of lift. This loading condition causes Fulcrum Moment = 5,000 in-lb
a nose-down tendency in flight, which is desirable during
flight at a high AOA and slow speeds. Note: The datum is assumed to be Wt x Arm = Moment
located at the fulcrum. (lb) x (in) = (in-lb)
As mentioned earlier, a safe zone within which the balance 50 x 100 = 5,000
point (CG) must fall is called the CG range. The extremities
of the range are called the forward CG limits and aft CG Figure 9-3. Determining moment.
limits. These limits are usually specified in inches, along the
longitudinal axis of the airplane, measured from a reference To establish a balance, a total of 5,000 in-lb must be applied
point called a datum reference. The datum is an arbitrary to the other end of the board. Any combination of weight
point, established by aircraft designers, which may vary in and distance which, when multiplied, produces a 5,000 in-lb
location between different aircraft. [Figure 9-2] moment will balance the board. For example (illustrated in
Figure 9-4), if a 100-pound weight is placed at a point (station)
The distance from the datum to any component part or any 25 inches from the datum, and another 50-pound weight is
object loaded on the aircraft, is called the arm. When the placed at a point (station) 50 inches from the datum, the sum
object or component is located aft of the datum, it is measured of the product of the two weights and their distances will total
in positive inches; if located forward of the datum, it is a moment of 5,000 in-lb, which will balance the board.
measured as negative inches, or minus inches. The location

CG Datum 100"
range
Fwd limit Aft limit 50"
Datum
25"

50 100 50
lb lb lb

(–) (+)
Arm Arm

Moment = 700 in-lb Fulcrum
( + ) Arm 70"
2,500 2,500 5,000
in-lb in-lb in-lb

10 lb Wt x Arm = Moment 100 x 25 = 2,500
Sta 70
(lb) x (in) = (in-lb) 50 x 50 = 2,500

Sta 0 Total = 5,000

Figure 9-2. Weight and balance. Figure 9-4. Establishing a balance.

9-6

Weight and Balance Restrictions 2. Enter the moment for each item listed. Remember
An aircraft’s weight and balance restrictions should be “weight x arm = moment.”
closely followed. The loading conditions and empty weight
of a particular aircraft may differ from that found in the 3. Find the total weight and total moment.
AFM/POH because modifications or equipment changes may
have been made. Sample loading problems in the AFM/POH 4. To determine the CG, divide the total moment by the
are intended for guidance only; therefore, each aircraft must total weight.
be treated separately. Although an aircraft is certified for a
specified maximum gross takeoff weight, it will not safely NOTE: The weight and balance records for a particular
take off with this load under all conditions. Conditions that aircraft will provide the empty weight and moment, as well
affect takeoff and climb performance, such as high elevations, as the information on the arm distance. [Figure 9-5]
high temperatures, and high humidity (high density altitudes)
may require a reduction in weight before flight is attempted. Item Weight Arm Moment
Other factors to consider prior to takeoff are runway length,
runway surface, runway slope, surface wind, and the presence Aircraft Empty Weight 2,100 78.3 164,430
of obstacles. These factors may require a reduction in weight Front Seat Occupants 340 85.0 28,900
prior to flight. Rear Seat Occupants 350 121.0 42,350
Fuel 450 75.0 33,750
Some aircraft are designed so that it is difficult to load them Baggage Area 1 80 150.0 12,000
in a manner that will place the CG out of limits. These are
usually small aircraft with the seats, fuel, and baggage areas Total 3,320 281,430
located near the CG limit. Pilots must be aware that while
within CG limits these aircraft can be overloaded in weight. 281,430 ÷ 3,320 = 84.8
Other aircraft can be loaded in such a manner that they will
be out of CG limits even though the useful load has not been Figure 9-5. Example of weight and balance computations.
exceeded. Because of the effects of an out-of-balance or
overweight condition, a pilot should always be sure that an The total loaded weight of 3,320 pounds does not exceed
aircraft is properly loaded. the maximum gross weight of 3,400 pounds, and the CG of
84.8 is within the 78–86 inch range; therefore, the aircraft is
Determining Loaded Weight and CG loaded within limits.

There are various methods for determining the loaded weight Graph Method
and CG of an aircraft. There is the computational method, as Another method for determining the loaded weight and CG is
well as methods that utilize graphs and tables provided by the use of graphs provided by the manufacturers. To simplify
the aircraft manufacturer. calculations, the moment may sometimes be divided by 100,
1,000, or 10,000. [Figures 9-6, 9-7, and 9-8]
Computational Method
The following is an example of the computational method Front seat occupants....................................340 pounds
involving the application of basic math functions.
Rear seat occupants ......................................300 pounds
Aircraft Allowances:
Fuel.................................................................40 gallons
Maximum gross weight...................... 3,400 pounds
Baggage area 1 ...............................................20 pounds
CG range............................................. 78–86 inches
Sample Loading Problem Weight (lb) Moment
Given: (in-lb/1,000)
1. Basic Empty Weight (Use data pertaining 1,467
Weight of front seat occupants............. 340 pounds to aircraft as it is presently equipped.) 240 57.3
Includes unusable fuel and full oil 11.5
Weight of rear seat occupants.............. 350 pounds 340
2. Usable Fuel (At 6 lb/gal) 300 12.7
Fuel........................................................... 75 gallons Standard Tanks (40 gal maximum) 21.8
Long Range Tanks (50 gal maximum) 20
Weight of baggage in area 1....................80 pounds Integral Tanks (62 gal maximum) 1.9
Integral Reduced Fuel (42 gal) 2,367
1. List the weight of the aircraft, occupants, fuel, and 105.2
baggage. Remember that aviation gas (AVGAS) 3. Pilot and Front Passenger (Station 34 to 46)
weighs 6 pounds per gallon and is used in this
example. 4. Rear Passengers

5. Baggage Area 1 or Passenger on Child’s
Seat (Station 82 to 108, 120 lb maximum)

6. Baggage Area 2
(Station 108 to 142, 50 lb maximum)

7. Weight and Moment

Figure 9-6. Weight and balance data.

9-7

Loading Graph

0 Load Moment/1,000 (Kilogram-Millimeters)
50 100 150 200 250 300 350 400
400
350 200
340
300 62 gal***(234.7 liters) Passengers 175
250 gal (227.1 liters) 150
200 Rear 125
Load Weight (Pounds)150 60 100
Pilot & FFruoenlt(6P labs/sgeal;ng0.e7r2 kg/liter100
50 gal**(189.3 liters)
Load Weight (Kilograms)50
0 42 gal reduced***(159 liters)
40 gal*(189.3 liters)

30 gal (113.6 liters) 75

20 gal (75.7 liters) Maximum Usable Fuel 50
* Standard Tanks 25
Baggage Area 1 or ** Long Range Tanks 0
Passenger on Child’s Seat
*** Internal Tanks
10 gal (37.9 liters)
25 30
Baggage Area 2

5 10 12.7 15 20
Load Moment/1,000 (Inch-Pounds)

Figure 9-7. Loading graph. Center of gravity moment envelope

Loaded Airplane Moment/1,000 (Kilogram-Millimeters)

600 700 800 900 1,000 1,100 1,200 1,300

Loaded Airplane Weight (Pounds)2,400 Normal 1,100
Utility Category2,367 Category 1,050
2,300 1,000
Loaded Airplane Weight (Kilograms)
2,200 950
900
2,100 850
800
2,000 750
700
1,900

1,800

1,700

1,600

1,50045 50 55 60 65 70 75 80 85 90 95 100 105 110
Loaded Aircraft Moment/1,000 (Inch-Pounds) 105.2

Figure 9-8. CG moment envelope.

9-8

The same steps should be followed as in the computational the lines intersect within the envelope, the aircraft is loaded
method except the graphs provided will calculate the within limits. In this sample loading problem, the aircraft is
moments and allow the pilot to determine if the aircraft is loaded within limits.
loaded within limits. To determine the moment using the
loading graph, find the weight and draw a line straight across Table Method
until it intercepts the item for which the moment is to be The table method applies the same principles as the
calculated. Then draw a line straight down to determine the computational and graph methods. The information
moment. (The red line on the loading graph represents the and limitations are contained in tables provided by the
moment for the pilot and front passenger. All other moments manufacturer. Figure 9-9 is an example of a table and a
were determined in the same way.) Once this has been done weight and balance calculation based on that table. In this
for each item, total the weight and moments and draw a line problem, the total weight of 2,799 pounds and moment of
for both weight and moment on the CG envelope graph. If 2,278/100 are within the limits of the table.

Usable Fuel Occupants Minimum Maximum
Weight Moment Moment
Main Wing Tanks Arm 75 Front Seat Rear Seats
Arm 85 Arm 121 100 100
2,057
Gallons Weight Moment FUEL TANK Weight Moment Weight Moment 2,400 1,848 2,065
100 100 100 2,410 1,856 2,074
5 30 120 120 2,420 1,863 2,083
10 60 22 130 102 130 145 2,430 1,871 2,091
15 90 45 140 110 140 157 2,440 1,879 2,100
20 120 68 150 119 150 169 2,450 1,887 2,108
25 150 90 160 128 160 182 2,460 1,894 2,117
30 180 112 170 136 170 194 2,470 1,092 2,125
35 210 135 180 144 180 206 2,480 1,911 2,134
40 240 158 190 153 190 218 2,490 1,921
44 264 180 200 162 200 230 2,143
198 170 242 2,500 1,932 2,151
2,510 1,942 2,160
SEATING AREA 2,520 1,953 2,168
2,530 1,963 2,176
2,540 1,974 2,184
2,550 1,984 2,192
2,560 1,995 2,200
2,570 2,005 2,208
2,580 2,016 2,216
2,590 2,026
2,224
OIL BAGGAGE AREA 2,600 2,037 2,232
2,610 2,048 2,239
Quarts *Oil Moment Baggage or 5th Minimum Maximum 2,620 2,058 2,247
Weight 100 Seat Occupant Weight Moment Moment 2,630 2,069 2,255
2,640 2,080 2,263
Arm 140 100 100 2,650 2,090 2,271
2,660 2,101 2,279
10 19 5 FUEL TANK Weight Moment 2,100 1,617 1,800 2,670 2,112 2,287
*Included in basic empty weight 100 2,110 1,625 1,808 2,680 2,123 2,295
10 2,120 1,632 1,817 2,690 2,133
20 14 2,130 1,640 1,825 2,303
Empty Weight ~ 2,015 30 28 2,140 1,648 1,834 2,700 2,144 2,311
MOM/ 100 ~ 1,554 40 42 2,150 1,656 1,843 2,710 2,155 2,319
50 56 2,160 1,663 1,851 2,720 2,166 2,326
Moment Limits vs Weight 60 70 2,170 1,671 1,860 2,730 2,177 2,334
70 84 2,180 1,679 1,868 2,740 2,188 2,342
80 98 2,190 1,686 1,877 2,750 2,199 2,350
90 112 2,760 2,210 2,358
Moment limits are based on the following weight and 100 126 2,200 1,694 1,885 2,770 2,221 2,366
center of gravity limit data (landing gear down). 110 140 2,210 1,702 1,894 2,780 2,232 2,374
120 154 2,220 1,709 1,903 2,790 2,243
Weight Forward AFT 168 2,230 1,717 1,911 2,381
Condition CG Limit CG Limit 182 2,240 1,725 1,920 2,389
2,250 1,733 1,928 2,397
2,950 lb (takeoff 82.1 84.7 2,260 1,740 1,937 2,800 2,254 2,405
or landing) 77.5 85.7 2,270 1,748 1,945 2,810 2,265 2,413
2,525 lb 77.0 85.7 2,280 1,756 1,954 2,820 2,276 2,421
2,475 lb or less 2,290 1,763 1,963 2,830 2,287 2,426
2,300 1,771 1,971 2,840 2,298 2,436
Sample Loading Problem Weight Moment 2,310 1,779 1,980 2,850 2,309 2,444
2,320 1,786 1,988 2,860 2,320 2,452
Basic empty weight 2,015 1,554 2,330 1,794 1,997 2,870 2,332 3,460
Fuel main tanks (44 gal) 264 198 2,340 1,802 2,005 2,880 2,343 2,468
*Front seat passengers 300 254 2,350 1,810 2,014 2,890 2,354 2,475
*Rear seat passengers 190 230 2,360 1,817 2,023 2,900 2,365 2,483
Baggage 30 42 2,370 1,825 2,031 2,910 2,377 2,491
2,380 1,833 2,040 2,920 2,388 2,499
Total 2,799 2,278/100 2,390 1,840 2,048 2,930 2,399
2,940 2,411
2,950 2,422

* Interpolate or, as in this case, add appropriate numbers.

Figure 9-9. Loading schedule placard.

9-9

Computations With a Negative Arm most satisfactory solution to this problem is to shift baggage,
Figure 9-10 is a sample of weight and balance computation passengers, or both. The pilot should be able to determine
using an airplane with a negative arm. It is important to the minimum load shift needed to make the aircraft safe for
remember that a positive times a negative equals a negative, flight. Pilots should be able to determine if shifting a load to
and a negative would be subtracted from the total moments. a new location will correct an out-of-limit condition. There
are some standardized calculations that can help make these
Item Weight Arm Moment determinations.

Licensed Empty Weight 1,011.9 68.6 69,393.0 Weight Shifting
Oil (6 quarts) 11.0 −31.0 −341.0 When weight is shifted from one location to another, the
Fuel (18 gallons) 9,072.0 total weight of the aircraft is unchanged. The total moments,
Fuel, Auxiliary (18 gallons) 108.0 84.0 9,072.0 however, do change in relation and proportion to the
Pilot 108.0 84.0 direction and distance the weight is moved. When weight is
Passenger 170.0 81.0 13,770.0 moved forward, the total moments decrease; when weight
Baggage 170.0 81.0 13,770.0 is moved aft, total moments increase. The moment change
105.0 is proportional to the amount of weight moved. Since many
70.0 7,350.0 aircraft have forward and aft baggage compartments, weight
may be shifted from one to the other to change the CG. If
Total 1,648.9 122,086.0 starting with a known aircraft weight, CG, and total moments,
CG 74.0 calculate the new CG (after the weight shift) by dividing the
new total moments by the total aircraft weight.
Figure 9-10. Sample weight and balance using a negative.
To determine the new total moments, find out how many
Computations With Zero Fuel Weight moments are gained or lost when the weight is shifted.
Figure 9-11 is a sample of weight and balance computation Assume that 100 pounds has been shifted from station 30 to
using an aircraft with a zero fuel weight. In this example, station 150. This movement increases the total moments of
the total weight of the aircraft less fuel is 4,240 pounds, the aircraft by 12,000 in-lb.
which is under the zero fuel weight of 4,400 pounds. If the
total weight of the aircraft without fuel had exceeded 4,400 Moment when
pounds, passengers or cargo would have needed to be reduced at station 150 = 100 lb x 150 in = 15,000 in-lb
to bring the weight at or below the max zero fuel weight.
Moment when
Item Weight Arm Moment at station 30 = 100 lb x 30 in = 3,000 in-lb

Basic Empty Weight 3,230 CG 90.5 292,315.0 Moment change = [15,000 - 3,000] = 12,000 in-lb
Front Seat Occupants 335 89.0 29,815.0
By adding the moment change to the original moment (or
3rd & 4th Seat Occupants 350 126.0 44,100.0 subtracting if the weight has been moved forward instead of
Forward Facing aft), the new total moments are obtained. Then determine the
5th & 6th Seat Occupants 200 157.0 31,400.0 new CG by dividing the new moments by the total weight:
Nose Baggage 100 10.0 1,000.0
Aft Baggage 4,575.0 Total moments =
25 183.0 616,000 in-lb + 12,000 in-lb = 628,000 in-lb
Zero Fuel Weight Max 4,400 pounds
4,240 CG 95.1 403,205.0 CG = 628,000 in-lb = 78.5 in
Subtotal 8,000 lb

Fuel 822 113.0 92,886.0 The shift has caused the CG to shift to station 78.5.

