Aircraft 

Weight and Balance 

Handbook

U.S. Department of Transportation

FEDERAL AVIATION ADMINISTRATION

Flight Standards Service

2016 
ii
iii

The Aircraft Weight and Balance Handbook has been prepared in recognition of the importance of weight and balance 

technology in conducting safe and efficient flight. The objective of this handbook is twofold: to provide the airframe and 

powerplant mechanic (A&P) with the method of determining the empty weight and empty weight center of gravity (EWCG) 

of an aircraft and to furnish the flight crew with information on loading and operating the aircraft to ensure its weight is 

within the allowable limit and the center of gravity (CG) is within the allowable range.

Any time there is a conflict between the information in this handbook and specific information issued by an aircraft 

manufacturer, the manufacturer’s data takes precedence over information in this handbook. Occasionally, the word “must” 

or similar language is used where the desired action is deemed critical. The use of such language is not intended to add to, 

interpret, or relieve a duty imposed by Title 14 of the Code of Federal Regulations (14 CFR). 

It is essential for persons using this handbook to become familiar with and apply the pertinent parts of 14 CFR. The current 

Flight Standards Service airman training and testing material and learning statements for all airman certificates and ratings 

can be obtained from www.faa.gov.

This handbook supersedes FAA-H-8083-1A, Aircraft Weight and Balance Handbook, dated 2007. 

This handbook is available for download, in PDF format, from www.faa.gov.

This handbook is published by the United States Department of Transportation, Federal Aviation Administration, Airman 

Testing Standards Branch, AFS-630, P.O. Box 25082, Oklahoma City, OK 73125.

Comments regarding this publication should be sent, in email form, to the following address:

AFS630comments@faa.gov.

John S. Duncan

Director, Flight Standards Service

Preface
iv
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This handbook begins with the basic principle of aircraft weight and balance control, emphasizing its importance and 

including examples of documentation furnished by the aircraft manufacturer and by the FAA to ensure the aircraft weight 

and balance records contain the proper data.

Procedures for the preparation and the actual weighing of an aircraft are described, as are the methods of determining the 

location of the empty weight center of gravity (EWCG) relative to both the datum and the mean aerodynamic chord (MAC).

Loading computations for general aviation aircraft are discussed using both loading graphs and tables of weight and 

moment indexes.

Information is included that allows an FAA-certifi ated mechanic or repairman to determine the weight and center of gravity 

(CG) changes caused by repairs and alterations or removal and installation of equipment. This includes instructions for 

conducting adverse-loaded CG checks, also explaining the way to determine the amount and location of ballast needed to 

bring the CG within allowable limits.

The unique requirements for helicopter weight and balance control are discussed, including the determination of lateral CG 

and the way both lateral and longitudinal CG change as fuel is consumed.

One chapter includes the methods and examples of solving weight and balance problems using a hand-held electronic 

calculator, E6-B flight computer, and a dedicated electronic flight compute

Introduction
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The Aircraft Weight and Balance Handbook (FAA-H-8083-1B) was produced by the Federal Aviation Administration 

(FAA) with the assistance of Safety Research Corporation of America (SRCA). The FAA wishes to acknowledge the 

following contributors:

Larry Jackson of Jackson Aircraft Weighing Service, for content and photographs used in Chapter 3

The White Planes Picture Company (www.whiteplanes.com), for images used in Chapter 4

Terri Sipantzi of Precision Windsports, for the weight-shift control sample weight and loading diagram used in Chapter 4

AirBorne Windsports Pty Ltd., for weight-shift control sample weight and loading data sheet used in Chapter 4

Jim Stevens of Aerosports Connection, for powered parachute sample weight and balance information sheet used in 

Chapter 4

Sporty’s Pilot Shop, for image of an E6B Flight Computer used in Chapter 10

Acknowledgments
viii
ix

Preface....................................................................iii

Introduction.............................................................v

Acknowledgments................................................vii

Table of Contents ..................................................ix

Chapter 1

Weight and Balance Control...............................1-1

Introduction....................................................................1-1

Responsibility for Weight and Balance Control ............1-2

Terminology...................................................................1-2

Weight Control ..............................................................1-2

Effects of Weight ...........................................................1-3

Weight Changes.............................................................1-3

Stability and Balance Control........................................1-4

Weight Control for Aircraft Other Than Fixed and 

Rotor-wing .....................................................................1-5

Weight-Shift Control Aircraft ....................................1-5

Powered Parachutes....................................................1-6

Balloons......................................................................1-6

Underestimating the Importance of Weight and 

Balance...........................................................................1-6

Chapter 2

Weight and Balance Theory ...............................2-1

Introduction....................................................................2-1

Weight and Balance Theory...........................................2-2

Arm ............................................................................2-2

Moment ......................................................................2-2

The Law of the Lever.....................................................2-2

Shifting the Balance Point or CG...................................2-4

Solution by Chart .......................................................2-4

Basic Weight and Balance Equation ..........................2-4

Solution by Formula...................................................2-5

Mean Aerodynamic Chord.............................................2-5

FAA-Furnished Weight and Balance Information ........2-6

CG Range ...................................................................2-6

Empty Weight CG Range (EWCG) ...........................2-6

Maximum Weights.....................................................2-6

Number of Seats.........................................................2-6

Maximum Baggage ...................................................2-6

Fuel Capacity..............................................................2-6

Oil Capacity (Wet Sump)...........................................2-6

Data Pertinent to all Models.....................................2-10

Manufacturer-Furnished Information ..........................2-10

Chapter 3

Weighing the Aircraft and Determining the 

Empty Weight Center of Gravity ........................3-1

Introduction....................................................................3-1

Requirements .................................................................3-2

Equipment for Weighing................................................3-2

Preparation for Weighing...............................................3-3

Scale Preparation........................................................3-3

Weigh Clean Aircraft Inside Hangar..........................3-4

Equipment List ...........................................................3-4

Ballast.........................................................................3-4

Standard Weights .......................................................3-4

Draining the Fuel........................................................3-4

Oil...............................................................................3-4

Other Fluids................................................................3-4

Configuration of the Aircraf ......................................3-5

Jacking the Aircraft ....................................................3-5

Leveling the Aircraft ..................................................3-5

Safety Considerations.................................................3-5

Determining the CG.......................................................3-5

EWCG Formulas............................................................3-6

Datum Forward of the Airplane—Nosewheel 

Landing Gear..............................................................3-6

Datum Aft of the Main Wheels—Nosewheel 

Landing Gear..............................................................3-6

Location of Datum.........................................................3-7

Datum Forward of the Main Wheels—Tailwheel 

Landing Gear..............................................................3-7

Datum Aft of the Main Wheels—Tailwheel 

Landing Gear..............................................................3-7

Center of Gravity (CG) and Mean Aerodynamic 

Chord (MAC).................................................................3-8

Table of Contents
x

Center of Gravity (CG)...............................................3-8

Mean Aerodynamic Chord (MAC) ............................3-8

Chapter 4

Light Sport Aircraft Weight and Balance 

Control..................................................................4-1

Introduction....................................................................4-1