Ramp Weight Max 5,224 pounds 5,062 CG 98.0 496,091.0

Subtotal Ramp Weight

* Less Fuel for Start, −24 113.0 −2,712.0
Taxi, and Takeoff 5,038 CG 97.9 493,379.0
Subtotal
Takeoff Weight

Less Fuel to Destination −450 113.0 −50,850.0

Max Landing Weight 4,940 pounds 4,588 CG 96.5 442,529.0

Actual Landing Weight

*Fuel for start, taxi, and takeoff is normally 24 pounds.

Figure 9-11. Sample weight and balance using an aircraft with a
published zero fuel weight.

Shifting, Adding, and Removing Weight
A pilot must be able to solve any problems accurately that
involve the shift, addition, or removal of weight. For example,
the pilot may load the aircraft within the allowable takeoff
weight limit, then find a CG limit has been exceeded. The

9-10

A simpler solution may be obtained by using a computer Example
or calculator and a proportional formula. This can be done
because the CG will shift a distance that is proportional to Given:
the distance the weight is shifted. Aircraft total weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6,100 lb
CG station. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80.0 in

Example Determine the location of the CG if 100 pounds is removed from
station 150.

Weight shifted = ∆CG (change of CG) Solution:
Total weight Distance weight is shifted
Weight removed ∆CG
100 = ∆CG New total weight = Distance between weight and old CG
8,000 120
100 lb ∆CG
∆CG = 1.5 in 6,100 lb − 100 lb = 150 in − 80 in

The change of CG is added to (or subtracted from when appropriate) 100 lb = ∆CG
the original CG to determine the new CG: 6,000 lb 70 in
77 + 1.5 = 78.5 inches aft of datum

The shifting weight proportion formula can also be used to determine CG = 1.2 in forward
how much weight must be shifted to achieve a particular shift of the CG.
The following problem illustrates a solution of this type. Subtract ∆CG from old CG
New CG = 80 in − 1.2 in = 78.8 in

Example when the aircraft burns fuel in flight, thereby reducing the
weight located at the fuel tanks. Most small aircraft are
Given: designed with the fuel tanks positioned close to the CG;
Aircraft total weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7,800 lb therefore, the consumption of fuel does not affect the CG to
CG station. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81.5 in any great extent.
Aft CG limit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80.5 in
The addition or removal of cargo presents a CG change
Determine how much cargo must be shifted from the aft cargo compart- problem that must be calculated before flight. The problem
ment at station 150 to the forward cargo compartment at station 30 to may always be solved by calculations involving total
move the CG to exactly the aft limit. moments. A typical problem may involve the calculation
of a new CG for an aircraft which, when loaded and ready
Solution: for flight, receives some additional cargo or passengers just
before departure time.
Weight to be shifted = CG
Total weight Distance weight is shifted

Weight to be shifted = 1.0 in
7,800 lb 120 in

Weight to be shifted = 65 lb

Weight Addition or Removal In the previous examples, the CG is either added or
subtracted from the old CG. Deciding which to accomplish is
In many instances, the weight and balance of the aircraft will best handled by mentally calculating which way the CG will
be changed by the addition or removal of weight. When this shift for the particular weight change. If the CG is shifting
happens, a new CG must be calculated and checked against aft, the CG is added to the old CG; if the CG is shifting
the limitations to see if the location is acceptable. This type forward, the CG is subtracted from the old CG.
of weight and balance problem is commonly encountered

Example Chapter Summary

Given: Operating an aircraft within the weight and balance limits
Aircraft total weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6,860 lb is critical to flight safety. Pilots must ensure that the CG is
CG station. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80.0 in and remains within approved limits for all phases of a flight.
Additional information on weight, balance, CG, and aircraft
Determine the location of the CG if 140 pounds of baggage is added to stability can be found in FAA-H-8083-1, Aircraft Weight
station 150. and Balance Handbook. Those pilots flying helicopters or
gyroplanes should consult the Rotorcraft Flying Handbook,
Solution: FAA-H-8083-21, for specific information relating to aircraft
type.
Added weight = ∆CG
New total weight Distance between weight and old CG

140 lb = ∆CG
6,860 lb + 140 lb 150 in − 80 in

140 lb = ∆CG
7,000 lb 70 in

CG = 1.4 in aft

Add ∆CG to old CG
New CG = 80 in + 1.4 in = 81.4 in

9-11

9-12

AircraftChapter 10
Performance

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.

10-1

It must be emphasized that the manufacturers’ information pounds per square inch (psi). The density of air has significant
and data furnished in the AFM/POH is not standardized. effects on the aircraft’s performance. As air becomes less
Some provide the data in tabular form, while others use dense, it reduces:
graphs. In addition, the performance data may be presented
on the basis of standard atmospheric conditions, pressure • Power, because the engine takes in less air.
altitude, or density altitude. The performance information in
the AFM/POH has little or no value unless the user recognizes • Thrust, because the propeller is less efficient in thin
those variations and makes the necessary adjustments. air.

To be able to make practical use of the aircraft’s capabilities • Lift, because the thin air exerts less force on the
and limitations, it is essential to understand the significance airfoils.
of the operational data. The pilot must be cognizant of the
basis for the performance data, as well as the meanings of The pressure of the atmosphere varies with time and altitude.
the various terms used in expressing performance capabilities Due to the changing atmospheric pressure, a standard
and limitations. reference was developed. The standard atmosphere at sea level
is a surface temperature of 59 degrees Fahrenheit (°F) or 15
Since the characteristics of the atmosphere have a major degrees Celsius (°C) and a surface pressure of 29.92 inches of
effect on performance, it is necessary to review two dominant mercury ("Hg) or 1013.2 millibars (mb). [Figure 10-1]
factors—pressure and temperature.
Standard Inches of Millibars Standard
Structure of the Atmosphere Sea Level Mercury 1016 Sea Level
Pressure 30 847 Pressure
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 29.92"Hg 25 1013 mb
its land and water. However, air differs from land and water
inasmuch as it is a mixture of gases. It has mass, weight, and 20 677
indefinite shape.
15 508
Air, like any other fluid, is able to flow and change its shape
when subjected to even minute pressures because of the lack of 10 339
strong molecular cohesion. For example, gas will completely
fill any container into which it is placed, expanding or 5 170
contracting to adjust its shape to the limits of the container. Atmospheric Pressure

00

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 Figure 10-1. Standard sea level pressure.

Though there are various kinds of pressure, pilots are mainly A standard temperature lapse rate is one in which the
concerned with atmospheric pressure. It is one of the basic temperature decreases at the rate of approximately 3.5 °F or
factors in weather changes, helps to lift the aircraft, and 2 °C per thousand feet up to 36,000 feet. Above this point,
actuates some of the most important flight instruments in the temperature is considered constant up to 80,000 feet. A
the aircraft. These instruments often include the altimeter, standard pressure lapse rate is one in which pressure decreases
the airspeed indicator (ASI), the vertical speed indicator, and at a rate of approximately 1 "Hg per 1,000 feet of altitude gain
the manifold pressure gauge. to 10,000 feet. [Figure 10-2] The International Civil Aviation
Organization (ICAO) has established this as a worldwide
Though air is very light, it has mass and is affected by the standard, and it is often referred to as International Standard
attraction of gravity. Therefore, like any other substance, it Atmosphere (ISA) or ICAO Standard Atmosphere. Any
has weight; because it has weight, it has force. Since it is a temperature or pressure that differs from the standard lapse
fluid substance, this force is exerted equally in all directions, rates is considered nonstandard temperature and pressure.
and its effect on bodies within the air is called pressure. Adjustments for nonstandard temperatures and pressures are
Under standard conditions at sea level, the average pressure provided on the manufacturer’s performance charts.
exerted by the weight of the atmosphere is approximately 14.7

10-2

Standard Atmosphere

Altitude (ft) Pressure Temperature 2. By applying a correction factor to the indicated altitude
("Hg) (°C) (°F) according to the reported “altimeter setting.”
0
1,000 29.92 15.0 59.0 Density Altitude
2,000 28.86 13.0 55.4
3,000 27.82 11.0 51.9 The more appropriate term for correlating aerodynamic
4,000 26.82 48.3 performance in the nonstandard atmosphere is density
5,000 25.84 9.1 44.7 altitude—the altitude in the standard atmosphere
6,000 24.89 7.1 41.2 corresponding to a particular value of air density.
7,000 23.98 5.1 37.6
8,000 23.09 3.1 34.0 Density altitude is pressure altitude corrected for nonstandard
9,000 22.22 1.1 30.5 temperature. As the density of the air increases (lower
10,000 21.38 −0.9 26.9 density altitude), aircraft performance increases. Conversely,
11,000 20.57 −2.8 23.3 as air density decreases (higher density altitude), aircraft
12,000 19.79 −4.8 19.8 performance decreases. A decrease in air density means a
13,000 19.02 −6.8 16.2 high density altitude; an increase in air density means a lower
14,000 18.29 −8.8 12.6 density altitude. Density altitude is used in calculating aircraft
15,000 17.57 −10.8 performance. Under standard atmospheric condition, air at
16,000 16.88 −12.7 9.1 each level in the atmosphere has a specific density; under
17,000 16.21 −14.7 5.5 standard conditions, pressure altitude and density altitude
18,000 15.56 −16.7 1.9 identify the same level. Density altitude, then, is the vertical
19,000 14.94 −18.7 −1.6 distance above sea level in the standard atmosphere at which
20,000 14.33 −20.7 −5.2 a given density is to be found.
13.74 −22.6 −8.8
−24.6 −12.3

Figure 10-2. Properties of standard atmosphere.

Since all aircraft performance is compared and evaluated with The computation of density altitude must involve consideration
respect to the standard atmosphere, all aircraft instruments of pressure (pressure altitude) and temperature. Since aircraft
are calibrated for the standard atmosphere. Thus, certain performance data at any level is based upon air density under
corrections must apply to the instrumentation, as well as the standard day conditions, such performance data apply to air
aircraft performance, if the actual operating conditions do density levels that may not be identical to altimeter indications.
not fit the standard atmosphere. In order to account properly Under conditions higher or lower than standard, these levels
for the nonstandard atmosphere, certain related terms must cannot be determined directly from the altimeter.
be defined.
Density altitude is determined by first finding pressure
Pressure Altitude altitude, and then correcting this altitude for nonstandard
temperature variations. Since density varies directly with
Pressure altitude is the height above the standard datum pressure, and inversely with temperature, a given pressure
plane (SDP). The aircraft altimeter is essentially a sensitive altitude may exist for a wide range of temperature by allowing
barometer calibrated to indicate altitude in the standard the density to vary. However, a known density occurs for any
atmosphere. If the altimeter is set for 29.92 "Hg SDP, the one temperature and pressure altitude. The density of the air,
altitude indicated is the pressure altitude—the altitude in the of course, has a pronounced effect on aircraft and engine
standard atmosphere corresponding to the sensed pressure. performance. Regardless of the actual altitude at which the
aircraft is operating, it will perform as though it were operating
The SDP is a theoretical level where the pressure of the at an altitude equal to the existing density altitude.
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 For example, when set at 29.92 "Hg, the altimeter may
above sea level. Pressure altitude is important as a basis for indicate a pressure altitude of 5,000 feet. According to the
determining aircraft performance, as well as for assigning AFM/POH, the ground run on takeoff may require a distance
flight levels to aircraft operating at above 18,000 feet. of 790 feet under standard temperature conditions.

The pressure altitude can be determined by either of two However, if the temperature is 20 °C above standard, the
methods: expansion of air raises the density level. Using temperature
correction data from tables or graphs, or by deriving the
1. By setting the barometric scale of the altimeter to
29.92 "Hg and reading the indicated altitude, or

10-3

density altitude with a computer, it may be found that the 15,000
density level is above 7,000 feet, and the ground run may be
closer to 1,000 feet. 13,000 14,000

Air density is affected by changes in altitude, temperature, 14,000 10,000 11,000 12,000
and humidity. High density altitude refers to thin air while 13,000 Pressure altitude (feet)
low density altitude refers to dense air. The conditions that 12,000
result in a high density altitude are high elevations, low 11,000
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 10,000 Standard temperature9,000
by inputting the pressure altitude and outside air temperature
at flight level. Density altitude can also be determined by 9,000 8,000
referring to the table and chart in Figures 10-3 and 10-4.
8,000
Method for Determining Alternate Method for Determining Density altitude (feet)
Pressure Altitude 7,000
Pressure Altitude 6,000 7,000
To field 6,000
Altimeter Altitude elevation
setting correction 5,000
To get
28.0 1,825 pressure altitude 4,000

28.1 1,725 From field 3,000
elevation
28.2 1,630 2,000 5,000

28.3 1,535 1,000

28.4 1,435 Sea level
C -20° -10° 0° 10° 20° 30° 40°
28.5 1,340 F 0° 10° 20° 30° 40° 50° 60° 70° 80° 90° 100°2,000 3,000 4,000

28.6 1,245 Outside air temperature (OAT)

28.7 1,150 Figure 10-4. Density altitude chart.
Effects of Pressure on Density
28.8 1,050 Since air is a gas, it can be compressed or expanded. When
air is compressed, a greater amount of air can occupy a
28.9 955 given volume. Conversely, when pressure on a given volume
of air is decreased, the air expands and occupies a greater
Field elevation is sea level29.0 865 space. That is, the original column of air at a lower pressure
Subtract Add contains a smaller mass of air. In other words, the density is
29.1 770 decreased. In fact, density is directly proportional to pressure.
If the pressure is doubled, the density is doubled, and if the
29.2 675 pressure is lowered, so is the density. This statement is true 1,000
only at a constant temperature. level
29.3 580

29.4 485 Sea

29.5 390 -1,000 -2,000

29.6 300

29.7 205

29.8 110

29.9 20

29.92 0

30.0 −75

30.1 −165

30.2 −255

30.3 −350

30.4 −440

30.5 −530

30.6 −620

30.7 −710

30.8 −805

30.9 −895

31.0 −965

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.

10-4

Effects of Temperature on Density take off in a very short distance is an important factor to the
Increasing the temperature of a substance decreases its density. pilot who operates in and out of short, unimproved airfields.
Conversely, decreasing the temperature increases the density. The ability to carry heavy loads, fly at high altitudes at fast
Thus, the density of air varies inversely with temperature. speeds, or travel long distances is essential performance for
This statement is true only at a constant pressure. operators of airline and executive type aircraft.