LSA Definition of Term ............................................4-2

Weight and Balance .......................................................4-3

WSC Aircraft..............................................................4-3

Powered Parachutes....................................................4-4

Weight and Balance Computations (Amateur-Built 

Aircraft)..........................................................................4-4

Chapter 5

Single-Engine Aircraft Weight and Balance 

Computations ......................................................5-1

Introduction....................................................................5-1

Determining the Loaded Weight and CG ......................5-2

Manual Computational Method..................................5-2

Loading Graph Method ..............................................5-4

Moment Indexes .....................................................5-4

Loading Graph........................................................5-4

Compute Weight and Balance Using the 

Loading Graph........................................................5-4

Chapter 6

Multiengine Aircraft Weight and Balance 

Computations ......................................................6-1

Introduction....................................................................6-1

Determine the Loaded CG .............................................6-2

Chart Method Using Weight, Arm, and Moments.....6-2

Determining the CG in Percentage of Mean 

Aerodynamic Chord (MAC) ......................................6-3

The Chart Method Using Weight and 

Moment Indexes.........................................................6-3

Chapter 7

Center of Gravity Change After a 

Repair or Alteration.............................................7-1

Introduction....................................................................7-1

Equipment List...............................................................7-2

Major Alteration and Repair..........................................7-3

Weight and Balance Revision Record ...........................7-3

Weight Changes Caused by a Repair or Alteration .......7-5

Computations Using Weight, Arm, and Moment.......7-5

Computations Using Weight and Moment Indexes ...7-5

Determining the CG in Percentage of Mean 

Aerodynamic Chord (Percent MAC) .........................7-6

Empty Weight CG (EWCG) Range...............................7-6

Adverse-Load CG Checks .............................................7-6

Forward Adverse-Load CG Check.............................7-7

Aft Adverse-Load CG Check.....................................7-7

Ballast ............................................................................7-7

Temporary Ballast......................................................7-8

Temporary Ballast Formula........................................7-8

Permanent Ballast.......................................................7-8

Chapter 8

Weight and Balance Control—Helicopter .........8-1

Introduction....................................................................8-1

Determining the Loaded CG of a Helicopter.................8-3

Effects of Offloading Passengers and Using Fue ..........8-3

Chapter 9

Weight and Balance Control—Commuter 

Category and Large Aircraft...............................9-1

Introduction....................................................................9-1

Establishing the Initial Weight of an Aircraft ...............9-2

Determining the Empty Weight and Empty Weight 

CG (EWCG)...................................................................9-2

Documenting Changes to an Aircraft’s Weight and 

Balance ..........................................................................9-3

Determining the Loaded CG of the Airplane in 

Percent MAC .................................................................9-3

Operational Empty Weight (OEW) ...............................9-5

Reestablishing the OEW ............................................9-5

Fleet Operating Empty Weights (FOEW) .................9-5

Onboard Aircraft Weighing System...........................9-5

Determining the Correct Stabilizer Trim Setting .......9-6

Determining CG Changes Caused by Modifying the 

Cargo..............................................................................9-6

Effects of Loading or Offloading Cargo ....................9-6

Effects of Shifting Cargo From One Hold to 

Another.......................................................................9-8

Determining Cargo Pallet Loads and Floor 

Loading Limits...........................................................9-9

Determining the Maximum Amount of Payload 

That Can Be Carried.................................................9-10

Determining the Landing Weight.............................9-10

Determining Fuel Dump Time in Minutes...............9-12

Weight and Balance of Commuter Category 

Airplanes ..................................................................9-12

Determining the Loaded Weight and CG.................9-12

Determining the Changes in CG When Passengers 

Are Shifted ...............................................................9-16

Determining Changes in Weight and CG When 

the Aircraft Is Operated in Its Cargo 

Configuratio ............................................................9-18

Determining the CG Shift When Cargo Is Moved 

From One Section to Another ..................................9-18

Determining the CG Shift When Cargo Is Added 

or Removed ..............................................................9-18

Determining Which Limits Are Exceeded ...............9-18
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Chapter 10

Use of Computer for Weight and Balance 

Computations ....................................................10-1

Introduction..................................................................10-1

Electronic Calculator ...................................................10-2

E6-B Flight Computer..................................................10-2

Dedicated Electronic Flight Computer ........................10-3

Typical Weight and Balance Problems....................10-4

Determining CG in Inches From the Datum ............10-4

Nosewheel Airplane With Datum Ahead of the 

Main Wheels.........................................................10-4

Nosewheel Airplane With Datum Behind the 

Main Wheels.........................................................10-4

Tailwheel Airplane With Datum Ahead of the 

Main Wheels.........................................................10-4

Tailwheel Airplane With Datum Behind the 

Main Wheels.........................................................10-4

Determining CG, Given Weights, and Arms ...........10-5

Determining CG, Given Weights, and Moment 

Indexes......................................................................10-5

Determining CG in Percent Mean Aerodynamic 

Chord (MAC)...........................................................10-5

Determining Lateral CG of a Helicopter..................10-5

Determining ΔCG Caused by Shifting Weights.......10-6

Determining Weight Shifted to Cause Specified

ΔCG..........................................................................10-6

Determining Distance Weight Is Shifted to Move 

CG a Specific Distanc .............................................10-6

Determining Total Weight of an Aircraft With a 

Specified ΔCG When Cargo Is Move ....................10-7

Determining Amount of Ballast Needed to Move 

CG to a Desired Location.........................................10-7

Appendix A..........................................................A-1

Appendix B..........................................................B-1

Glossary ..............................................................G-1

Index ......................................................................I-1
xii
1-1

Introduction

There are many factors in the safe and efficient operation of 

aircraft, including proper weight and balance control. The 

weight and balance system commonly employed among 

aircraft consists of three equally important elements: the 

weighing of the aircraft, the maintaining of the weight and 

balance records, and the proper loading of the aircraft. An 

inaccuracy in any one of these elements defeats the purpose 

of the system. The final loading calculations are meaningless 

if either the aircraft has been improperly weighed or the 

records contain an error.

Improper loading decreases the effic ency and performance of 

an aircraft from the standpoint of altitude, maneuverability, 

rate of climb, and speed. It may even be the cause of failure 

to complete the flight or, for that matter, failure to start the 

flight.Because of abnormal stresses placed upon the structure 

of an improperly loaded aircraft, or because of changed fly ng 

characteristics of the aircraft, loss of life and destruction of 

valuable equipment may result.

Weight and Balance 

Control

Chapter 1
1-2

Aircraft can perform safely and achieve their designed 

efficiency only when they are operated and maintained in the 

way their designers intended. This safety and efficiency is 

determined to a large degree by holding the aircraft’s weight 

and balance parameters within the limits specifiedfor its design. 

The remainder of this handbook describes how this is done.

Responsibility for Weight and Balance 

Control

The responsibility for proper weight and balance control 

begins with the engineers and designers and extends to the 

technicians who maintain the aircraft and the pilots who 

operate them. Modern aircraft are engineered utilizing stateof-the-art technology and materials to achieve maximum 

reliability and performance for the intended category. As 

much care and expertise must be exercised in operating 

and maintaining these efficient aircraft as was taken in their 

design and manufacturing:

1. The designers of an aircraft set the maximum weight 

based on the amount of lift the wings or rotors can 

provide under the operational conditions for which 

the aircraft is designed. The structural strength of the 

aircraft also limits the maximum weight the aircraft 

can safely carry. The designers carefully determine 

the ideal center of gravity (CG) and calculate the 

maximum allowable deviation from this specific 

location.

2. The manufacturer provides the aircraft operator with 

the empty weight of the aircraft and the location of its 

empty weight center of gravity (EWCG) at the time 

the certified aircraft leaves the factory. Amateur-built 

aircraft must have this information determined and 

available at the time of certification

3. The FAA-certificated mechanic or repairman who 

maintains the aircraft keeps the weight and balance 

records current, recording any changes that have been 

made because of repairs or alterations.

4. The pilot in command (PIC) has the responsibility 

prior to every flight to know the maximum allowable 

weight of the aircraft and its CG limits. This allows 

the pilot to determine during the preflight inspection 

that the aircraft is loaded so that the CG is within the 

allowable limits.

Terminology

Pilots and FAA-certificated mechanics or repairmen must 

ensure they understand the terms as they relate to the 

aircraft in question. For small aircraft terminology, use 

the information found in sources associated with Civil 

Air Regulation (CAR) 3 certification or General Aviation 

Manufacturers Association (GAMA) Specification No. 1 

for part 23 aircraft or part 27 for rotorcraft. For terminology 

applied to large part 25 aircraft, information can be found 

in Advisory Circular (AC) 120-27, Aircraft Weight and 

Balance Control. The glossary contains the most current 

terms and defin tions. Current regulations are available from 

the Superintendent of Documents; U.S. Government Printing 

Office; Washington, DC 20402. They are also located on the 

FAA website at www.faa.gov. Earlier regulations may be 

available in libraries or in the Federal Register.

Weight Control 

Weight is a major factor in airplane construction and 

operation, and it demands respect from all pilots and 

particular diligence by all maintenance personnel. Excessive 

weight reduces the efficiency of an aircraft and the available 

safety margin if an emergency condition should arise.

When an aircraft is designed, it is made as light as the 

required structural strength allows, and the wings or rotors are 

designed to support the maximum allowable weight. When 

the weight of an aircraft is increased, the wings or rotors must 

produce additional lift and the structure must support not 

only the additional static loads, but also the dynamic loads 

imposed by flight maneuvers. For example, the wings of a 

3,000-pound airplane must support 3,000 pounds in level 

flight, but when the airplane is turned smoothly and sharply 

using a bank angle of 60°, the dynamic load requires the 

wings to support twice this or 6,000 pounds.

Severe uncoordinated maneuvers or flight into turbulence 

can impose dynamic loads on the structure great enough 

to cause failure. In accordance with Title 14 of the Code 

of Federal Regulations (14 CFR) part 23, the structure of a 

normal category airplane must be strong enough to sustain 

a load factor of 3.8 times its weight. Every pound of weight 

added to a normal category aircraft requires that the structure 

be strong enough to support 3.8 pounds. An aircraft operated 

in the utility category must sustain a load factor of 4.4 times 

its weight, and acrobatic category aircraft must be strong 

enough to withstand 6.0 times their weight.

The lift produced by a wing is determined by its airfoil shape, 

angle of attack, speed through the air, and air density. When 

an aircraft takes off from an airport with a high density 

altitude, it must accelerate to a speed faster than would be 

required at sea level to produce enough lift to allow takeoff; 

therefore, a longer takeoff run is necessary. The distance 

needed may be longer than the available runway. When 

operating from a high density altitude airport, the Pilot’s 

Operating Handbook (POH) or Airplane Flight Manual 

(AFM) must be consulted to determine the maximum weight 

allowed for the aircraft under the conditions of altitude, 

temperature, wind, and runway conditions.
1-3

Effects of Weight

Most modern aircraft are so designed that, when all seats 

are occupied, the baggage compartment is full, and all fuel 

tanks are full, the aircraft is grossly overloaded. This type 

of design requires the pilot to give great consideration to 

the requirements of each specific flight. If maximum range 

is required, occupants or baggage must be left behind, or if 

the maximum load must be carried, the range, dictated by 

the amount of fuel on board, must be reduced.

Overloading an aircraft can create a variety of problems:

• The aircraft needs a higher takeoff speed, which results 

in a longer takeoff run.

• Both the rate and angle of climb are reduced.

• The service ceiling is lowered.

• The cruising speed is reduced.

• The cruising range is shortened.

• Maneuverability is decreased.

• A longer landing roll is required because the landing 

speed is higher.

• Excessive loads are imposed on the structure, 

especially the landing gear.

The POH or AFM includes tables or charts that give the pilot 

an indication of the performance expected for any weight. 

An important part of careful preflight planning includes a 

check of these charts to determine if the aircraft is loaded so 

the proposed flight can be safely made

Weight Changes

The maximum allowable weight for an aircraft is determined 

by design considerations. However, the maximum operational 

weight may be less than the maximum allowable weight due 

to such considerations as high density altitude or high-drag 

field conditions caused by wet grass or water on the runway. 

The maximum operational weight may also be limited by the 

departure or arrival airport’s runway length.

One important preflight consideration is the distribution 

of the load in the aircraft. Loading the aircraft so the gross 

weight is less than the maximum allowable is not enough. 

This weight must be distributed to keep the CG within the 

limits specified in the POH or AFM

If the CG is too far forward, a heavy passenger can be moved 

to one of the rear seats or baggage may be shifted from a 

forward baggage compartment to a rear compartment. If the 

CG is too far aft, passenger weight or baggage can be shifted 

forward. The fuel load should be balanced laterally. The pilot 

should pay special attention to the POH or AFM regarding 

the operation of the fuel system in order to keep the aircraft 

balanced in flight

Weight and balance of a helicopter is far more critical 

than for an airplane. Some helicopters may be properly 

loaded for takeoff, but near the end of a long flight when 

the fuel tanks are almost empty, the CG may have shifted 

enough for the helicopter to be out of balance laterally or 

longitudinally. Before making any long flight, the CG with 

the fuel available for landing must be checked to ensure it is 

within the allowable range.

Airplanes with tandem seating normally have a limitation 

requiring solo flight to be made from the front seat in some 

airplanes or the rear seat in others. Some of the smaller 

helicopters also require solo flight be made from a specific

seat, either the right, left, or center. These seating limitations 

are noted by a placard, usually on the instrument panel, and 

they should be strictly followed.