In the atmosphere, both temperature and pressure decrease The primary factors most affected by performance are the
with altitude, and have conflicting effects upon density. takeoff and landing distance, rate of climb, ceiling, payload,
However, the fairly rapid drop in pressure as altitude is range, speed, maneuverability, stability, and fuel economy.
increased usually has the dominant effect. Hence, pilots can Some of these factors are often directly opposed: for example,
expect the density to decrease with altitude. high speed versus short landing distance, long range versus
great payload, and high rate of climb versus fuel economy. It
Effects of Humidity (Moisture) on Density is the preeminence of one or more of these factors that dictates
The preceding paragraphs are based on the presupposition of differences between aircraft and explains the high degree of
perfectly dry air. In reality, it is never completely dry. The specialization found in modern aircraft.
small amount of water vapor suspended in the atmosphere
may be negligible under certain conditions, but in other The various items of aircraft performance result from the
conditions humidity may become an important factor in the combination of aircraft and powerplant characteristics. The
performance of an aircraft. Water vapor is lighter than air; aerodynamic characteristics of the aircraft generally define
consequently, moist air is lighter than dry air. Therefore, as the the power and thrust requirements at various conditions of
water content of the air increases, the air becomes less dense, flight, while powerplant characteristics generally define the
increasing density altitude and decreasing performance. It is power and thrust available at various conditions of flight.
lightest or least dense when, in a given set of conditions, it The matching of the aerodynamic configuration with the
contains the maximum amount of water vapor. powerplant is accomplished by the manufacturer to provide
maximum performance at the specific design condition (e.g.,
Humidity, also called relative humidity, refers to the amount range, endurance, and climb).
of water vapor contained in the atmosphere, and is expressed
as a percentage of the maximum amount of water vapor Straight-and-Level Flight
the air can hold. This amount varies with the temperature; All of the principal components of flight performance involve
warm air can hold more water vapor, while colder air can steady-state flight conditions and equilibrium of the aircraft.
hold less. Perfectly dry air that contains no water vapor has For the aircraft to remain in steady, level flight, equilibrium
a relative humidity of zero percent, while saturated air that must be obtained by a lift equal to the aircraft weight and a
cannot hold any more water vapor has a relative humidity powerplant thrust equal to the aircraft drag. Thus, the aircraft
of 100 percent. Humidity alone is usually not considered an drag defines the thrust required to maintain steady, level
essential factor in calculating density altitude and aircraft flight. As presented in Chapter 4, Aerodynamics of Flight,
performance; however, it does contribute. all parts of an aircraft contribute to the drag, either induced
(from lifting surfaces) or parasite drag.
The higher the temperature, the greater amount of water
vapor that the air can hold. When comparing two separate air While the parasite drag predominates at high speed, induced
masses, the first warm and moist (both qualities making air drag predominates at low speed. [Figure 10-5] For example,
lighter) and the second cold and dry (both qualities making
it heavier), the first must be less dense than the second. Total drag
Pressure, temperature, and humidity have a great influence
on aircraft performance because of their effect upon density. Drag L/DMAX Parasite drag
There is no rule-of-thumb or chart used to compute the effects Stall
of humidity on density altitude, but it must be taken into
consideration. Expect a decrease in overall performance in
high humidity conditions.

Performance Induced drag

Performance is a term used to describe the ability of an aircraft Speed
to accomplish certain things that make it useful for certain Figure 10-5. Drag versus speed.
purposes. For example, the ability of an aircraft to land and

10-5

if an aircraft in a steady flight condition at 100 knots is then Although the terms “power” and “thrust” are sometimes
accelerated to 200 knots, the parasite drag becomes four used interchangeably, erroneously implying that they are
times as great, but the power required to overcome that synonymous, it is important to distinguish between the two
drag is eight times the original value. Conversely, when the when discussing climb performance. Work is the product of
aircraft is operated in steady, level flight at twice as great a a force moving through a distance and is usually independent
speed, the induced drag is one-fourth the original value, and of time. Work is measured by several standards; the most
the power required to overcome that drag is only one-half common unit is called a foot-pound. If a one pound mass
the original value. is raised one foot, a work unit of one foot-pound has been
performed. The common unit of mechanical power is
When an aircraft is in steady, level flight, the condition of horsepower; one horsepower is work equivalent to lifting
equilibrium must prevail. The unaccelerated condition of flight 33,000 pounds a vertical distance of one foot in one minute.
is achieved with the aircraft trimmed for lift equal to weight The term power implies work rate or units of work per unit
and the powerplant set for a thrust to equal the aircraft drag. 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 maximum level flight speed for the aircraft will be obtained the force that imparts a change in the velocity of a mass. This
when the power or thrust required equals the maximum power force is measured in pounds but has no element of time or
or thrust available from the powerplant. [Figure 10-6] The rate. It can be said then, that during a steady climb, the rate
minimum level flight airspeed is not usually defined by thrust of climb is a function of excess thrust.
or power requirement since conditions of stall or stability and
control problems generally predominate. This relationship means that, for a given weight of an aircraft,
the angle of climb depends on the difference between thrust
Maximum available power and drag, or the excess power. [Figure 10-7] Of course, when
the excess thrust is zero, the inclination of the flightpath is
High cruise speed zero, and the aircraft will be in steady, level flight. When the
thrust is greater than the drag, the excess thrust will allow
Low cruise speed a climb angle depending on the value of excess thrust. On
the other hand, when the thrust is less than the drag, the
Min. speed deficiency of thrust will allow an angle of descent.
Power required
Maximum level flight speedT
Power required
Reserve thrusthrust avail
required or drag
Speed able Thrust
Stall
Figure 10-6. Power versus speed.
Speed for max. angle of climb
Climb Performance
Climb performance is a result of using the aircrafts potential Speed
energy provided by one, or a combination of two factors. The
first is the use of excess power above that required for level Figure 10-7. Thrust versus climb angle.
flight. An aircraft equipped with an engine capable of 200
horsepower (at a given altitude) but using 130 horsepower The most immediate interest in the climb angle performance
to sustain level flight (at a given airspeed) has 70 excess involves obstacle clearance. The most obvious purpose for
horsepower available for climbing. A second factor is that which it might be used is to clear obstacles when climbing
the aircraft can tradeoff its kinetic energy and increase its out of short or confined airports.
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.

10-6

The maximum angle of climb would occur where there the optimum. Of course, climb performance would be most
exists the greatest difference between thrust available and critical with high gross weight, at high altitude, in obstructed
thrust required; i.e., for the propeller-powered airplane, the takeoff areas, or during malfunction of a powerplant. Then,
maximum excess thrust and angle of climb will occur at some optimum climb speeds are necessary.
speed just above the stall speed. Thus, if it is necessary to
clear an obstacle after takeoff, the propeller-powered airplane Weight has a very pronounced effect on aircraft performance.
will attain maximum angle of climb at an airspeed close to—if If weight is added to an aircraft, it must fly at a higher angle
not at—the takeoff speed. of attack (AOA) to maintain a given altitude and speed. This
increases the induced drag of the wings, as well as the parasite
Of greater interest in climb performance are the factors that drag of the aircraft. Increased drag means that additional
affect the rate of climb. The vertical velocity of an aircraft thrust is needed to overcome it, which in turn means that less
depends on the flight speed and the inclination of the reserve thrust is available for climbing. Aircraft designers
flightpath. In fact, the rate of climb is the vertical component go to great effort to minimize the weight since it has such a
of the flightpath velocity. marked effect on the factors pertaining to performance.

llerFor rate of climb, the maximum rate would occur wherePower available and power requiredA change in an aircraft’s weight produces a twofold effect on
there exists the greatest difference between power availableReserve climb performance. First, a change in weight will change the
and power required. [Figure 10-8] The above relationshippower drag and the power required. This alters the reserve power
means that, for a given weight of an aircraft, the rate of climb available, which in turn, affects both the climb angle and the
depends on the difference between the power available andPower required climb rate. Secondly, an increase in weight will reduce the
the power required, or the excess power. Of course, when the maximum rate of climb, but the aircraft must be operated at a
excess power is zero, the rate of climb is zero and the aircraft higher climb speed to achieve the smaller peak climb rate.
is in steady, level flight. When power available is greater
than the power required, the excess power will allow a rate An increase in altitude also will increase the power required
of climb specific to the magnitude of excess power. and decrease the power available. Therefore, the climb
performance of an aircraft diminishes with altitude. The
Power available jet speeds for maximum rate of climb, maximum angle of climb,
Power available prope and maximum and minimum level flight airspeeds vary with
altitude. As altitude is increased, these various speeds finally
Speed for max. rate of climb converge at the absolute ceiling of the aircraft. At the absolute
ceiling, there is no excess of power and only one speed will
Speed allow steady, level flight. Consequently, the absolute ceiling
of an aircraft produces zero rate of climb. The service ceiling
Figure 10-8. Power versus climb rate. is the altitude at which the aircraft is unable to climb at a rate
During a steady climb, the rate of climb will depend on excess greater than 100 feet per minute (fpm). Usually, these specific
power while the angle of climb is a function of excess thrust. performance reference points are provided for the aircraft at
a specific design configuration. [Figure 10-9]
The climb performance of an aircraft is affected by certain
variables. The conditions of the aircraft’s maximum climb In discussing performance, it frequently is convenient to use
angle or maximum climb rate occur at specific speeds, the terms power loading, wing loading, blade loading, and disk
and variations in speed will produce variations in climb loading. Power loading is expressed in pounds per horsepower
performance. There is sufficient latitude in most aircraft that and is obtained by dividing the total weight of the aircraft by
small variations in speed from the optimum do not produce the rated horsepower of the engine. It is a significant factor
large changes in climb performance, and certain operational in an aircraft’s takeoff and climb capabilities. Wing loading
considerations may require speeds slightly different from 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.

10-7

specific endurance = flight hours
pounds of fuel

Standard altitude (feet) 24,000 Absolute ceiling or
22,000 Service ceiling specific endurance = flight hours/hour
20,000
18,000 Best rate pounds of fuel/hour
16,000 of climb (Vy)
14,000 or
12,000
10,000 specific endurance = 1

8,000 Best angle fuel flow
6,000 of climb (Vx)
4,000
2,000 Fuel flow can be defined in either pounds or gallons. If
maximum endurance is desired, the flight condition must
Sea level70 80 90 100 110 120 provide a minimum fuel flow. In Figure 10-10 at point A
Indicated airspeed (knots) the airspeed is low and fuel flow is high. This would occur
during ground operations or when taking off and climbing.
Figure 10-9. Absolute and service ceiling. As airspeed is increased, power requirements decrease due
to aerodynamic factors and fuel flow decreases to point B.
Range Performance This is the point of maximum endurance. Beyond this point
The ability of an aircraft to convert fuel energy into flying increases in airspeed come at a cost. Airspeed increases
distance is one of the most important items of aircraft require additional power and fuel flow increases with
performance. In flying operations, the problem of efficient additional power.
range operation of an aircraft appears in two general forms:
Cruise flight operations for maximum range should be
1. To extract the maximum flying distance from a given conducted so that the aircraft obtains maximum specific range
fuel load throughout the flight. The specific range can be defined by
the following relationship.
2. To fly a specified distance with a minimum expenditure
of fuel specific range = NM
pounds of fuel

A common element for each of these operating problems is the or
specific range; that is, nautical miles (NM) of flying distance
versus the amount of fuel consumed. Range must be clearly specific range = NM/hour
distinguished from the item of endurance. Range involves pounds of fuel/hour
consideration of flying distance, while endurance involves
consideration of flying time. Thus, it is appropriate to define or
a separate term, specific endurance.

Fuel flow/Power required (HP) Reference line
At altitude
Maximum endurance at Applicable for a particular
minimum power required Weight
Altitude
A Configuration

B Maximum range at L/DMAX

Speed

Figure 10-10. Airspeed for maximum endurance.
10-8

specific range = knots flight and on performance charts power can be substituted
for fuel flow. This fact allows for the determination of range
fuel flow through analysis of power required versus speed.

If maximum specific range is desired, the flight condition must The maximum endurance condition would be obtained at the
point of minimum power required since this would require the
provide a maximum of speed per fuel flow. While the peak value lowest fuel flow to keep the airplane in steady, level flight.
Maximum range condition would occur where the ratio of
of specific range would provide maximum range operation, speed to power required is greatest. [Figure 10-10]

long-range cruise operation is generally recommended at some The maximum range condition is obtained at maximum lift/
drag ratio (L/DMAX), and it is important to note that for a given
slightly higher airspeed. Most long-range cruise operations are aircraft configuration, the L/DMAX occurs at a particular AOA
and lift coefficient, and is unaffected by weight or altitude.
conducted at the flight condition that provides 99 percent of A variation in weight will alter the values of airspeed and
power required to obtain the L/DMAX. [Figure 10-11]
the absolute maximum specific range. The advantage of such
L/DMAX
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 Power required
aircraft. Lower BwaeisigchtweHiigghhter weight

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 Constant altitude
to maintain the recommended long-range cruise condition
throughout the flight. Since fuel is consumed during cruise, the Speed
gross weight of the aircraft will vary and optimum airspeed,
altitude, and power setting can also vary. Cruise control means Figure 10-11. Effect of weight.
the control of the optimum airspeed, altitude, and power setting
to maintain the 99 percent maximum specific range condition. The variations of speed and power required must be monitored
At the beginning of cruise flight, the relatively high initial by the pilot as part of the cruise control procedure to maintain
weight of the aircraft will require specific values of airspeed, the L/DMAX. When the aircraft’s fuel weight is a small part of
altitude, and power setting to produce the recommended cruise the gross weight and the aircraft’s range is small, the cruise
condition. As fuel is consumed and the aircraft’s gross weight control procedure can be simplified to essentially maintaining
decreases, the optimum airspeed and power setting may a constant speed and power setting throughout the time of
decrease, or, the optimum altitude may increase. In addition, cruise flight. However, a long-range aircraft has a fuel weight
the optimum specific range will increase. Therefore, the pilot that is a considerable part of the gross weight, and cruise
must provide the proper cruise control procedure to ensure control procedures must employ scheduled airspeed and
that optimum conditions are maintained. power changes to maintain optimum range conditions.

Total range is dependent on both fuel available and specific The variations of speed and power required must be
range. When range and economy of operation are the principal monitored by the pilot as part of the cruise control procedure
goals, the pilot must ensure that the aircraft is operated at the to maintain the L/DMAX. When the aircraft’s fuel weight is a
recommended long-range cruise condition. By this procedure, small part of the gross weight and the aircraft’s range is small,
the aircraft will be capable of its maximum design-operating the cruise control procedure can be simplified to essentially
radius, or can achieve flight distances less than the maximum maintaining a constant speed and power setting throughout
with a maximum of fuel reserve at the destination. the time of cruise flight. However, a long-range aircraft has a
fuel weight that is a considerable part of the gross weight, and
A propeller-driven aircraft combines the propeller with the cruise control procedures must employ scheduled airspeed
reciprocating engine for propulsive power. Fuel flow is and power changes to maintain optimum range conditions.
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

10-9

The effect of altitude on the range of a propeller-driven aircraft Region of Reversed Command
is illustrated in Figure 10-12. A flight conducted at high altitude The aerodynamic properties of an aircraft generally determine
has a greater true airspeed (TAS), and the power required is the power requirements at various conditions of flight, while
proportionately greater than when conducted at sea level. The the powerplant capabilities generally determine the power
drag of the aircraft at altitude is the same as the drag at sea level, available at various conditions of flight. When an aircraft
but the higher TAS causes a proportionately greater power is in steady, level flight, a condition of equilibrium must
required. NOTE: The straight line that is tangent to the sea level prevail. An unaccelerated condition of flight is achieved
power curve is also tangent to the altitude power curve. when lift equals weight, and the powerplant is set for thrust
equal to drag. The power required to achieve equilibrium in
Power required constant-altitude flight at various airspeeds is depicted on a
power required curve. The power required curve illustrates
Sea level the fact that at low airspeeds near the stall or minimum
At altitude controllable airspeed, the power setting required for steady,
level flight is quite high.

L/DMAX Constant weight 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.

Speed Flight in the region of reversed command means flight in
which a higher airspeed requires a lower power setting and
Figure 10-12. Effect of altitude on range. a lower airspeed requires a higher power setting to hold
altitude. It does not imply that a decrease in power will
The effect of altitude on specific range also can be appreciated produce lower airspeed. The region of reversed command is
from the previous relationships. If a change in altitude causes encountered in the low speed phases of flight. Flight speeds
identical changes in speed and power required, the proportion below the speed for maximum endurance (lowest point
of speed to power required would be unchanged. The fact on the power curve) require higher power settings with a
implies that the specific range of a propeller-driven aircraft decrease in airspeed. Since the need to increase the required
would be unaffected by altitude. Actually, this is true to the power setting with decreased speed is contrary to the normal
extent that specific fuel consumption and propeller efficiency command of flight, the regime of flight speeds between the
are the principal factors that could cause a variation of speed for minimum required power setting and the stall speed
specific range with altitude. If compressibility effects are (or minimum control speed) is termed the region of reversed
negligible, any variation of specific range with altitude is command. In the region of reversed command, a decrease in
strictly a function of engine/propeller performance. airspeed must be accompanied by an increased power setting
in order to maintain steady flight.
An aircraft equipped with a reciprocating engine will
experience very little, if any, variation of specific range up Figure 10-13 shows the maximum power available as a
to its absolute altitude. There is negligible variation of brake curved line. Lower power settings, such as cruise power,
specific fuel consumption for values of brake horsepower would also appear in a similar curve. The lowest point on
below the maximum cruise power rating of the engine that the power required curve represents the speed at which the
is the lean range of engine operation. Thus, an increase in lowest brake horsepower will sustain level flight. This is
altitude will produce a decrease in specific range only when termed the best endurance airspeed.
the increased power requirement exceeds the maximum cruise
power rating of the engine. One advantage of supercharging An airplane performing a low airspeed, high pitch attitude
is that the cruise power may be maintained at high altitude, power approach for a short-field landing is an example
and the aircraft may achieve the range at high altitude with of operating in the region of reversed command. If an
the corresponding increase in TAS. The principal differences unacceptably high sink rate should develop, it may be
in the high altitude cruise and low altitude cruise are the TAS possible for the pilot to reduce or stop the descent by applying
and climb fuel requirements. power. But without further use of power, the airplane would
probably stall or be incapable of flaring for the landing.