As an aircraft ages, its weight usually increases due to debris 

and dirt collecting in hard-to-reach locations and moisture 

absorbed in the cabin insulation. This increase in weight is 

normally small, but it can be determined only by accurately 

weighing the aircraft.

Changes of fixed equipment may have a major effect upon 

the weight of the aircraft. Many aircraft are overloaded by 

the installation of extra radios or instruments. Fortunately, 

the replacement of older, heavy electronic equipment with 

newer, lighter types results in a weight reduction. This 

weight change, however helpful, can cause the CG to shift, 

which must be computed and annotated in the weight and 

balance record.

Repairs and alterations are the major sources of weight 

changes. It is the responsibility of the FAA-certificate 

mechanic or repairman making any repair or alteration to 

know the weight and location of a change, to compute the 

CG, record the new empty weight and EWCG in the aircraft 

weight and balance record, and update the equipment lists.

If the newly calculated EWCG should happen to fall outside 

the EWCG range, it is necessary to perform an adverseloading check. This requires a forward and rearward 

adverse-loading check and a maximum weight check. 

These weight and balance extreme conditions represent the 

maximum forward and rearward CG position for the aircraft. 

An adverse-loading check is a deliberate attempt to load an 

aircraft in a manner that creates the most critical balance 

condition and still remains within the design CG limits of the 

aircraft. If any of the checks fall outside the loaded CG range, 
1-4

Figure 1-1. Longitudinal forces acting on an airplane in flight. Lift

Variable

Nose-up force

Fixed dependent upon airspeed

Nose-down force

independent of airspeed

CG

Figure 1-2. If the CG is too far aft at the low stall airspeed, there 

might not be enough elevator nose-down authority to get the nose 

down for recovery.

Lift

Insufficient elevator

Nose-down force

CG too far aft

CG

Figure 1-3. If the CG is too far forward, there is not enough elevator 

nose-up force to flare the airplane for landing. Lift

Insufficient elevator

Nose-up force

CG too far forward

CG

the aircraft must be reconfigured or placarded to prevent the 

pilot from loading the aircraft improperly. It is sometimes 

possible to install a fixed ballast in order for the aircraft to 

operate again within the normal CG range.

The FAA-certificated mechanic or repairman conducting an 

annual or condition inspection must ensure the weight and 

balance data in the aircraft records is current and accurate. It 

is the responsibility of the PIC to use the most current weight 

and balance data when operating the aircraft.

Stability and Balance Control

Balance control refers to the location of the CG of an aircraft. 

This is of primary importance to aircraft stability, which is a 

factor in flight safety. The CG is the point at which the total 

weight of the aircraft is assumed to be concentrated, and the 

CG must be located within specific limits for safe flight. Both 

lateral and longitudinal balance are important, but the prime 

concern is longitudinal balance; that is, the location of the 

CG along the longitudinal or lengthwise axis.

An airplane is designed to have stability that allows it to be 

trimmed to maintain straight-and-level flight with hands off 

the controls. Longitudinal stability is maintained by ensuring 

the CG is slightly ahead of the center of lift. This produces a 

fixed nose-down force independent of the airspeed. This is 

balanced by a variable nose-up force, which is produced by a 

downward aerodynamic force on the horizontal tail surfaces 

that varies directly with the airspeed. [Figure 1-1]

If a rising air current should cause the nose to pitch up, the 

airplane slows and the downward force on the tail decreases. 

The weight concentrated at the CG pulls the nose back down. 

If the nose should drop in flight, the airspeed increases and 

the increased downward tail load brings the nose back up 

to level flight

As long as the CG is maintained within the allowable limits 

for its weight, the airplane has adequate longitudinal stability 

and control. If the CG is too far aft, it is too near the center 

of lift; the airplane is unstable and difficult to recover from 

a stall. [Figure 1-2] If the unstable airplane should enter a 

spin, the spin could become flat making recovery difficult

or impossible. If the CG is too far forward, the downward 

tail load needs to be increased to maintain level flight. This 

increased tail load has the same effect as carrying additional 

weight; the aircraft must fly at a higher angle of attack and 

drag increases.

A more serious problem caused by the CG being too far 

forward is the lack of sufficient elevator authority. At low 

takeoff speeds, the elevator might not produce enough 

nose-up force to rotate; on landing there may not be enough 

elevator force to flare the airplane. [Figure 1-3] Both takeoff 

and landing runs are lengthened if the CG is too far forward.

The basic aircraft design is such that lateral symmetry is 

assumed to exist. For each item of weight added to the left 

of the center line of the aircraft (also known as buttock 

line zero or BL‐0), there is generally an equal weight at a

corresponding location on the right.

The lateral balance can be upset by uneven fuel loading 

or burnoff. The position of the lateral CG is not normally 

computed for an airplane, but the pilot must be aware of 
1-5

Figure 1-4. Lateral imbalance causes wing heaviness, which may 

be corrected by deflecting the aileron. The additional lift causes 

additional drag, and the airplane flies inefficiently.

Empty Full

Additional lift and drag

Additional weight

Figure 1-5. Fuel in the tanks of a swept-wing airplane affects both 

lateral and longitudinal balance. As fuel is used from an outboard 

tank, the CG shifts forward.

Inboard fuel: nose heavy

Outboard fuel: tail heavy

the adverse effects that result from a laterally unbalanced 

condition. [Figure 1-4] This is corrected by using the aileron 

trim tab until enough fuel has been used from the tank on 

the heavy side to balance the airplane. The deflected trim tab 

deflects the aileron to produce additional lift on the heavy 

side, but it also produces additional drag, and the airplane 

flies inefficientl

Helicopters are more often affected by lateral imbalance 

than airplanes. If a helicopter is loaded with heavy occupants 

and fuel on the same side, it could be out of balance enough 

to make it unsafe to fly. It is also possible that if external 

loads are carried in such a position that requires large lateral 

displacement of the cyclic control to maintain level fl ght, the 

fore-and-aft cyclic control effectiveness is limited.

Swept-wing airplanes are more critical due to fuel imbalance 

because as the fuel is used from the outboard tanks, the CG 

shifts forward. As fuel is used from the inboard tanks, the CG 

shifts aft. [Figure 1-5] For this reason, fuel-use scheduling 

in swept-wing airplanes operation is critical.

Weight Control for Aircraft Other Than 

Fixed and Rotor-wing

Some light aircraft utilize different methods of determining 

weight and balance from the traditional fixed and rotor-wing 

aircraft. These aircraft achieve flight control by methods 

different from the fixed-wing airplane or helicopter. Most 

notable of these are weight-shift control (WSC) aircraft (also 

known as trikes), powered parachutes, and balloons. These 

aircraft typically do not specify either an EWCG or a CG 

range. They require only a certified or approved maximum 

weight. To understand why this is so, a look at how flight

control is achieved is helpful.

Airplanes and WSC aircraft control flight under the influenc 

of the same four forces (lift, gravity, thrust, and drag), and 

around the same three axes (pitch, yaw, and roll). However, 

each aircraft accomplishes this control in a very different 

manner. This difference helps explain why the fixed-wing

airplane requires an established weight and a known CG, 

whereas the WSC aircraft only requires the known weight.

The fixed-wing airplane has movable controls that alter lift 

on various airfoil surfaces to vary pitch, roll, and yaw. In 

turn, these changes in lift affect the characteristics of the 

flight parameters. Weight normally decreases in flight due 

to fuel consumption, and the airplane CG changes with this 

weight reduction. An airplane utilizes its variable fligh 

controls to compensate and maintain controllability through 

the various flight modes and as the CG changes. An airplane 

has a CG range or envelope within which it must remain if 

the flight controls are to remain effective and the airplane 

safely operated.

Weight-Shift Control Aircraft

The WSC aircraft has a relatively set platform wing without 

a tail. The pilot achieves control by shifting weight. In the 

design of this aircraft, the weight of the airframe and its 

payload is attached to the wing at a single point in a pendulous 

arrangement. The pilot, through the flight controls, controls the 

arm of this pendulum and thereby controls the aircraft. When 

a change in flight parameter is desired, the pilot displaces the 

aircraft’s weight by the appropriate distance and direction. 

This change momentarily disrupts the equilibrium between the 

four forces acting on the aircraft. The wing, due to its inherent 

stability, then moves appropriately to reestablish the desired 

relationship between these forces; the wing flexes and alter 

its shape. As the shape is changed, lift is varied at different 

points on the wing to achieve the desired flight parameters
1-6

The flight controls primarily affect the pitch-and-roll axes. 

Since there is no vertical tail plane, there is minimal or no 

yaw control. Unlike in an airplane, the CG experienced by 

the WSC aircraft wing remains constant. Since the weight 

of the airframe acts through a single point (the wing attach 

point), the range over which the weight may act is fixed at 

the pendulum arm or length. Even though weight decreases 

as fuel is consumed, weight remains focused at the wing 

attach point. Because the range is fixed, there is no need to 

establish a calculated range. The pilot should consult the 

POH or Aircraft Operating Instructions (AOI) for maximum 

takeoff weight and minimum and maximum seat weight 

limits prior to each flight

Powered Parachutes

The powered parachute is also a pendulum-style aircraft. Its 

airframe CG is fixed at the pendulum attach point. It is more 

limited in controllability than the WSC aircraft because it 

lacks an aerodynamic pitch control. Pitch (and lift) control 

is primarily a function of the power control. Increased power 

results in increased lift; cruise power amounts to level flight;

decreased power causes a descent. Due to this characteristic, 

the aircraft is basically a one-airspeed aircraft. Once again, 

because the CG is fixed at the attach point to the wing, there 

is no CG range. As with WSC, the pilot should consult the 

POH or AOI for maximum takeoff weight and minimum and 

maximum seat weight limits prior to each flight

Roll control on a powered parachute is achieved by changing 

the shape of the wing. The change is achieved by varying 

the length of steering lines attached to the outboard trailing 

edges of the wing. The trailing edge of the parachute is pulled 

down slightly on one side or the other to create increased 

drag along that side. This change in drag creates roll and 

yaw, permitting the aircraft to be steered.

Balloons

The balloon is controlled by the pilot only in the vertical 

dimension; this is in contrast to all other aircraft. He or she 

achieves this control through the use of lift and weight. 

Wind provides all other movement. The CG of the gondola 

remains constant beneath the balloon envelope. As in WSC 

and powered-parachute aircraft, there is no CG limitation.

Underestimating the Importance of 

Weight and Balance

Many pilots, from sport pilot to commercial pilot, tend to 

underestimate the importance of proper weight and balance 

of their aircraft. Load sheets are taken for granted and hasty 

calculations are made of the aircraft’s CG. Unfortunately, 

each year there are a number of accidents related to weight 

and balance issues. Many of these occurrences could have 

been avoided had more attention been given to weight and 

balance.

Every student pilot is taught how to work a weight and 

balance problem and that it is important to make sure 

every flight is loaded “within the envelope” (no more than 

maximum gross weight) for both takeoff and landing. But 

does he or she really understand just why this is so and the 

disastrous effect of being out of the envelope? Two examples 

of documented cases are provided below in an effort to 

indicate the serious nature of maintaining the proper weight 

and balance. In case studies when weight and balance was 

listed as the major factor of the accident, many were fatal. 

For instance, a small aircraft was loaded with hunters, gear, 

and dogs (none of the dogs were secured inside the aircraft). 

During takeoff, all the dogs went to the aft of the airplane. 

This shifted the CG well aft of its allowable limit. The 

airplane stalled and crashed. The airplane was destroyed 

with casualties.

Another accident occurred when a group of skydivers were 

sitting on the floor toward the aft portion of the airplane 

(they were unsecured). During takeoff, the CG was again 

well beyond its aft limit. The airplane stalled and crashed. 

The airplane was destroyed with casualties.

There is a safety factor built into the formula for maximum 

gross weight. Any airplane can fly when it takes off at a 

weight greater than maximum gross weight if the runway 

is long enough and the density altitude is low enough. 

However, landing is a different matter. All airplanes are built 

to withstand an occasional hard landing, but what would 

happen if the hard landing were combined with a substantially 

overweight airplane? Something would probably break at that 

time or the structure would be weakened enough to break 

sometime in the future when everything might seem normal to 

a pilot unaware of the previous situation. Even more disastrous 

than an overweight, hard landing is reaching or exceeding 

the structural integrity of the metal and/or composite design 

values when maneuvering or when turbulence is encountered. 

Hidden damage could result, causing an unexpected 

catastrophic failure at some future time.

If an airplane is certificated with a maximum gross weight 

of 6,000 pounds (its weight on the ground) and is rolled into 

a 60° bank, the forces exerted make it feel as if it weighed 

12,000 pounds. At its maximum certificated gross weight, 

there is no problem because the aircraft is operated within its 

certificated maneuvering loads. But loaded to 8,000 pounds 

with a 60° bank or an abrupt pullup, it suddenly weighs 

16,000 pounds and might not be able to perform! Even if it 

could, there would probably be internal stress damage that 

would show up on future flights
2-1

Introduction

Weight and balance in aircraft is based on the law of the lever. 

This chapter discusses the application of the law of the lever 

and its applications relative to locating the balance point of a 

beam or lever on which various weights are located or shifted. 

The chapter also discusses the documentation pertaining to 

weight and balance that is furnished by the Federal Aviation 

Administration (FAA) and aircraft manufacturers.

Weight and Balance Theory

Chapter 2
2-2

Figure 2-1. Balance lever.

B = 200

A = 100

Arm A (−50") Arm B (+25")

Fulcrum

Datum − Moment + Moment

−Forces

+Forces

Figure 2-2. Balance point locations.