10-10

Region of Maximum power available Takeoff and Landing Performance
reversed
command Excess power The majority of pilot-caused aircraft accidents occur during
the takeoff and landing phase of flight. Because of this fact,
Power the pilot must be familiar with all the variables that influence
Power setting the takeoff and landing performance of an aircraft and must
required strive for exacting, professional procedures of operation
during these phases of flight.
Best endurance speed
Takeoff and landing performance is a condition of accelerated
Speed and decelerated motion. For instance, during takeoff, an
aircraft starts at zero speed and accelerates to the takeoff
Figure 10-13. Power required curve. speed to become airborne. During landing, the aircraft touches
down at the landing speed and decelerates to zero speed. The
Merely lowering the nose of the airplane to regain flying important factors of takeoff or landing performance are:
speed in this situation, without the use of power, would result
in a rapid sink rate and corresponding loss of altitude. • The takeoff or landing speed is generally a function
of the stall speed or minimum flying speed.
If during a soft-field takeoff and climb, for example, the pilot
attempts to climb out of ground effect without first attaining • The rate of acceleration/deceleration during the
normal climb pitch attitude and airspeed, the airplane may takeoff or landing roll. The speed (acceleration and
inadvertently enter the region of reversed command at a deceleration) experienced by any object varies directly
dangerously low altitude. Even with full power, the airplane with the imbalance of force and inversely with the
may be incapable of climbing or even maintaining altitude. mass of the object. An airplane on the runway moving
The pilot’s only recourse in this situation is to lower the pitch at 75 knots has four times the energy it has traveling
attitude in order to increase airspeed, which will inevitably at 37 knots. Thus, an airplane requires four times as
result in a loss of altitude. much distance to stop as required at half the speed.

Airplane pilots must give particular attention to precise • The takeoff or landing roll distance is a function of
control of airspeed when operating in the low flight speeds both acceleration/deceleration and speed.
of the region of reversed command.
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]

Figure 10-14. Takeoff distance chart.

10-11

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

Runway surfaces vary widely from one airport to another. Ensure that runways are adequate in length for takeoff
The runway surface encountered may be concrete, asphalt, acceleration and landing deceleration when less than ideal
gravel, dirt, or grass. The runway surface for a specific airport surface conditions are being reported.
is noted in the Airport/Facility Directory (A/FD). Any surface
that is not hard and smooth will increase the ground roll The gradient or slope of the runway is the amount of change
during takeoff. This is due to the inability of the tires to roll in runway height over the length of the runway. The gradient
smoothly along the runway. Tires can sink into soft, grassy, is expressed as a percentage such as a 3 percent gradient. This
or muddy runways. Potholes or other ruts in the pavement means that for every 100 feet of runway length, the runway
can be the cause of poor tire movement along the runway. height changes by 3 feet. A positive gradient indicates the
Obstructions such as mud, snow, or standing water reduce the runway height increases, and a negative gradient indicates the
airplane’s acceleration down the runway. Although muddy runway decreases in height. An upsloping runway impedes
and wet surface conditions can reduce friction between acceleration and results in a longer ground run during takeoff.
the runway and the tires, they can also act as obstructions However, landing on an upsloping runway typically reduces
and reduce the landing distance. [Figure 10-15] Braking the landing roll. A downsloping runway aids in acceleration
effectiveness is another consideration when dealing with on takeoff resulting in shorter takeoff distances. The opposite
various runway types. The condition of the surface affects is true when landing, as landing on a downsloping runway
the braking ability of the airplane. increases landing distances. Runway slope information is
contained in the A/FD. [Figure 10-16]
The amount of power that is applied to the brakes without
skidding the tires is referred to as braking effectiveness.

Airport name Runway surface
Runway slope and direction of slope

Runway

Figure 10-16. Airport/facility directory (A/FD) information.
10-12

Water on the Runway and Dynamic Hydroplaning braking may be ineffective, so aerodynamic braking should
Water on the runways reduces the friction between the tires be used to its fullest advantage.
and the ground, and can reduce braking effectiveness. The
ability to brake can be completely lost when the tires are Takeoff Performance
hydroplaning because a layer of water separates the tires from The minimum takeoff distance is of primary interest in
the runway surface. This is also true of braking effectiveness the operation of any aircraft because it defines the runway
when runways are covered in ice. requirements. The minimum takeoff distance is obtained by
taking off at some minimum safe speed that allows sufficient
When the runway is wet, the pilot may be confronted with margin above stall and provides satisfactory control and
dynamic hydroplaning. Dynamic hydroplaning is a condition initial rate of climb. Generally, the lift-off speed is some fixed
in which the aircraft tires ride on a thin sheet of water rather percentage of the stall speed or minimum control speed for the
than on the runway’s surface. Because hydroplaning wheels aircraft in the takeoff configuration. As such, the lift-off will
are not touching the runway, braking and directional control be accomplished at some particular value of lift coefficient
are almost nil. To help minimize dynamic hydroplaning, and AOA. Depending on the aircraft characteristics, the lift-
some runways are grooved to help drain off water; most off speed will be anywhere from 1.05 to 1.25 times the stall
runways are not. speed or minimum control speed.

Tire pressure is a factor in dynamic hydroplaning. Using To obtain minimum takeoff distance at the specific lift-off
speed, the forces that act on the aircraft must provide the
the simple formula in Figure 10-17, a pilot can calculate maximum acceleration during the takeoff roll. The various
forces acting on the aircraft may or may not be under the
the minimum speed, in knots, at which hydroplaning will control of the pilot, and various procedures may be necessary
in certain aircraft to maintain takeoff acceleration at the
begin. In plain language, the minimum hydroplaning speed highest value.

is determined by multiplying the square root of the main gear The powerplant thrust is the principal force to provide the
acceleration and, for minimum takeoff distance, the output
tire pressure in psi by nine. For example, if the main gear tire thrust should be at a maximum. Lift and drag are produced
as soon as the aircraft has speed, and the values of lift and
pressure is at 36 psi, the aircraft would begin hydroplaning drag depend on the AOA and dynamic pressure.

at 54 knots. Tire Pressure

Minimum dynamic hydroplaning speed (rounded off) =

9 x Tire pressure (in psi) 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.

36 = 6 For example, the effect of gross weight on takeoff distance
9 x 6 = 54 knots is significant and proper consideration of this item must be
made in predicting the aircraft’s takeoff distance. Increased
Figure 10-17. Tire pressure. gross weight can be considered to produce a threefold effect
on takeoff performance:
Landing at higher than recommended touchdown speeds will
expose the aircraft to a greater potential for hydroplaning. 1. Higher lift-off speed
And once hydroplaning starts, it can continue well below the
minimum initial hydroplaning speed. 2. Greater mass to accelerate

On wet runways, directional control can be maximized 3. Increased retarding force (drag and ground friction)
by landing into the wind. Abrupt control inputs should be
avoided. When the runway is wet, anticipate braking problems If the gross weight increases, a greater speed is necessary to
well before landing and be prepared for hydroplaning. Opt for produce the greater lift necessary to get the aircraft airborne
a suitable runway most aligned with the wind. Mechanical 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.

10-13

A change in gross weight will change the net accelerating Percent increase 80 Reference line
force and change the mass that is being accelerated. If the in takeoff or 70
aircraft has a relatively high thrust-to-weight ratio, the change 60
in the net accelerating force is slight and the principal effect landing distance 50
on acceleration is due to the change in mass. 40
Ratio of wind 30
For example, a 10 percent increase in takeoff gross weight velocity to takeoff 20
would cause: or landing speed 10

• A 5 percent increase in takeoff velocity. 30% 20% 10% Tailwind

• At least a 9 percent decrease in rate of acceleration. Headwind 10 10% 20% 30%
20 Ratio of wind
• At least a 21 percent increase in takeoff distance. 30
40 velocity to takeoff
With ISA conditions, increasing the takeoff weight of the 50 or landing speed
average Cessna 182 from 2,400 pounds to 2,700 pounds (11 60
percent increase) results in an increased takeoff distance from Percent decrease
440 feet to 575 feet (23 percent increase). in takeoff or

For the aircraft with a high thrust-to-weight ratio, the landing distance
increase in takeoff distance might be approximately 21 to
22 percent, but for the aircraft with a relatively low thrust- Figure 10-18. Effect of wind on takeoff and landing.
to-weight ratio, the increase in takeoff distance would be
approximately 25 to 30 percent. Such a powerful effect control, or have a very low initial rate of climb. In some cases,
requires proper consideration of gross weight in predicting an excessive AOA may not allow the aircraft to climb out
takeoff distance. of ground effect. On the other hand, an excessive airspeed
at takeoff may improve the initial rate of climb and “feel”
The effect of wind on takeoff distance is large, and proper of the aircraft, but will produce an undesirable increase in
consideration also must be provided when predicting takeoff takeoff distance. Assuming that the acceleration is essentially
distance. The effect of a headwind is to allow the aircraft unaffected, the takeoff distance varies with the square of the
to reach the lift-off speed at a lower groundspeed while the takeoff velocity.
effect of a tailwind is to require the aircraft to achieve a
greater groundspeed to attain the lift-off speed. Thus, ten percent excess airspeed would increase the takeoff
distance 21 percent. In most critical takeoff conditions, such
A headwind that is 10 percent of the takeoff airspeed will an increase in takeoff distance would be prohibitive, and the
reduce the takeoff distance approximately 19 percent. pilot must adhere to the recommended takeoff speeds.
However, a tailwind that is 10 percent of the takeoff airspeed
will increase the takeoff distance approximately 21 percent. In The effect of pressure altitude and ambient temperature
the case where the headwind speed is 50 percent of the takeoff is to define the density altitude and its effect on takeoff
speed, the takeoff distance would be approximately 25 percent performance. While subsequent corrections are appropriate
of the zero wind takeoff distance (75 percent reduction). for the effect of temperature on certain items of powerplant
performance, density altitude defines specific effects on
The effect of wind on landing distance is identical to its takeoff performance. An increase in density altitude can
effect on takeoff distance. Figure 10-18 illustrates the general produce a twofold effect on takeoff performance:
effect of wind by the percent change in takeoff or landing
distance as a function of the ratio of wind velocity to takeoff 1. Greater takeoff speed
or landing speed.
2. Decreased thrust and reduced net accelerating force
The effect of proper takeoff speed is especially important
when runway lengths and takeoff distances are critical. The If an aircraft of given weight and configuration is operated at
takeoff speeds specified in the AFM/POH are generally greater heights above standard sea level, the aircraft requires
the minimum safe speeds at which the aircraft can become the same dynamic pressure to become airborne at the takeoff
airborne. Any attempt to take off below the recommended lift coefficient. Thus, the aircraft at altitude will take off at the
speed means that the aircraft could stall, be difficult to same indicated airspeed (IAS) as at sea level, but because of
the reduced air density, the TAS will be greater.

10-14

The effect of density altitude on powerplant thrust depends aircraft during the landing roll may require various procedures
much on the type of powerplant. An increase in altitude to maintain landing deceleration at the peak value.
above standard sea level will bring an immediate decrease in
power output for the unsupercharged reciprocating engine. A distinction should be made between the procedures for
However, an increase in altitude above standard sea level will minimum landing distance and an ordinary landing roll with
not cause a decrease in power output for the supercharged considerable excess runway available. Minimum landing
reciprocating engine until the altitude exceeds the critical distance will be obtained by creating a continuous peak
operating altitude. For those powerplants that experience deceleration of the aircraft; that is, extensive use of the brakes
a decay in thrust with an increase in altitude, the effect for maximum deceleration. On the other hand, an ordinary
on the net accelerating force and acceleration rate can be landing roll with considerable excess runway may allow
approximated by assuming a direct variation with density. extensive use of aerodynamic drag to minimize wear and tear
Actually, this assumed variation would closely approximate on the tires and brakes. If aerodynamic drag is sufficient to
the effect on aircraft with high thrust-to-weight ratios. cause deceleration, it can be used in deference to the brakes
in the early stages of the landing roll; i.e., brakes and tires
Proper accounting of pressure altitude and temperature is suffer from continuous hard use, but aircraft aerodynamic
mandatory for accurate prediction of takeoff roll distance. drag is free and does not wear out with use. The use of
The most critical conditions of takeoff performance are the aerodynamic drag is applicable only for deceleration to 60
result of some combination of high gross weight, altitude, or 70 percent of the touchdown speed. At speeds less than
temperature, and unfavorable wind. In all cases, the pilot 60 to 70 percent of the touchdown speed, aerodynamic drag
must make an accurate prediction of takeoff distance from is so slight as to be of little use, and braking must be utilized
the performance data of the AFM/POH, regardless of the to produce continued deceleration. Since the objective during
runway available, and strive for a polished, professional the landing roll is to decelerate, the powerplant thrust should
takeoff procedure. be the smallest possible positive value (or largest possible
negative value in the case of thrust reversers).
In the prediction of takeoff distance from the AFM/POH data,
the following primary considerations must be given: In addition to the important factors of proper procedures,
many other variables affect the landing performance. Any
• Pressure altitude and temperature—to define the effect item that alters the landing speed or deceleration rate during
of density altitude on distance the landing roll will affect the landing distance.

• Gross weight—a large effect on distance The effect of gross weight on landing distance is one of the
principal items determining the landing distance. One effect
• Wind—a large effect due to the wind or wind of an increased gross weight is that a greater speed will be
component along the runway required to support the aircraft at the landing AOA and lift
coefficient. For an example of the effect of a change in gross
• Runway slope and condition—the effect of an incline weight, a 21 percent increase in landing weight will require
and retarding effect of factors such as snow or ice a ten percent increase in landing speed to support the greater
weight.
Landing Performance
In many cases, the landing distance of an aircraft will define When minimum landing distances are considered, braking
the runway requirements for flight operations. The minimum friction forces predominate during the landing roll and, for
landing distance is obtained by landing at some minimum the majority of aircraft configurations, braking friction is the
safe speed, which allows sufficient margin above stall and main source of deceleration.
provides satisfactory control and capability for a go-around.
Generally, the landing speed is some fixed percentage of The minimum landing distance will vary in direct proportion
the stall speed or minimum control speed for the aircraft to the gross weight. For example, a ten percent increase in
in the landing configuration. As such, the landing will be gross weight at landing would cause a:
accomplished at some particular value of lift coefficient
and AOA. The exact values will depend on the aircraft • Five percent increase in landing velocity
characteristics but, once defined, the values are independent
of weight, altitude, and wind. • Ten percent increase in landing distance

To obtain minimum landing distance at the specified landing A contingency of this is the relationship between weight and
speed, the forces that act on the aircraft must provide maximum braking friction force.
deceleration during the landing roll. The forces acting on the

10-15

The effect of wind on landing distance is large and deserves greatest required landing distances and critical levels of
proper consideration when predicting landing distance. Since energy dissipation required of the brakes. In all cases, it is
the aircraft will land at a particular airspeed independent of necessary to make an accurate prediction of minimum landing
the wind, the principal effect of wind on landing distance is distance to compare with the available runway. A polished,
the change in the groundspeed at which the aircraft touches professional landing procedure is necessary because the
down. The effect of wind on deceleration during the landing landing phase of flight accounts for more pilot-caused aircraft
is identical to the effect on acceleration during the takeoff. accidents than any other single phase of flight.