Item Weight (lb) Arm (in) Moment (lb-in)

−5,000

+5,000

0

Weight A

Weight B

100

200

300

−50

+25

Figure 2-3. Balance lever datum located off the lever.

C = 200

A = 100 B = 100

+50

+90

+110

+150

CG

Datum

Weight and Balance Theory

Two elements are vital in the weight and balance considerations 

of an aircraft.

• The total weight of the aircraft must be no greater 

than the maximum weight allowed by the FAA for 

the make and model of the aircraft.

• The center of gravity (CG), or the point at which 

all of the weight of the aircraft is considered to be 

concentrated, must be maintained within the allowable 

range for the operational weight of the aircraft.

Arm 

The arm is usually measured and expressed in inches and 

refers to the horizontal distance between the CG of an item 

or object and the datum, a point from where all measurements 

are taken. Arms to the left of the datum are negative (–) and 

those to the right of the datum are positive (+). The datum is an 

imaginary vertical plane from which all horizontal distances 

are measured for balance purposes. The position of the 

reference datum varies by aircraft design and manufacturer. 

When the datum is located off of the lever and to the left, all of 

the arms are positive and computational errors are minimized. 

Note: When the datum is established ahead of the aircraft, for 

example at the aircraft nose, all of the arms are positive and 

computational errors are minimized. 

Moment

A moment is a force that tries to cause rotation and is the 

product of the arm, in inches, and the weight, in pounds. 

Moments are generally expressed in pound-inches (lb-in) 

and may be either positive or negative. 

The Law of the Lever

Weight and balance problems are based on the physical law 

of the lever. This law states that a lever is balanced when the 

weight on one side of the fulcrum (a pivot point for the lever) 

multiplied by its arm is equal to the weight on the opposite 

side multiplied by its arm. In other words, the lever is balanced 

when the sum of the moments about the fulcrum is zero. This 

is the condition in which the positive moments (those that 

try to rotate the lever clockwise) are equal to the negative 

moments (those that try to rotate it counterclockwise). In an 

aircraft, the balance point is referred to as the CG. 

One of the easiest ways to understand weight and balance is 

to consider a lever with weights placed at various locations. 

The balance point or CG of the lever can be changed by either 

moving the weights closer or farther from the fulcrum or by 

increasing or decreasing the weights. The balance point or 

CG of a lever may be determined by using these four steps:

1. Measure the arm of each weight in inches from the 

datum.

2. Multiply each arm by its weight in pounds to determine 

the moment in pound-inches of each weight.

3. Determine the total of all weights and of all the 

moments. (Disregard the weight of the lever).

4. Divide the total moment by the total weight to 

determine the balance point.

Consider these facts about the lever in Figure 2-1. The 

100-pound weight A is located 50 inches to the left of the 

fulcrum (the datum, in this instance), and it has a moment of

100 × –50 = –5,000 lb-in. The 200-pound weight B is located 

25 inches to the right of the fulcrum, and its moment is

200 × +25 = +5,000 lb-in. In Figure 2-2, the sum of the 

moments is –5,000 + 5,000 = 0, and the lever is balanced. The 

forces that try to rotate it clockwise have the same magnitude 

as those that try to rotate it counterclockwise. If either weight 

is moved or changed, the balance point or CG changes and 

the lever becomes unbalanced.

In Figure 2-3, the datum is located off the lever to the left 

of weight A. Using the information provided in Figure 2-3,
2-3

Figure 2-4. Finding balance point with datum located off the lever.

Item Weight (lb) Arm (in) Moment CG

5,000

9,000

30,000

44,000 110

Weight A

Weight B

Weight C

100

100

200

400

50

90

150

Figure 2-5. Locating balance point.

C = 200

A = 100 B = 100

−20

−60

 110

+40

CG

Original datum New datum

Figure 2-6. Proving balance point with three weights is correct.

Item Weight (lb) Arm (in) Moment (lb-in)

−6,000

−2,000

+8,000

0

Weight A

Weight B

Weight C

100

100

200

−60

−20

+40

Figure 2-7. Locating balance point with datum at C.

C = 200

A = 100 B = 100

−60

−100

Datum

Figure 2-8. Determining new balance point.

Item Weight (lb) Arm (in) Moment CG

−10,000

−6,000

0

−16,000 −40

Weight A

Weight B

Weight C

100

100

200

400

−100

−60

Figure 2-9. Locating balance point with datum left of original.

C = 200

A = 100 B = 100

−40

Datum

−20

−60

+40

determine the balance point by making a chart like the one 

in Figure 2-4.

As noted in Figure 2-4, A weighs 100 pounds and is 50 inches 

from the datum; B weighs 100 pounds and is 90 inches from 

the datum; C weighs 200 pounds and is 150 inches from the 

datum. The total of the weights is 400 pounds, and the total 

moment is 44,000 lb-in.

Determine the balance point by dividing the total moment 

by the total weight. A balance point is equal to the CG and 

can be mathematically written as:

CG = total moment

 total weight

To prove this is the correct balance, move the datum to a 

location 110 inches to the right of the original datum and 

determine the arm of each weight from this new datum. 

[Figure 2-5] Then, make a new chart similar to the one in 

Figure 2-6. If the balance point is correct, the sum of the 

moments is zero.

The new arm of weight A is 60 inches (the difference between 

110 and 50), and since this weight is to the left of the datum, 

its arm is negative or –60 inches. The new arm of weight B is 

20 inches (110 – 90), and it is also to the left of the datum, so 

it is –20; the new arm of weight C is 40 inches (150 – 110). 

It is to the right of the datum and is therefore positive. 

The lever is balanced when the sum of the moments is zero. 

The location of the datum used for determining the arms of 

the weights is not important; it may be in various locations, 

but all of the measurements must be made from the same 

datum location.

The procedure for finding the balance point is the same 

anywhere the datum is located. In Figure 2-7, the datum is 

located at C. Weight A has an arm of –100 inches (negative 

because it is to the left) of the datum and weight B has an 

arm of –60 inches from the datum. The table in Figure 2-8

is used to determine the new balance point.

To verify that this is the correct balance point, move the datum 

40 inches to the left of the original datum and determine the 

arm of each weight from this new datum as in Figure 2-9.
2-4

Figure 2-10. Proving the new balance point is correct.

Item Weight (lb) Arm (in) Moment (lb-in)

−6,000

−2,000

+8,000

0

Weight A

Weight B

Weight C

100

100

200

−60

−20

40

Figure 2-11. Locating balance point with three weights.

B = 200 C = 200

A = 100

80

72

100

CG

Datum

Before Weight Shift

Figure 2-12. Proving the new balance point is correct.

Item Weight (lb) Arm (in) Moment (lb-in)

−5,000

+10,000

+5,000

Weight A

Weight B

Weight C

100

200

−50

+50

Figure 2-13. Weight distribution to balance lever.

B = 200

A = 100 C = 100

−50 50

−25

CG

Datum

After Weight Shift

Figure 2-14. Weight shift provides correct CG.