The effect of pressure altitude and ambient temperature is to In the prediction of minimum landing distance from the AFM/
define density altitude and its effect on landing performance. POH data, the following considerations must be given:
An increase in density altitude increases the landing speed
but does not alter the net retarding force. Thus, the aircraft • Pressure altitude and temperature—to define the effect
at altitude lands at the same IAS as at sea level but, because of density altitude
of the reduced density, the TAS is greater. Since the aircraft
lands at altitude with the same weight and dynamic pressure, • Gross weight—which defines the CAS for landing.
the drag and braking friction throughout the landing roll have
the same values as at sea level. As long as the condition is • Wind—a large effect due to wind or wind component
within the capability of the brakes, the net retarding force along the runway
is unchanged, and the deceleration is the same as with the
landing at sea level. Since an increase in altitude does not alter • Runway slope and condition—relatively small
deceleration, the effect of density altitude on landing distance correction for ordinary values of runway slope, but a
is due to the greater TAS. significant effect of snow, ice, or soft ground

The minimum landing distance at 5,000 feet is 16 percent A tail wind of ten knots increases the landing distance by
greater than the minimum landing distance at sea level. The about 21 percent. An increase of landing speed by ten percent
approximate increase in landing distance with altitude is increases the landing distance by 20 percent. Hydroplaning
approximately three and one-half percent for each 1,000 feet makes braking ineffective until a decrease of speed to that
of altitude. Proper accounting of density altitude is necessary determined using Figure 10-17.
to accurately predict landing distance.
For instance, a pilot is downwind for runway 18, and the
The effect of proper landing speed is important when runway tower asks if runway 27 could be accepted. There is a light
lengths and landing distances are critical. The landing speeds rain and the winds are out of the east at ten knots. The pilot
specified in the AFM/POH are generally the minimum safe accepts because he or she is approaching the extended
speeds at which the aircraft can be landed. Any attempt to land centerline of runway 27. The turn is tight and the pilot must
at below the specified speed may mean that the aircraft may descend (dive) to get to runway 27. After becoming aligned
stall, be difficult to control, or develop high rates of descent. with the runway and at 50 feet AGL, the pilot is already 1,000
On the other hand, an excessive speed at landing may improve feet down the 3,500 feet runway. The airspeed is still high by
the controllability slightly (especially in crosswinds), but about ten percent (should be at 70 knots and is at about 80
causes an undesirable increase in landing distance. knots). The wind of ten knots is blowing from behind.

A ten percent excess landing speed causes at least a 21 percent First, the airspeed being high by about ten percent (80 knots
increase in landing distance. The excess speed places a greater versus 70 knots), as presented in the performance chapter,
working load on the brakes because of the additional kinetic results in a 20 percent increase in the landing distance.
energy to be dissipated. Also, the additional speed causes In performance planning, the pilot determined that at 70
increased drag and lift in the normal ground attitude, and the knots the distance would be 1,600 feet. However, now it
increased lift reduces the normal force on the braking surfaces. is increased by 20 percent and the required distance is now
The deceleration during this range of speed immediately after 1,920 feet.
touchdown may suffer, and it is more probable for a tire to be
blown out from braking at this point. The newly revised landing distance of 1,920 feet is also
affected by the wind. In looking at Figure 10-18, the affect
The most critical conditions of landing performance are of the wind is an additional 20 percent for every ten miles
combinations of high gross weight, high density altitude, per hour (mph) in wind. This is computed not on the original
and unfavorable wind. These conditions produce the estimate but on the estimate based upon the increased
airspeed. Now the landing distance is increased by another

10-16

320 feet for a total requirement of 2,240 feet to land the VS1—the calibrated power-off stalling speed or the minimum
airplane after reaching 50 feet AGL. steady flight speed at which the aircraft is controllable in a
specified configuration.
That is the original estimate of 1,600 under planned
conditions plus the additional 640 feet for excess speed VY—the speed at which the aircraft will obtain the maximum
and the tailwind. Given the pilot overshot the threshhold increase in altitude per unit of time. This best rate-of-climb
by 1,000 feet, the total length required is 3, 240 on a 3,500 speed normally decreases slightly with altitude.
foot runway; 260 feet to spare. But this is in a perfect
environment. Most pilots become fearful as the end of the VX—the speed at which the aircraft will obtain the highest
runway is facing them just ahead. A typical pilot reaction altitude in a given horizontal distance. This best angle-of-
is to brake—and brake hard. Because the aircraft does not climb speed normally increases slightly with altitude.
have antilock braking features like a car, the brakes lock,
and the aircraft hydroplanes on the wet surface of the runway VLE—the maximum speed at which the aircraft can be safely
until decreasing to a speed of about 54 knots (the square flown with the landing gear extended. This is a problem
root of the tire pressure (√36) x 9). Braking is ineffective involving stability and controllability.
when hydroplaning.
VLO—the maximum speed at which the landing gear can
The 260 feet that a pilot might feel is left over has long since be safely extended or retracted. This is a problem involving
evaporated as the aircraft hydroplaned the first 300–500 feet the air loads imposed on the operating mechanism during
when the brakes locked. This is an example of a true story, extension or retraction of the gear.
but one which only changes from year to year because of new
participants and aircraft with different N-numbers. VFE—the highest speed permissible with the wing flaps in a
prescribed extended position. This is because of the air loads
In this example, the pilot actually made many bad decisions. imposed on the structure of the flaps.
Bad decisions, when combined, have a synergy greater
than the individual errors. Therefore, the corrective actions VA—the calibrated design maneuvering airspeed. This is
become larger and larger until correction is almost impossible. the maximum speed at which the limit load can be imposed
Aeronautical decision-making will be discussed more fully in (either by gusts or full deflection of the control surfaces)
Chapter 17, Aeronautical Decision-Making (ADM). without causing structural damage. Operating at or below
manuevering speed does not provide structural protection
Performance Speeds against multiple full control inputs in one axis or full control
inputs in more than one axis at the same time.
True Airspeed (TAS)—the speed of the aircraft in relation to
the air mass in which it is flying. VNO—the maximum speed for normal operation or the
maximum structural cruising speed. This is the speed at
Indicated Airspeed (IAS)—the speed of the aircraft as which exceeding the limit load factor may cause permanent
observed on the ASI. It is the airspeed without correction deformation of the aircraft structure.
for indicator, position (or installation), or compressibility
errors. VNE—the speed which should never be exceeded. If flight is
attempted above this speed, structural damage or structural
Calibrated Airspeed (CAS)—the ASI reading corrected for failure may result.
position (or installation), and instrument errors. (CAS is
equal to TAS at sea level in standard atmosphere.) The color Performance Charts
coding for various design speeds marked on ASIs may be
IAS or CAS. Performance charts allow a pilot to predict the takeoff, climb,
cruise, and landing performance of an aircraft. These charts,
Equivalent Airspeed (EAS)—the ASI reading corrected provided by the manufacturer, are included in the AFM/POH.
for position (or installation), or instrument error, and for Information the manufacturer provides on these charts has
adiabatic compressible flow for the particular altitude. (EAS been gathered from test flights conducted in a new aircraft,
is equal to CAS at sea level in standard atmosphere.) under normal operating conditions while using average
piloting skills, and with the aircraft and engine in good
VS0—the calibrated power-off stalling speed or the minimum working order. Engineers record the flight data and create
steady flight speed at which the aircraft is controllable in the performance charts based on the behavior of the aircraft
landing configuration. during the test flights. By using these performance charts,

10-17

a pilot can determine the runway length needed to take off The remainder of this section covers performance information
and land, the amount of fuel to be used during flight, and the for aircraft in general and discusses what information the
time required to arrive at the destination. It is important to charts contain and how to extract information from the charts
remember that the data from the charts will not be accurate by direct reading and interpolation methods. Every chart
if the aircraft is not in good working order or when operating contains a wealth of information that should be used when
under adverse conditions. Always consider the necessity to flight planning. Examples of the table, graph, and combined
compensate for the performance numbers if the aircraft is not graph formats for all aspects of flight will be discussed.
in good working order or piloting skills are below average.
Each aircraft performs differently and, therefore, has different Interpolation
performance numbers. Compute the performance of the Not all of the information on the charts is easily extracted.
aircraft prior to every flight, as every flight is different. (See Some charts require interpolation to find the information for
appendix for examples of performance charts for a Cessna specific flight conditions. Interpolating information means
Model 172R and Challenger 605.) that by taking the known information, a pilot can compute
intermediate information. However, pilots sometimes round
Every chart is based on certain conditions and contains off values from charts to a more conservative figure.
notes on how to adapt the information for flight conditions.
It is important to read every chart and understand how to Using values that reflect slightly more adverse conditions
use it. Read the instructions provided by the manufacturer. provides a reasonable estimate of performance information
For an explanation on how to use the charts, refer to the and gives a slight margin of safety. The following illustration
example provided by the manufacturer for that specific chart. is an example of interpolating information from a takeoff
[Figure 10-19] distance chart. [Figure 10-20]

The information manufacturers furnish is not standardized. Density Altitude Charts
Information may be contained in a table format, and other Use a density altitude chart to figure the density altitude at the
information may be contained in a graph format. Sometimes departing airport. Using Figure 10-21, determine the density
combined graphs incorporate two or more graphs into one chart altitude based on the given information.
to compensate for multiple conditions of flight. Combined
graphs allow the pilot to predict aircraft performance for Sample Problem 1
variations in density altitude, weight, and winds all on one Airport Elevation...............................................5,883 feet
chart. Because of the vast amount of information that can be OAT...........................................................................70 °F
extracted from this type of chart, it is important to be very Altimeter...........................................................30.10" Hg
accurate in reading the chart. A small error in the beginning
can lead to a large error at the end.

, ,

,

Figure 10-19. Conditions notes chart.
10-18

Conditions Flaps 10° TAKEOFF DISTANCE
Full throttle prior to brake release MAXIMUM WEIGHT 2,400 LB
Paved level runway
Zero wind

Takeoff 0 °C 10 °C 20 °C 30 °C 40 °C
speed KIAS Grnd Total feet
Weight Lift AT Press Grnd Total feet Grnd Total feet Grnd Total feet Roll to clear Grnd Total feet
(lb) off 50 ft ALT Roll to clear Roll to clear Roll to clear (feet) 50 ft OBS Roll to clear
(feet) (feet) 50 ft OBS (feet) 50 ft OBS (feet) 50 ft OBS (feet) 50 ft OBS
1,065 1,945
2,400 51 56 S.L. 795 1,460 860 1,570 925 1,685 995 1,810 1,170 2,155
1,000 875 1,605 940 1,725 1,015 1,860 1,090 2,000 1,290 2,395
2,000 960 1,770 1,035 1,910 1,115 2,060 1,200 2,220 1,425 2,685
3,000 1,055 1,960 1,140 2,120 1,230 2,295 1,325 2,480 1,575 3,030
4,000 1,165 2,185 1,260 2,365 1,355 2,570 1,465 2,790 1,745 3,455
5,000 1,285 2,445 1,390 2,660 1,500 2,895 1,620 3,160 1,940 3,990
6,000 1,425 2,755 1,540 3,015 1,665 3,300 1,800 3,620 ---
7,000 1,580 3,140 1,710 3,450 1,850 3,805 2,000 4,220 --- ---
8,000 1,755 3,615 1,905 4,015 2,060 4,480 - - - - - - ---
Density Altitude Chart

To find the takeoff distance for a pressure altitude of 2,500 feet 15
at 20 °C, average the ground roll for 2,000 feet and 3,000 feet.
13,000 14,000 Altimeter setting
1,115 + 1,230 14 ("Hg)
= 1,173 feet
Pressure altitude
2 conversion factor

Figure 10-20. Interpolating charts. 13

First, compute the pressure altitude conversion. Find 30.10 10,000 11,000 12,000
under the altimeter heading. Read across to the second Pressure altitude (feet)
column. It reads “–165.” Therefore, it is necessary to subtract 28.0 1,824
165 from the airport elevation giving a pressure altitude of 12 28.1 1,727
5,718 feet. Next, locate the outside air temperature on the
scale along the bottom of the graph. From 70°, draw a line up Approximate density altitude (thousand feet) 28.2 1,630
to the 5,718 feet pressure altitude line, which is about two- 11 28.3 1,533
thirds 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 28.4 1,436
and read the approximate density altitude. The approximate 10 28.5 1,340
density altitude in thousands of feet is 7,700 feet. Standard temperature
9,000 28.6 1,244
Takeoff Charts
Takeoff charts are typically provided in several forms and allow 9 28.7 1,148
a pilot to compute the takeoff distance of the aircraft with no
flaps or with a specific flap configuration. A pilot can also 8,000 28.8 1,053
compute distances for a no flap takeoff over a 50 foot obstacle
scenario, as well as with flaps over a 50 foot obstacle. The 8 28.9 957
takeoff distance chart provides for various aircraft weights, 29.0 863
altitudes, temperatures, winds, and obstacle heights. 6,000 7,000
7 29.1 768
Sample Problem 2 29.2 673

Pressure Altitude...............................................2,000 feet 6 29.3 579
29.4 485
OAT..........................................................................22 °C 5,000
29.5 392
Takeoff Weight.............................................2,600 pounds 5 29.6 298
1,000 2,000 3,000 4,000
Headwind...............................................................6 knots 29.7 205Sea level
4 29.8 112
Obstacle Height.......................................50 foot obstacle
29.9 20
Refer to Figure 10-22. This chart is an example of a combined
takeoff distance graph. It takes into consideration pressure 3 29.92 0
altitude, temperature, weight, wind, and obstacles all on one
chart. First, find the correct temperature on the bottom left- 30.0 −73

2 30.1 −165

30.2 −257

1 30.3 −348
30.4 −440
-1,000
SC.L-.18 -12° 30.5 −531
-7° -1° 4° 10° 16° 21° 27° 32° 38° 30.6 −622

F 0° 10° 20° 30° 40° 50° 60° 70° 80° 90°100° 30.7 −712
30.8 −803
Outside air temperature

Figure 10-21. Density altitude chart.

10-19

Reference line 6,000
Reference line
Takeoff speed
Tailwind

Reference line

IntGeuridmeelidnieastenot applicable for
Weight Associated conditions
pounds
Lift-off 50 ft Power Full throttle 2,600 rpm 5,000
2,950 kts MPH kts MPH
2,800 Mixture Lean to appropriate fuel
2,600 66 76 72 83 pressure
2,400 64 74 70 81
2,200 Flaps Up Obstacle heights
63 72 68 78
61 70 66 76 Landing Retract after positive 4,000
58 67 63 73 gear climb established 3,000

Cowl Open
flaps

Pressure altitude - feet ISA Headwind
10,080,0060,0040,0020,000S0.L.
2,000
C -40° -30° -20° -10° 0° 10° 20° 30° 40° 50°
Outside air temperature 1,000

F -40° -20° 0° 20° 40° 60° 80° 100° 120° 2,800 2,600 2,400 2,200 0 10 20 30 0 0
Weight 50
(pounds)
Wind component Obstacle

(knots) height (feet)

Figure 10-22. Takeoff distance graph.

hand side of the graph. Follow the line from 22° C straight left to right across the table. The takeoff speed is in the second
up until it intersects the 2,000 foot altitude line. From that column and, in the third column under pressure altitude, find
point, draw a line straight across to the first dark reference the pressure altitude of 3,000 feet. Carefully follow that line
line. Continue to draw the line from the reference point in to the right until it is under the correct temperature column
a diagonal direction following the surrounding lines until it of 30 °C. The ground roll total reads 1,325 feet and the total
intersects the corresponding weight line. From the intersection required to clear a 50 foot obstacle is 2,480 feet. At this point,
of 2,600 pounds, draw a line straight across until it reaches the there is an 18 knot headwind. According to the notes section
second reference line. Once again, follow the lines in a diagonal under point number two, decrease the distances by ten percent
manner until it reaches the six knot headwind mark. Follow for each 9 knots of headwind. With an 18 knot headwind, it
straight across to the third reference line and from here, draw is necessary to decrease the distance by 20 percent. Multiply
a line in two directions. First, draw a line straight across to 1,325 feet by 20 percent (1,325 x .20 = 265), subtract the
figure the ground roll distance. Next, follow the diagonal lines product from the total distance (1,325 – 265 = 1,060). Repeat
again until it reaches the corresponding obstacle height. In this this process for the total distance over a 50 foot obstacle. The
case, it is a 50 foot obstacle. Therefore, draw the diagonal line ground roll distance is 1,060 feet and the total distance over
to the far edge of the chart. This results in a 600 foot ground a 50 foot obstacle is 1,984 feet.
roll distance and a total distance of 1,200 feet over a 50 foot
obstacle. To find the corresponding takeoff speeds at lift-off Climb and Cruise Charts
and over the 50 foot obstacle, refer to the table on the top of Climb and cruise chart information is based on actual
the chart. In this case, the lift-off speed at 2,600 pounds would flight tests conducted in an aircraft of the same type. This
be 63 knots and over the 50 foot obstacle would be 68 knots. information is extremely useful when planning a cross-
country to predict the performance and fuel consumption of
Sample Problem 3 the aircraft. Manufacturers produce several different charts
for climb and cruise performance. These charts include
Pressure Altitude...............................................3,000 feet everything from fuel, time, and distance to climb, to best
power setting during cruise, to cruise range performance.
OAT.........................................................................30 °C
The first chart to check for climb performance is a fuel,
Takeoff Weight............................................2,400 pounds time, and distance-to-climb chart. This chart will give the
fuel amount used during the climb, the time it will take to
Headwind............................................................18 knots accomplish the climb, and the ground distance that will
be covered during the climb. To use this chart, obtain the
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

10-20

Notes Conditions Flaps 10° TAKEOFF DISTANCE
Full throttle prior to brake release MAXIMUM WEIGHT 2,400 LB
Paved level runway
Zero wind SHORT FIELD

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.

Takeoff 0 °C 10 °C 20 °C 30 °C 40 °C
speed KIAS
Weight Lift AT Press Grnd Total feet Grnd Total feet Grnd Total feet Grnd Total feet Grnd Total feet
(lb) off 50 ft ALT Roll to clear Roll to clear Roll to clear Roll to clear Roll to clear
51 56 (FT) (FT) 50 ft OBS (FT) 50 ft OBS (FT) 50 ft OBS (FT) 50 ft OBS (FT) 50 ft OBS
2,400 795 1,460 860 1,570 925 1,685 995 1,810 1,065 1,945
49 54 S.L. 875 1,605 940 1,725 1,015 1,860 1,090 2,000 1,170 2,155
2,200 1,000 960 1,770 1,035 1,910 1,115 2,060 1,200 2,220 1,290 2,395
46 51 2,000 1,055 1,960 1,140 2,120 1,230 2,295 1,325 2,480 1,425 2,685
2,000 3,000 1,165 2,185 1,260 2,365 1,355 2,570 1,465 2,790 1,575 3,030
4,000 1,285 2,445 1,390 2,660 1,500 2,895 1,620 3,160 1,745 3,455
5,000 1,425 2,755 1,540 3,015 1,665 3,300 1,800 3,620 1,940 3,990
6,000 1,580 3,140 1,710 3,450 1,850 3,805 2,000 4,220 ---
7,000 1,755 3,615 1,905 4,015 2,060 4,480 ��� --- ---
8,000 650 1,195 700 1,280 750 1,375 805 --- 865 ---
S.L. 710 1,310 765 1,405 825 1,510 885 1,470 950 1,575
1,000 780 1,440 840 1,545 905 1,660 975 1,615 1,045 1,735
2,000 855 1,585 925 1,705 995 1,835 1,070 1,785 1,150 1,915
3,000 945 1,750 1,020 1,890 1,100 2,040 1,180 1,975 1,270 2,130
4,000 1,040 1,945 1,125 2,105 1,210 2,275 1,305 2,200 1,405 2,375
5,000 1,150 2,170 1,240 2,355 1,340 2,555 1,445 2,465 1,555 2,665
6,000 1,270 2,440 1,375 2,655 1,485 2,890 1,605 2,775 1,730 3,020
7,000 1,410 2,760 1,525 3,015 1,650 3,305 1,785 3,155 1,925 3,450
8,000 525 565 1,035 605 1,110 650 3,630 695 4,005
S.L. 570 970 615 1,135 665 1,215 710 1,185 765 1,265
1,000 625 1,060 675 1,240 725 1,330 780 1,295 840 1,385
2,000 690 1,160 740 1,365 800 1,465 860 1,425 920 1,525
3,000 755 1,270 815 1,500 880 1,615 945 1,570 1,015 1,685
4,000 830 1,400 900 1,660 970 1,790 2,145 1,735 1,120 1,865
5,000 920 1,545 990 1,845 1,070 1,990 2,405 1,925 1,235 2,070
6,000 1,015 1,710 1,095 2,055 1,180 2,225 2,715 2,145 1,370 2,315
7,000 1,125 1,900 1,215 2,305 1,310 2,500 1,410 2,405 1,520 2,605
8,000 2,125 2,715 2,950

Figure 10-23. Takeoff distance short field charts.

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.

20,000 Pressure ALT - feet Fuel - gallons inutes nautical miles
18,000 Time - m Distance
-
16,000

14,000 Cruise

12,000 Associated conditions
Maximum continuous power*
10,000 3,600 lb gross weight
Flaps up
8,000 90 KIAS
6,000 No wind
* 2,700 rpm & 36 in M.P. (3-blade prop)
4,000
2,575 rpm & 36 in M.P. (2-blade prop)
2,000 Departure
Sea level 50

-40° -30° -20° -10° 0° 10° 20° 30° 40°C 0 10 20 30 40
Outside air temperature Fuel, time and distance to climb

Figure 10-24. Fuel time distance climb chart.

10-21

Sample Problem 4 Conditions Flaps up NORMAL CLIMB
Gear up 110 KIAS
Departing Airport Pressure Altitude.................6,000 feet 2,500 RPM
30" Hg
Departing Airport OAT............................................25 °C 120 PPH fuel flow
Cowl flaps open
Cruise Pressure Altitude..................................10,000 feet Standard temperature

Cruise OAT..............................................................10 °C Notes 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
First, find the information for the departing airport. Find the
OAT for the departing airport along the bottom, left-hand side temperature.
of the graph. Follow the line from 25 °C straight up until it 3. Distances shown are based on zero wind.
intersects the line corresponding to the pressure altitude of
6,000 feet. Continue this line straight across until it intersects Press Rate of From sea level
all three lines for fuel, time, and distance. Draw a line straight ALT climb
down from the intersection of altitude and fuel, altitude and Weight (feet) FPM Time Fuel used Distance
time, and a third line at altitude and distance. It should read (pounds) (minutes) (pounds) (nautical
three and one-half gallons of fuel, 6.5 minutes of time, and S.L. 605
nine NM. Next, repeat the steps to find the information for 4,000 4,000 570 0 0 miles)
the cruise altitude. It should read six and one-half gallons 3,700 8,000 530 7 14 0
of fuel, 11.5 minutes of time, and 15 NM. Take each set of 12,000 485 14 28
numbers for fuel, time, and distance and subtract them from 3,400 16,000 430 22 44 13
one another (6.5 – 3.5 = 3 gallons of fuel). It will take three 20,000 365 31 62 27
gallons of fuel and 5 minutes of time to climb to 10,000 700 41 82 43
feet. During that climb, the distance covered is six NM. S.L. 665 0 63
Remember, according to the notes at the top of the chart, these 4,000 625 6 87
numbers do not take into account wind, and it is assumed 8,000 580 12
maximum continuous power is being used. 12,000 525 19 00
16,000 460 26 12 11
The next example is a fuel, time, and distance-to-climb table. 20,000 810 34 24 23
For this table, use the same basic criteria as for the previous 775 0 37 37
chart. However, it is necessary to figure the information in a S.L. 735 5 52 53
different manner. Refer to Figure 10-25 to work the following 4,000 690 10 68 72
sample problem. 8,000 635 16
12,000 565 22 00
Sample Problem 5 16,000 29 10 9
20,000 21 20
Departing Airport Pressure Altitude..................Sea level 32 31
44 45
Departing Airport OAT............................................22 °C 57 61

Cruise Pressure Altitude....................................8,000 feet Figure 10-25. Fuel time distance climb.

Takeoff Weight.............................................3,400 pounds 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
To begin, find the given weight of 3,400 in the first column of accounting for the additional ten percent, the findings should
the chart. Move across to the pressure altitude column to find read 11 minutes, 23.1 pounds of fuel, and 22 NM. Notice that
the sea level altitude numbers. At sea level, the numbers read the fuel is reported in pounds of fuel, not gallons. Aviation fuel
zero. Next, read the line that corresponds with the cruising weighs six pounds per gallon, so 23.1 pounds of fuel is equal
altitude of 8,000 feet. Normally, a pilot would subtract these to 3.85 gallons of fuel (23.1 ÷ 6 = 3.85).
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 The next example is a cruise and range performance chart.
to climb from sea level to 8,000 feet is 10 minutes. It is also This type of table is designed to give TAS, fuel consumption,
known that 21 pounds of fuel will be used and 20 NM will be endurance in hours, and range in miles at specific cruise
covered during the climb. However, the temperature is 22 °C, configurations. Use Figure 10-26 to determine the cruise and
which is 7° above the standard temperature of 15 °C. The notes range performance under the given conditions.
section of this chart indicate that the findings must be increased
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,

10-22

"

Notes Conditions Gross weight—2,300 lb. Sample Problem 7
Standard conditions
Zero wind Pressure Altitude at Cruise................................6,000 feet
Lean mixture
Maximum cruise is normally limited to 75% power. OAT..................................................36 °F above standard

ALT RPM % TAS GAL/ 38 gal 48 gal Refer to Figure 10-27 for this sample problem. First, locate
BHP MPH Hour (no reserve) (no reserve) 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
2,500 2,700 86 134 9.7 Endr. Range Endr. Range under the 20 °C (or 36 °F) column. At 6,000 feet, the rpm
2,600 79 129 8.6 hours miles hours miles setting of 2,450 will maintain 65 percent continuous power
2,500 72 123 7.8 525 660 at 21.0 "Hg with a fuel flow rate of 11.5 gallons per hour and
2,400 65 117 7.2 3.9 570 4.9 720 airspeed of 161 knots.
2,300 58 111 6.7 4.4 600 5.6 760
2,200 52 103 6.3 4.9 620 6.2 780 Another type of cruise chart is a best power mixture range
82 134 9.0 5.3 630 6.7 795 graph. This graph gives the best range based on power
5,000 2,700 75 128 8.1 5.7 625 7.2 790 setting and altitude. Using Figure 10-28, find the range at
2,600 68 122 7.4 6.1 7.7 65 percent power with and without a reserve based on the
2,500 61 116 6.9 provided conditions.
2,400 55 108 6.5 4.2 565 5.3 710
2,300 49 100 6.0 4.7 600 5.9 760 Sample Problem 8
2,200 78 133 8.4 5.1 625 6.4 790
71 127 7.7 5.5 635 6.9 805 OAT....................................................................Standard
7,500 2,700 64 121 7.1 5.9 635 7.4 805
2,600 58 113 6.7 6.3 630 7.9 795 Pressure Altitude...............................................5,000 feet
2,500 52 105 6.2
2,400 70 129 7.6 4.5 600 5.7 755
2,300 67 125 7.3 4.9 625 6.2 790
61 118 6.9 5.3 645 6.7 810
10,000 2,650 55 110 6.4 5.7 645 7.2 820
2,600 49 100 6.0 6.1 640 7.7 810
2,500
2,400 5.0 640 6.3 810
2,300 5.2 650 6.5 820
5.5 655 7.0 830
5.9 650 7.5 825
6.3 635 8.0 800

Figure 10-26. Cruise and range performance. First, move up the left side of the graph to 5,000 feet and
standard temperature. Follow the line straight across the
read that the endurance in hours is 5.5 hours and the range graph until it intersects the 65 percent line under both the
in miles is 635 miles. reserve and no reserve categories. Draw a line straight down
from both intersections to the bottom of the graph. At 65
Cruise power setting tables are useful when planning cross- percent power with a reserve, the range is approximately
country flights. The table gives the correct cruise power 522 miles. At 65 percent power with no reserve, the range
settings, as well as the fuel flow and airspeed performance should be 581 miles.
numbers at that altitude and airspeed.
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

CRUISE POWER SETTING
65% MAXIMUM CONTINUOUS POWER (OR FULL THROTTLE)

2,800 POUNDS

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

Press IOAT Engine Man. Fuel TAS IOAT Engine Man. Fuel TAS IOAT Engine Man. Fuel TAS
ALT speed press flow per speed press flow per speed press flow per
engine kts MPH engine engine

°F °C RPM " HG PSI GPH 147 169 °F °C RPM " HG PSI GPH kts MPH °F °C RPM " HG PSI GPH kts MPH
149 171
S.L. 27 -3 2,450 20.7 6.6 11.5 152 175 63 17 2,450 21.2 6.6 11.5 150 173 99 37 2,450 21.8 6.6 11.5 153 176
2,000 19 -7 2,450 20.4 6.6 11.5 155 178 55 13 2,450 21.0 6.6 11.5 153 176 91 33 2,450 21.5 6.6 11.5 156 180
4,000 12 -11 2,450 20.1 6.6 11.5 157 181 48 9 2,450 20.7 6.6 11.5 156 180 84 29 2,450 21.3 6.6 11.5 159 183
6,000 5 -15 2,450 19.8 6.6 11.5 160 184 41 5 2,450 20.4 6.6 11.5 158 182 79 26 2,450 21.0 6.6 11.5 161 185
8,000 -2 -19 2,450 19.5 6.6 11.5 162 186 36 2 2,450 20.2 6.6 11.5 161 185 72 22 2,450 20.8 6.6 11.5 164 189
10,000 -8 -22 2,450 19.2 6.6 11.5 159 183 28 -2 2,450 19.9 6.6 11.5 163 188 64 18 2,450 20.3 6.5 11.4 166 191
12,000 -15 -26 2,450 18.8 6.4 11.3 156 180 21 -6 2,450 18.8 6.1 10.9 163 188 57 14 2,450 18.8 5.9 10.6 163 188
14,000 -22 -30 2,450 17.4 5.8 10.5 14 -10 2,450 17.4 5.6 10.1 160 184 50 10 2,450 17.4 5.4 9.8 160 184
16,000 -29 -34 2,450 16.1 5.3 9.7 2,450 16.1 5.1 9.4 156 180 43 6 2,450 16.1 4.9 9.1 155 178
7 -14

Notes 1. Full throttle manifold pressure settings are approximate.
2. Shaded area represents operation with full throttle.

Figure 10-27. Cruise power setting.