Item Weight (lb) Arm (in) Moment (lb-in)

−5,000

−5,000

+10,000

0

Weight A

Weight B

Weight C

100

200

200

−50

−25

+50

The new arm for weight A would be –100 + 40 = –60; for 

weight B, –60 + 40 = –20; and point C, is +40. The lever is 

balanced and the balance point is correct when the sum of 

the moments is zero. [Figure 2-10]

Shifting the Balance Point or CG

One common weight and balance problem involves moving 

or shifting weight from one point to another in order to move 

the balance point or CG to a desired location. This can be 

demonstrated by using a lever with three weights to work 

out the problem.

Solution by Chart

As the lever is loaded in Figure 2-11, it balances at a point 

72 inches from the CG of weight A.

To shift weight B so the lever balances about its center, 50 

inches from the CG of weight A, first determine the arm of 

weight B that produces a moment that causes the total moment 

of all three weights around this desired balance point to be 

zero. The combined moment of weights A and C around this 

new balance point is 5,000 lb-in, so the moment of weight B 

must be –5,000 lb-in for the lever to balance. [Figure 2-12]

Determine the arm of weight B by dividing its moment, 

–5,000 lb-in, by its weight of 200 pounds. The arm is –25 

inches. To balance the lever at its center, weight B must be 

placed so its CG is 25 inches to the left of the center of the 

lever. [Figure 2-13]

Figure 2-14 indicates that the shift in weight depicted in 

Figure 2-13 allows the lever to balance as the sum of the 

moments is zero.

Basic Weight and Balance Equation

The following formulas can be used to determine the distance 

weight must be shifted to obtain a desired change in the CG 

location. The equation can also be rearranged to fi d the 

amount of weight required to be shifted to move the CG 

to a desired location, to find the distance the CG is moved 

when a specified amount of weight is shifted, or to find the 

total weight that would allow shifting a specified amount of 

weight to move the CG a given distance.

Weight to be shifted =

  C

 Total weight Distance weight is shifted

Total weight = Weight shifted × Distance weight is shifted

  C

Weight shifted =

 Total weight shifted ×  C

 Distance weight is shifted

∆ CG = Weight shifted × Distance weight is shifte

 Total weight

Distance weight is shifted = Total weight ×  C

 Weight shifted
2-5

Figure 2-14. Center of gravity expressed as percent mean 

aerodynamic chord.

MAC MAC

30% MAC

15% MAC Datum

Quarter chord

Half

semi-span

Solution by Formula

The problem in Figure 2-11 can be solved by using variations 

of this basic equation. First, rearrange the formula to 

determine the distance weight B must be shifted:

Distance weight B is shifted = Total weight ×  C

 Weight shifted

 = 500 × –22

 200

 = –55 inches

The CG of the lever in Figure 2-11 was 72 inches from the 

datum. This CG can be shifted to the center of the lever as in 

Figure 2-13 by moving weight B. If the 200-pound weight B 

is moved 55 inches to the left, the CG shifts from +72 inches 

to +50 inches, a distance of 22 inches.

When the distance the weight is to be shifted is known, 

the amount of weight to be shifted to move the CG to any 

location can be determined by another arrangement of the 

basic equation. Use the following arrangement of the formula 

to determine the amount of weight that has to be shifted from 

station 8 to station +25, to move the CG from station +72 

to station +50. 

Weight shifted = Total weight ×  C

 Distance weight is shifted

 = 500 × 22

 55

 = 200 inches

If the 200-pound weight B is shifted from station +80 to 

station +25, the CG moves from station +72 to station +50. 

A third arrangement of this basic equation is used to 

determine the amount the CG is shifted when a given amount 

of weight is moved for a specified distance (as it was done 

in Figure 2-11). The following formula is used to determine 

the amount the CG is shifted when 200-pound weight B is 

moved from +80 to +25.

∆ CG = Weight shifted × Distance it is shifte

 Total weight

 = 200 × 55

 500

 = 22 inches

Moving weight B from +80 to +25 moves the CG 22 inches 

from its original location at +72 to its new location at +50 

as seen in Figure 2-13.

To complete the calculations, return to the original formula 

and enter the appropriate numbers.

Weight to be shifted =  C

 Total weight Distance weight is shifted

 200 = 22

 500 55

 .4 = .4

The equation is balanced.

Mean Aerodynamic Chord

The CG point affects the stability of the aircraft. To ensure 

the aircraft is safe to fly, the CG must fall within specified

limits established by the manufacturer. 

On some aircraft, the CG is expressed as a percentage of the 

length of the mean aerodynamic chord (MAC) or “percent 

MAC.” [Figure 2-14] In order to make such a calculation, 

the position of the leading edge of the MAC must be known 

ahead of time. 

CG limits are specified forward and aft and/or lateral (left and 

right) limits within which the aircraft’s CG must be located 

during flight. The area between the limits is called the CG 

range of the aircraft. 
2-6

The position of the fore and aft CG limits is measured as a 

percentage of MAC from the MAC leading edge. Usually for 

a single or two-seat aircraft, the most forward position would 

be forward of 30–35 percent MAC. Thus, the allowable CG 

range in a light aircraft should not exceed 20 percent MAC.

Note: For a rectangular width of constant airfoil section 

dimensions, MAC is just the chord. For a symmetricallytapered wing, it is the average of the root chord and the tip 

chord.

FAA-Furnished Weight and Balance 

Information 

The information discussed to this point can be readily applied 

to any aircraft weight and balance problem. To apply the 

techniques, certain elements of information are required. 

This information is obtained from both FAA documents and 

manufacturer provided data. 

Before an aircraft CG can be computed, certain information 

must be known. This information, furnished by the FAA 

for every certificated aircraft in the Type Certificate Data 

Sheets (TCDS) or Aircraft Specifications, can be accessed 

at www.faa.gov. When the design of an aircraft is approved 

by the FAA, an Approved Type Certificate and TCDS are 

issued. The TCDS includes all of the pertinent specification 

for the aircraft; at each annual or 100-hour inspection, it is 

the responsibility of the inspecting mechanic or repairman 

to ensure that the aircraft adheres to them. A note about the 

TCDS: aircraft certifi ated before January 1, 1958, were 

issued Aircraft Specifica ions under the Civil Air Regulations 

(CARs), but when the Civil Aeronautical Administration 

(CAA) was replaced by the FAA, Aircraft Specification 

were replaced by the TCDS.