10-23

14 -13° Best Power Mixture Range No reserve Begin by finding the correct OAT on the bottom, left side of
12 -9° the graph. Move up that line until it intersects the pressure
45 minutes reserve at 55% altitude of 6,000 feet. Draw a line straight across to the
power best economy mixture 65 percent, best power line. This is the solid line, which
represents best economy. Draw a line straight down from
10 -5° Power 65% Power 65% this intersection to the bottom of the graph. The TAS at 65
Power 55% Power 55% percent best power is 140 knots. However, it is necessary
to subtract 8 knots from the speed since there are no wheel
8 -1° Power 75% Power 75% fairings. This note is listed under the title and conditions.
The TAS will be 132 knots.
6 3°
Crosswind and Headwind Component Chart
4 7° Notes Range may be reduced Every aircraft is tested according to Federal Aviation
2 11° by up to 7% if wheel Administration (FAA) regulations prior to certification. The
fairings are not installed aircraft is tested by a pilot with average piloting skills in
90° crosswinds with a velocity up to 0.2 VSO or two-tenths
S.L. 15° 500 550 600 500 550 600 650 of the aircraft’s stalling speed with power off, gear down,
450 and flaps down. This means that if the stalling speed of the
Range (nautical miles) aircraft is 45 knots, it must be capable of landing in a 9-knot,
Pressure ALT (1,000 feet) (Includes distance to climb and descend) 90° crosswind. The maximum demonstrated crosswind
Standard Temperature °C component is published in the AFM/POH. The crosswind and
Add 0.6 NM for each Associated conditions headwind component chart allows for figuring the headwind
Notes degree Celsius above and crosswind component for any given wind direction and
standard temperature Mixture Leaned per section 4 velocity.
and subtract 1 NM for Weight 2,300 lb.
each degree Celsius Wings No Sample Problem 10
below standard Fuel 48 gal usable
temperature. Wheel Fairings installed Runway..........................................................................17
Cruise Mid cruise
Wind........................................................140° at 25 knots
Figure 10-28. Best power mixture range.
Refer to Figure 10-30 to solve this problem. First, determine
setting. Using Figure 10-29, find the TAS performance based how many degrees difference there is between the runway
on the given information. and the wind direction. It is known that runway 17 means
a direction of 170°; from that subtract the wind direction of
Sample Problem 9 140°. This gives a 30° angular difference, or wind angle. Next,
OAT.........................................................................16 °C locate the 30° mark and draw a line from there until it intersects
Pressure Altitude...............................................6,000 feet
Power Setting................................65 percent, best power
Wheel Fairings..............................................Not installed

20,000 Notes Subtract 8 knots if wheel
fairings are not installed.

18,000 Standard temperature Power 55% 65%
16,000 75%
14,000
12,000 Pressure ALT (feet) Power Best power
10,000 Power Best economy
8,000 –3-2b-labldaedeprporpop
Associated conditions

2,27,0507 r5prpmmatat363I6NI.N.M.MP..P–. Weight 3,600 lb. gross weight

Flaps Up

Best power Mixture leaned to 100°

6,000 rich of peak EGT
4,000
2,000 Best economy Mixture leaned to peak EGT

Sea level 1,650° Max allowable EGT

Wheel Fairings installed

-40° -30° -20° -10° 0° 10° 20° 30° 40° 100 120 140 160 180 200
Outside air temperature (°C) True airspeed (knots)

Figure 10-29. Cruise performance graph.

10-24

0° 10° Sample Problem 11
70 20°
60 30° Pressure Altitude...............................................1,250 feet
50
Headwind component Wind velocity 40° Temperature.........................................................Standard
50°
Refer to Figure 10-31. This example makes use of a landing
40 60° distance table. Notice that the altitude of 1,250 feet is not
on this table. It is, therefore, necessary to interpolate to find
30 the correct landing distance. The pressure altitude of 1,250
70° 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
20 total distance of 1,075 for sea level and the total distance of
80° 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
10 feet total landing distance to clear a 50 foot obstacle. Repeat
this process to obtain the ground roll distance for the pressure
90° altitude. The ground roll should be 457.5 feet.
0 10 20 30 40 50 60 70
Sample Problem 12
Crosswind component
OAT.......................................................................... 57 °F
Figure 10-30. Crosswind component chart.
Pressure Altitude.............................................. 4,000 feet
the correct wind velocity of 25 knots. From there, draw a
line straight down and a line straight across. The headwind Landing Weight...........................................2,400 pounds
component is 22 knots and the crosswind component is 13
knots. This information is important when taking off and Headwind.............................................................. 6 knots
landing so that, first of all, the appropriate runway can be
picked if more than one exists at a particular airport, but also Obstacle Height..................................................... 50 feet
so that the aircraft is not pushed beyond its tested limits.
Using the given conditions and Figure 10-32, determine the
Landing Charts landing distance for the aircraft. This graph is an example of
Landing performance is affected by variables similar to those a combined landing distance graph and allows compensation
affecting takeoff performance. It is necessary to compensate for temperature, weight, headwinds, tailwinds, and varying
for differences in density altitude, weight of the airplane, and obstacle height. Begin by finding the correct OAT on the
headwinds. Like takeoff performance charts, landing distance scale on the left side of the chart. Move up in a straight
information is available as normal landing information, as line to the correct pressure altitude of 4,000 feet. From this
well as landing distance over a 50 foot obstacle. As usual, read intersection, move straight across to the first dark reference
the associated conditions and notes in order to ascertain the line. Follow the lines in the same diagonal fashion until the
basis of the chart information. Remember, when calculating correct landing weight is reached. At 2,400 pounds, continue
landing distance that the landing weight will not be the same in a straight line across to the second dark reference line.
as the takeoff weight. The weight must be recalculated to Once again, draw a line in a diagonal manner to the correct
compensate for the fuel that was used during the flight. wind component and then straight across to the third dark

Conditions Flaps lowered to 40° LANDING DISTANCE
Power off
Hard surface runway
Zero wind

Gross Approach speed At sea level & 59 °F At 2,500 ft & 59 °F At 5,000 ft & 41 °F At 7,500 ft & 32 °F
weight IAS, MPH
Ground roll Total to clear Ground roll Total to clear Ground roll Total to clear Ground roll Total to clear
lb 50 ft OBS 50 ft OBS 50 ft OBS 50 ft OBS

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

1. Decrease the distances shown by 10% for each 4 knots of headwind.Note
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.

Figure 10-31. Landing distance table.

10-25

Associated conditions Reference lineWeight Speed 3,500
Reference line(pounds)at 50 feet 3,000
Power Retarded to maintain 2,500
900 feet/on final approach Tailwind2,950kts MPH 2,000
Flaps Down 2,800 1,500
Landing gear Down Reference line2,60070 80 1,000
Runway Paved, level, dry surface 2,400 68 78
Approach speed IAS as tabulated appliOcbaGsbtluiaecdfleeolrihInenietsgerhntmostediate2,20065 75
Braking Maximum 63 72
60 69

Pressure altitude (feet) Headwind

10,080,0060,00040,020,000S0.L. ISA

C -40° -30° -20° -10° 0° 10° 20° 30° 40° 50° 2,800 2,600 2,400 2,200 0 10 20 30 0 500
Outside air temperature Weight 50
(pounds)
F -40° -20° 0° 20° 40° 60° 80° 100° 120° Wind component Obstacle

(knots) height (feet)

Figure 10-32. Landing distance graph. Stall Speed

reference line. From this point, draw a line in two separate Angle of bank
directions: one straight across to figure the ground roll and
one in a diagonal manner to the correct obstacle height. This Gross weight Level 30° 45° 60°
should be 900 feet for the total ground roll and 1,300 feet for 2,750 lb
the total distance over a 50 foot obstacle. 62 Gear and flaps up 88
Power On MPH 54 76
Stall Speed Performance Charts knots 75 67 74 106
Stall speed performance charts are designed to give an 65 58 64 92
understanding of the speed at which the aircraft will stall Off MPH 81 89
in a given configuration. This type of chart will typically knots 54 70 77 76
take into account the angle of bank, the position of the gear 47 Gear and flaps down 66
and flaps, and the throttle position. Use Figure 10-33 and Power On MPH 66 93
the accompanying conditions to find the speed at which the knots 57 58 64 81
airplane will stall. 50 56
Off MPH 71 78
Sample Problem 13 knots 62 68

Power........................................................................ OFF Figure 10-33. Stall speed table.

Flaps....................................................................... Down Performance charts provide valuable information to the
pilot. Take advantage of these charts. A pilot can predict the
Gear........................................................................ Down performance of the aircraft under most flying conditions, and
this enables a better plan for every flight. The Code of Federal
Angle of Bank............................................................. 45° Regulations (CFR) requires that a pilot be familiar with all
information available prior to any flight. Pilots should use
First, locate the correct flap and gear configuration. The the information to their advantage as it can only contribute
bottom half of the chart should be used since the gear and to safety in flight.
flaps are down. Next, choose the row corresponding to a
power-off situation. Now, find the correct angle of bank Transport Category Airplane
column, which is 45°. The stall speed is 78 mph, and the Performance
stall speed in knots would be 68 knots.
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.

10-26

Major Differences in Transport Category • VS—stalling speed or the minimum steady flight speed
Versus Non-Transport Category Performance at which the aircraft is controllable.
Requirements
• VMCG—minimum control speed on the ground, with
• Full temperature accountability—all of the performance one engine inoperative, (critical engine on two-engine
charts for the transport category aircraft require that airplanes) takeoff power on other engine(s), using
takeoff and climb performance be computed with the aerodynamic controls only for directional control
full effects of temperature considered. (must be less than V1).

• Climb performance expressed as percent gradient • VMCA—minimum control speed in the air, with one
of climb—the transport category aircraft’s climb engine inoperative, (critical engine on two-engine
performance is expressed as a percent gradient of aircraft) operating engine(s) at takeoff power,
climb rather than a figure calculated in fpm of climb. maximum of 5° bank into the good engine(s).
This percent gradient of climb is a much more practical
expression of performance since it is the aircraft’s • V1—critical engine failure speed or decision speed.
angle of climb that is critical in an obstacle clearance Engine failure below this speed shall result in an
situation. aborted takeoff; above this speed the takeoff run
should be continued.
• Change in lift-off technique—lift-off technique in
transport category aircraft allows the reaching of V2 • VR—speed at which the rotation of the aircraft is
(takeoff safety speed) after the aircraft is airborne. initiated to takeoff attitude. The speed cannot be less
This is possible because of the excellent acceleration than V1 or less than 1.05 times VMC. With an engine
and reliability characteristics of the engines on these failure, it must also allow for the acceleration to V2
aircraft and due to the larger surplus of power. at the 35-foot height at the end of the runway.

• Performance requirements applicable to all segments • VLOF—lift-off speed. The speed at which the aircraft
of aviation—all aircraft certificated by the FAA in the first becomes airborne.
transport category, whatever the size, must be operated
in accordance with the same performance criteria. • V2—the takeoff safety speed which must be attained
This applies to both commercial and non-commercial at the 35-foot height at the end of the required runway
operations. distance. This is essentially the best one-engine
operative angle of climb speed for the aircraft and
Performance Requirements should be held until clearing obstacles after takeoff,
The performance requirements that the transport category or until at least 400 feet above the ground.
aircraft must meet are:
• VFS—final segment climb speed, which is based upon
Takeoff one-engine inooerative climb, clean configuration, and
mximum continuos power setting.
• Takeoff speeds
All of the V speeds should be considered during every
• Takeoff runway required takeoff. The V1, VR, V2, and VFS speeds should be visibly
posted in the flightdeck for reference during the takeoff.
• Takeoff climb required
Takeoff speeds vary with aircraft weight. Before takeoff
• Obstacle clearance requirements speeds can be computed, the pilot must first determine the
maximum allowable takeoff weight. The three items that can
Landing limit takeoff weight are runway requirements, takeoff climb
requirements, and obstacle clearance requirements.
• Landing speeds
Runway Requirements
• Landing runway required The runway requirements for takeoff are affected by:

• Landing climb required • Pressure altitude

Takeoff Planning • Temperature
Listed below are the speeds that affect the transport category
aircraft’s takeoff performance. The flight crew must be • Headwind component
thoroughly familiar with each of these speeds and how they
are used in takeoff planning. • Runway gradient or slope

• Aircraft weight

10-27

The runway required for takeoff must be based upon the 3. Takeoff distance—the distance required to complete
possible loss of an engine at the most critical point, which is an all-engines operative takeoff to the 35-foot height.
at V1 (decision speed). By regulation, the aircraft’s takeoff It must be at least 15 percent less than the distance
weight has to accommodate the longest of three distances: required for a one-engine inoperative engine takeoff.
This distance is not normally a limiting factor as it is
1. Accelerate-go distance—the distance required to usually less than the one-engine inoperative takeoff
accelerate to V1 with all engines at takeoff power, distance.
experience an engine failure at V1 and continue the
takeoff on the remaining engine(s). The runway These three required takeoff runway considerations are
required includes the distance required to climb to 35 shown in Figure 10-34.
feet by which time V2 speed must be attained.
Balanced Field Length
2. Accelerate-stop distance—the distance required to In most cases, the pilot will be working with a performance
accelerate to V1 with all engines at takeoff power, chart for takeoff runway required, which will give “balanced
experience an engine failure at V1, and abort the takeoff field length” information. This means that the distance
and bring the aircraft to a stop using braking action
only (use of thrust reVv1ersing is not considered). VR VLO V2

Takeoff distance with
an engine failure

35'

25 25 25All engine acceleration One engine acceleration

V1

Accelerate-stop
distance

All engine acceleration Stop distance
VR VLO
V1

115% All engine takeoff
distance

All engine takeoff distance to 35' 35'

All engine takeoff field length 15% margin

Figure 10-34. Minimum required takeoff.

10-28

shown for the takeoff will include both the accelerate-go and For other than normal takeoff conditions, such as with
accelerate-stop distances. One effective means of presenting engine anti-ice, anti-skid brakes inoperative, or extremes in
the normal takeoff data is shown in the tabulated chart in temperature or runway slope, the pilot should consult the
Figure 10-35. appropriate takeoff performance charts in the performance
section of the AFM.
The chart in Figure 10-35 shows the runway distance required
under normal conditions and is useful as a quick reference There are other occasions of very high weight and
chart for the standard takeoff. The V speeds for the various temperature where the runway requirement may be dictated
weights and conditions are also shown. by the maximum brake kinetic energy limits that affect

Notes Conditions Cabin pressurization On TAKEOFF RUNWAY REQUIREMENTS
Zero slope runway Standard ISA conditions
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.

Takeoff Temp. Pressure altitude (feet) Headwind
gross weight (knots)
°F °C Sea level (V1) 1,000 (V1) 2,000 (V1) 3,000 (V1) 4,000 (V1) 5,000 (V1) 6,000 (V1) 0
at brake 20
release 30 -1.1 47 (121) 48 (121) 50 (120) 53 (121) 57 (122) 62 (123) 70 (123) 0
50 10 48 (121) 51 (121) 55 (121) 60 (122) 63 (123) 69 (124) 77 (125) 20
19,612 70 21 53 (122) 56 (122) 60 (123) 65 (124) 70 (125) 77 (125) 85 (126) 0
VR = 126 90 32 58 (123) 62 (124) 68 (124) 73 (125) 78 (126) 85 (127) 95 (129) 20
V2 = 134 30 -1.1 43 (121) 43 (121) 45 (120) 48 (121) 52 (122) 56 (123) 64 (123) 0
50 10 43 (121) 46 (121) 50 (122) 55 (122) 57 (123) 63 (124) 70 (125) 20
19,000 70 21 48 (122) 51 (122) 55 (123) 59 (124) 63 (125) 70 (125) 77 (126) 0
VR = 124 90 32 53 (123) 57 (124) 62 (124) 66 (125) 71 (126) 77 (127) 85 (129) 20
V2 = 131 30 -1.1 45 (118) 45 (118) 47 (117) 50 (118) 54 (119) 59 (120) 66 (120) 0
50 10 46 (118) 48 (118) 51 (118) 56 (119) 59 (120) 65 (121) 73 (121) 20
18,000 70 21 50 (118) 53 (119) 57 (120) 66 (121) 66 (121) 72 (122) 80 (123)
VR = 119 90 32 55 (120) 59 (121) 64 (121) 73 (122) 73 (123) 80 (124) 90 (124)
V2 = 127 30 -1.1 40 (118) 41 (118) 43 (117) 45 (118) 49 (119) 54 (120) 60 (120)
50 10 42 (118) 44 (118) 46 (118) 51 (119) 54 (120) 59 (121) 66 (121)
17,000 70 21 45 (118) 48 (119) 52 (120) 56 (121) 60 (121) 65 (122) 72 (123)
VR = 115 90 32 50 (120) 54 (121) 58 (121) 63 (122) 66 (123) 73 (124) 81 (124)
V2 = 124 30 -1.1 40 (114) 41 (114) 42 (113) 45 (113) 49 (114) 53 (115) 60 (115)
50 10 41 (115) 43 (114) 46 (114) 50 (115) 53 (115) 59 (116) 66 (117)
16,000 70 21 45 (114) 48 (115) 51 (115) 56 (116) 59 (116) 65 (116) 72 (117)
VR = 111 90 32 50 (115) 53 (116) 58 (116) 62 (117) 66 (118) 73 (118) 80 (119)
V2 = 120 30 -1.1 36 (114) 37 (114) 38 (113) 41 (113) 45 (114) 48 (115) 54 (115)
50 10 37 (115) 39 (114) 42 (114) 46 (115) 48 (115) 54 (116) 60 (117)
15,000 70 21 41 (114) 44 (115) 46 (115) 51 (116) 56 (116) 59 (116) 65 (117)
VR = 106 90 32 46 (115) 48 (116) 53 (116) 56 (117) 60 (118) 66 (118) 73 (119)
V2 = 116 30 -1.1 36 (108) 37 (108) 38 (107) 40 (108) 44 (109) 48 (110) 53 (111)
50 10 37 (110) 39 (108) 41 (109) 45 (110) 48 (110) 53 (111) 59 (112)
70 21 40 (108) 43 (110) 46 (111) 50 (111) 53 (112) 58 (111) 65 (113)
90 32 45 (111) 46 (112) 52 (112) 56 (113) 59 (114) 65 (114) 72 (114)
30 -1.1 32 (108) 33 (108) 34 (107) 36 (108) 40 (109) 44 (110) 48 (111)
50 10 34 (110) 35 (108) 37 (109) 41 (110) 44 (110) 48 (111) 54 (112)
70 21 36 (108) 39 (110) 42 (111) 45 (111) 48 (112) 53 (111) 59 (113)
90 32 41 (111) 44 (112) 47 (112) 51 (113) 54 (114) 59 (114) 65 (114)
30 -1.1 32 (104) 33 (103) 34 (103) 36 (103) 39 (105) 43 (106) 48 (106)
50 10 34 (105) 35 (103) 37 (104) 41 (105) 43 (106) 47 (107) 53 (107)
70 21 36 (104) 38 (105) 41 (105) 45 (106) 48 (107) 52 (107) 58 (108)
90 32 41 (106) 43 (107) 46 (107) 50 (108) 53 (108) 58 (109) 64 (110)
30 -1.1 29 (104) 30 (103) 31 (103) 32 (103) 35 (105) 39 (106) 44 (106)
50 10 31 (105) 32 (103) 33 (104) 37 (105) 39 (106) 43 (107) 48 (107)
70 21 32 (104) 34 (105) 37 (105) 41 (106) 44 (107) 47 (107) 53 (108)
90 32 37 (106) 39 (107) 42 (107) 45 (108) 48 (108) 53 (109) 58 (110)
30 -1.1 28 (98) 30 (98) 30 (98) 32 (98) 35 (99) 38 (101) 42 (101)
50 10 30 (100) 31 (98) 33 (99) 36 (100) 38 (101) 42 (102) 46 (102)
70 21 32 (99) 34 (100) 37 (101) 40 (102) 42 (102) 46 (102) 51 (103)
90 32 36 (101) 38 (102) 41 (102) 44 (103) 47 (104) 51 (104) 56 (105)
30 -1.1 25 (98) 27 (98) 27 (98) 29 (98) 32 (99) 34 (101) 38 (101)
50 10 27 (100) 29 (98) 30 (99) 32 (100) 34 (101) 38 (102) 42 (102)
70 21 29 (99) 31 (100) 33 (101) 36 (102) 38 (102) 42 (102) 46 (103)
90 32 32 (101) 34 (102) 37 (102) 40 (103) 43 (104) 46 (104) 51 (105)

Figure 10-35. Normal takeoff runway required.

10-29

the aircraft’s ability to stop. Under these conditions, the NOTE: Climb gradient can best be described as being a
accelerate-stop distance may be greater than the accelerate- specific gain of vertical height for a given distance covered
go. The procedure to bring performance back to a balanced horizontally. For instance, a 2.4 percent gradient means that
field takeoff condition is to limit the V1 speed so that it 24 feet of altitude would be gained for each 1,000 feet of
does not exceed the maximum brake kinetic energy speed distance covered horizontally across the ground.
(sometimes called VBE). This procedure also results in a
reduction in allowable takeoff weight. The following brief explanation of the one-engine inoperative
climb profile may be helpful in understanding the chart in
Climb Requirements Figure 10-36.

After the aircraft has reached the 35 foot height with one engine First Segment
inoperative, there is a requirement that it be able to climb This segment is included in the takeoff runway required
at a specified climb gradient. This is known as the takeoff charts, and is measured from the point at which the aircraft
flightpath requirement. The aircraft’s performance must be becomes airborne until it reaches the 35-foot height at the
considered based upon a one-engine inoperative climb up to end of the runway distance required. Speed initially is VLO
1,500 feet above the ground. The takeoff flightpath profile and must be V2 at the 35 foot height.
with required gradients of climb for the various segments
and configurations is shown in Figure 10-36.

Ground roll 1st segment—climb 2nd segment—climb 3rd segment— 4th
acceleration segment—

climb

V235'
400'
V1 VR VLOF 1,500'

Landing gear Down Retraction Retracted
Engine All operating
Down One inoperative
Airspeed Variable Takeoff V2 Variable 1.2V5FVSsomr in.
Flaps Retracted
Power
M.C.

Items 1st T/O 2nd T/O Transition Final T/O M.C. = Maximum continuous
segment segment (acceleration) segment
2 Engine V1 = Critical-engine-failure speed
Positive 2.4% Positive 1.2% V2 = Takeoff safety speed
* 3 Engine 3.0% 2.7% Positive 1.5% VS = Calibrated stalling speed, or
4 Engine 5.0% 3.0% Positive 1.7%
Wing flaps T.O. T.O. Up minimum steady flight speed at
Landing gear Down T.O.
Engines 1 Out Up Up Up which the aircraft is controllable
Power T.O. 1 Out 1 Out 1 Out
Air speed T.O. T.O. M.C. VR = Speed at which aircraft can
VLOF V2 V2 1.25 VS(Min) 1.25 VS(Min) start safely raising nose wheel
V2
off surface (Rotational Speed)

VLOF = Speed at point where airplane
lifts off

* Required Absolute Minimum Gradient of Flightpath

Figure 10-36. One engine inoperative takeoff.

10-30

Second Segment be met. The aircraft may well be capable of lifting off with
This is the most critical segment of the profile. The second one engine inoperative, but it must then be able to climb
segment is the climb from the 35 foot height to 400 feet and clear obstacles. Although second segment climb may
above the ground. The climb is done at full takeoff power not present much of a problem at the lower altitudes, at the
on the operating engine(s), at V2 speed, and with the flaps higher altitude airports and higher temperatures, the second
in the takeoff configuration. The required climb gradient segment climb chart should be consulted to determine the
in this segment is 2.4 percent for two-engine aircraft, 2.7 effects on maximum takeoff weights before figuring takeoff
percent for three-engine aircraft, and 3.0 percent for four- runway distance required.
engine aircraft.
Air Carrier Obstacle Clearance
Third or Acceleration Segment Requirements
During this segment, the airplane is considered to be
maintaining the 400 feet above the ground and accelerating Regulations require that large transport category turbine
from the V2 speed to the VFS speed before the climb profile powered aircraft certificated after September 30, 1958, be
is continued. The flaps are raised at the beginning of the taken off at a weight that allows a net takeoff flightpath
acceleration segment and power is maintained at the takeoff (one engine inoperative) that clears all obstacles either by
setting as long as possible (5 minutes maximum). a height of at least 35 feet vertically, or by at least 200 feet
horizontally within the airport boundaries and by at least 300
Fourth or Final Segment feet horizontally after passing the boundaries. The takeoff
This segment is from the 400 to 1,500 foot AGL altitude flightpath is considered to begin 35 feet above the takeoff
with power set at maximum continuous. The required climb surface at the end of the takeoff distance, and extends to a
in this segment is a gradient of 1.2 percent for two-engine point in the takeoff at which the aircraft is 1,500 feet above
airplanes, 1.55 for three-engine airplanes, and 1.7 percent the takeoff surface, or at which the transition from the takeoff
for four-engine airplanes. to the en route configuration is completed. The net takeoff
flightpath is the actual takeoff flightpath reduced at each point
Second Segment Climb Limitations by 0.8 percent for two engine aircraft, 0.9 percent for three-
The second segment climb requirements, from 35 to 400 engine aircraft, and 1.0 percent for four-engine aircraft.
feet, are the most restrictive (or hardest to meet) of the
climb segments. The pilot must determine that the second Air carrier pilots therefore are responsible not only for
segment climb is met for each takeoff. In order to achieve this determining that there is enough runway available for an
performance at the higher density altitude conditions, it may engine inoperative takeoff (balanced field length), and
be necessary to limit the takeoff weight of the aircraft. the ability to meet required climb gradients; but they must
also assure that the aircraft will safely be able to clear any
It must be realized that, regardless of the actual available obstacles that may be in the takeoff flightpath. The net
length of the takeoff runway, takeoff weight must be takeoff flightpath and obstacle clearance required are shown
adjusted so that the second segment climb requirements can in Figure 10-37.

MinimuOmNbeastcttatauckaleelotcafflkefealiografhfntfpcliaegthhptpaathth 35'

400'
35'

35'

Figure 10-37. Takeoff obstacle clearance.

10-31

The usual method of computing net takeoff flightpath • VREF—1.3 times the stalling speed in the landing
performance is to add up the total ground distances required configuration. This is the required speed at the 50-foot
for each of the climb segments and/or use obstacle clearance height above the threshold end of the runway.
performance charts in the AFM. Although this obstacle
clearance requirement is seldom a limitation at the normally • Approach climb—the speed which gives the best climb
used airports, it is quite often an important consideration performance in the approach confguration, with one
under critical conditions such as high takeoff weight and/or engine inoperative, and with maximum takeoff power
high density altitude. Consider that at a 2.4 percent climb on the operating engine(s). The required gradient of
gradient (2.4 feet up for every 100 feet forward) a 1,500 climb in this configuration is 2.1 percent for two-
foot altitude gain would take a horizontal distance of 10.4 engine aircraft, 2.4 percent for three-emgine aircraft,
NM to achieve. and 2.7 percent for four-engine aircraft.

Summary of Takeoff Requirements • Landing climb—the speed giving the best performance
In order to establish the allowable takeoff weight for a in the full landing configuration with maximum
transport category aircraft, at any airfield, the following must takeoff power on all engines. The gradient of climb
be considered: required in this configuration is 3.2 percent.

• Airfield pressure altitude Planning the Landing

• Temperature As in the takeoff, the landing speeds shown above should be
precomputed and visible to both pilots prior to the landing.
• Headwind component The VREF speed, or threshold speed, is used as a reference
speed throughout the traffic pattern or instrument approach
• Runway length as in the following example:

• Runway gradient or slope VREF plus 30K Downwind or procedure turn
VREF plus 20K Base leg or final course inbound to final
• Obstacles in the flightpath VREF plus 10K fix
VREF Final or final course inbound from fix
Once the above details are known and applied to the (ILS final)
appropriate performance charts, it is possible to determine Speed at the 50 foot height above the
the maximum allowable takeoff weight. This weight would threshold
be the lower of the maximum weights as allowed by:
Landing Requirements
• Balanced field length required
The maximum landing weight of an aircraft can be restricted
• Engine inoperative climb ability (second segment by either the approach climb requirements or by the landing
limited) runway available.

• Obstacle clearance requirement Approach Climb Requirements

In practice, restrictions to takeoff weight at low altitude The approach climb is usually more limiting (or more difficult
airports are usually due to runway length limitations; engine to meet) than the landing climb, primarily because it is based
inoperative climb limitations are most common at the higher upon the ability to execute a missed approach with one engine
altitude airports. All limitations to weight must be observed. inoperative. The required climb gradient can be affected
Since the combined weight of fuel and payload in the aircraft by pressure altitude and temperature and, as in the second
may amount to nearly half the maximum takeoff weight, segment climb in the takeoff, aircraft weight must be limited
it is usually possible to reduce fuel weight to meet takeoff as needed in order to comply with this climb requirement.
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.

10-32

Landing Runway Required Summary of Landing Requirements
The runway distance needed for landing can be affected by In order to establish the allowable landing weight for a
the following: transport category aircraft, the following details must be
considered:
• Pressure altitude
• Airfield pressure altitude
• Temperature
• Temperature
• Headwind component
• Headwind component
• Runway gradient or slope
• Runway length
• Aircraft weight
• Runway gradient or slope
In computing the landing distance required, some manufacturers
do not include all of the above items in their charts, since the • Runway surface condition
regulations state that only pressure altitude, wind, and aircraft
weight must be considered. Charts are provided for anti-skid With these details, it is possible to establish the maximum
on and anti-skid off conditions, but the use of reverse thrust is allowable landing weight, which will be the lower of the
not used in computing required landing distances. weights as dictated by:

The landing distance, as required by the regulations, is • Landing runway requirements
that distance needed to land and come to a complete stop
from a point 50 feet above the threshold end of the runway. • Approach climb requirements
It includes the air distance required to travel from the 50
foot height to touchdown (which can consume 1,000 feet In practice, the approach climb limitations (ability to climb
of runway distance), plus the stopping distance, with no in approach configuration with one engine inoperative)
margin left over. This is all that is required for 14 CFR part are seldom encountered because the landing weights upon
91 operators (non-air carrier), and all that is shown on some arrival at the destination airport are usually low. However,
landing distance required charts. as in the second segment climb requirement for takeoff, this
approach climb gradient must be met and landing weights
For air carriers and other commercial operators subject to must be restricted if necessary. The most likely conditions
14 CFR part 121, a different set of rules applies stating that that would make the approach climb critical would be the
the required landing distance from the 50 foot height cannot landings at high weights and high pressure altitudes and
exceed 60 percent of the actual runway length available. temperatures, which might be encountered if a landing were
In all cases, the minimum airspeed allowed at the 50 foot required shortly after takeoff.
height must be no less than 1.3 times the aircraft’s stalling
speed in the landing configuration. This speed is commonly Landing field requirements can more frequently limit an
called the aircraft’s VREF speed and varies with landing aircraft’s allowable landing weight than the approach climb
weight. Figure 10-38 is a diagram of these landing runway limitations. Again, however, unless the runway is particularly
requirements. 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.

VSO1.A3 pVpSrOoach Complete stop

50'

25

Required landing distance per 14 CFR part 91 40% of runway length
Required landing runway length per 14 CFR part 121

Figure 10-38. Landing runway requirements.

10-33

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.

10-34

Chapter 11

Weather Theory

Introduction

Weather is an important factor that influences aircraft
performance and flying safety. It is the state of the atmosphere
at a given time and place, with respect to variables such as
temperature (heat or cold), moisture (wetness or dryness),
wind velocity (calm or storm), visibility (clearness or
cloudiness), and barometric pressure (high or low). The term
weather can also apply to adverse or destructive atmospheric
conditions, such as high winds.
This chapter explains basic weather theory and offers pilots
background knowledge of weather principles. It is designed
to help them gain a good understanding of how weather
affects daily flying activities. Understanding the theories
behind weather helps a pilot make sound weather decisions
based on the reports and forecasts obtained from a Flight
Service Station (FSS) weather specialist and other aviation
weather services.
Be it a local flight or a long cross-country flight, decisions
based on weather can dramatically affect the safety of the
flight.

11-1