The weight and balance information on a TCDS includes CG 

range, empty weight CG range (EWCG), maximum weights, 

number of seats, maximum baggage, fuel capacity, oil 

capacity, and datum location. Data pertinent to an individual 

model is located in its respective section of the TCDS. Use 

the TCDS excerpts in Figure 2-15 to verify the following for 

a Cirrus Design Corporation SR 20 normal category:

CG Range

S/N 1005 through 1147:

Forward Limits: 138.7 inches at 2,110 lb with a straight line 

taper to 141.0 in at 2,694 lb and 143.0 in at 2,900 lb

Aft Limits: 144.6 in at 2,110 lb, with straight line taper to 

147.4 in at 2,570 lb, and to 147.9 in at 2,745 lb, and 148.2 

in at 2,900 lb

S/N 1148 through 1877, 1879 through 1885, and S/N 1005 

through 1147 if Cirrus Service Bulletin SB 20-01-00 is 

complied with:

Forward Limits: 138.7 in at 2,110 lb with a straight line taper 

to 141.0 in at 2,694 lb and 144.1 in at 3,000 lb

Aft Limits: 144.6 in at 2,110 lb, with straight line taper to 

147.4 in at 2,570 lb, and to 148.1 in at 2,900 lb, and 148.0 

in at 3,000 lb

S/N 1878, 1886 and Subsequent:

Forward Limits: 137.8 in at 2,100 lb with a straight line taper 

to 139.1 in at 2,700 lb, and to 140.7 in at 3,050 lb

Aft Limits: 148.1 in at 2,100 lb, with straight line to 148.1 

in at 3,050 lb

Empty Weight CG Range (EWCG)

When all of the seats and baggage compartments are located 

close together, it is not possible (as long as the EWCG is 

located within the EWCG range) to legally load the aircraft 

so that its operational CG falls outside this allowable range. 

If the seats and baggage areas extend over a wide range, the 

EWCG range is listed as “None.”

Maximum Weights

The maximum allowable takeoff and landing weights and 

the maximum allowable ramp weight are given. This basic 

information may be altered by a note. Notes are found in data 

pertinent to all models. An example would be Note 6 at the 

end of Figure 2-15.

Number of Seats

The number of seats and their arms are given in such terms 

as: 4 (2 at 143.5 aft of datum, 2 at 180 aft of datum).

Maximum Baggage 

Maximum baggage for this model is 130 pounds at 208 

inches.

Fuel Capacity

This important information is given in such terms as: 60.5 

gal at 153.75 in. Usable: 56 gal (See Note 1). Notes can be 

found in data pertinent to all models.

Oil Capacity (Wet Sump)

The quantity of the full oil supply and its arm are given: 8 

quarts at 76.2 in.
2-7

Figure 2-15. Sample excerpt from TCDS A00009CH.

Page No. 1 2 3 4 5 6 7 

Rev. No. 11 12 12 14 14 16 16 

DEPARTMENT OF TRANSPORTATION

FEDERAL AVIATION ADMINISTRATION 

A00009CH

Revision 16 

Cirrus Design Corporation 

SR20

SR22

April 22, 2009 

TYPE CERTIFICATE DATA SHEET NO. A00009CH

This data sheet, which is part of Type Certificate No. A00009CH, prescribes conditions and limitations under which the product 

for the which type certificate was issued meets the airworthiness requirements of the Federal Aviation Regulations. 

Type Certificate Holder: Cirrus Design Corporation 

4515 Taylor Circle 

Duluth, MN 55811 

I - Model SR20, (Normal Category), Approved October 23, 1998

 Engine Teledyne Continental IO-360-ES, Type Certificate Data Sheet (TCDS) E1CE 

 Fuel 100/100LL minimum grade aviation gasoline 

 Engine Limits Maximum Take-off 2700 RPM (200 hp) 

 Maximum Continuous Power 2700 RPM (200 hp) 

 Propeller and 

 Propeller limits 1. Hartzell Propeller Inc. P/N BHC-J2YF-1BF/F7694 

TCDS P37EA 

Maximum Diameter: 76 inches 

Minimum Diameter: 73 inches 

Number of Blades: 2 

Low Pitch: 14.6°+/-0.5° 

High Pitch: 35.0°+/-1.0° 

Not to be operated above 24 inches of manifold pressure between 1900 and 2200 RPM. 

Spinner: Hartzell P/N A-2295(P) NOTE: Spinner may be painted or polished. 

2. Hartzell Propeller Inc. P/N PHC-J3YF-1MF/F7392-1 

TCDS P36EA 

Maximum Diameter: 74 inches 

Minimum Diameter: 72 inches 

Number of Blades: 3 

Low Pitch: 14.1°+/-0.5° 

High Pitch: 35.0°+/-1.0° 

No operating limitations to 2800 RPM 

Spinner: Hartzell P/N A-2295-1P 

3. Hartzell Propeller Inc. P/N PHC-J3YF-1RF/F7392-1 

TCDS P36EA 

Maximum Diameter: 74 inches 

Minimum Diameter: 72 inches 

Number of Blades: 3 

Low Pitch: 13.9°+/-0.5° 

High Pitch: 35.0°+/-1.0° 

No operating limitations to 2800 RPM 

Spinner: Hartzell P/N A-2295-1(P) NOTE: Spinner may be painted or polished. 
2-8

Figure 2-15. Sample excerpt from TCDS A00009CH (continued).

A00009CH -2- 

 Airspeed Limits S/N 1005 thru 1147:

Vne Never Exceed Speed 200 KIAS 

Vno Maximum Structural Cruising Speed 165 KIAS 

Vo (2900 lbs) Operating Maneuvering Speed 135 KIAS 

Vo (2600 lbs) Operating Maneuvering Speed 126 KIAS 

Vo (2200 lbs) Operating Maneuvering Speed 116 KIAS 

Vfe Maximum Flap Extension Speed 100 KIAS 

Vpd Maximum Parachute Deployment Speed 135 KIAS 

 S/N 1148 thru 1877, 1879 thru 1885, and S/N 1005 thru 1147 if Cirrus Service Bulletin SB 20-01-

00 is complied with:

Vne Never Exceed Speed 200 KIAS 

Vno Maximum Structural Cruising Speed 165 KIAS 

Vo (3000 lbs) Operating Maneuvering Speed 131 KIAS 

Vo (2600 lbs) Operating Maneuvering Speed 122 KIAS 

Vo (2300 lbs) Operating Maneuvering Speed 114 KIAS 

Vfe Maximum Flap Extension Speed 100 KIAS 

Vpd Maximum Parachute Deployment Speed 135 KIAS 

S/N 1878, 1886 and subsequent:

Vne Never Exceed Speed 200 KIAS 

Vno Maximum Structural Cruising Speed 163 KIAS 

Vo (3050 lbs) Operating Maneuvering Speed 130 KIAS 

Vfe Maximum Flap Extension Speed 104 KIAS 

Vpd Maximum Parachute Deployment Speed 133 KIAS 

 C.G. Range S/N 1005 thru 1147:

Forward Limits: 138.7 inches at 2110 lbs with a straight line taper to 141.0 inches at 2694 lbs, and 

143.0 inches at 2900 lbs. 

Aft Limits: 144.6 inches at 2110 lbs, with straight line taper to 147.4 inches at 2570 lbs, and to 

147.9 inches at 2745 lbs, and 148.2 inches at 2900 lbs. 

 S/N 1148 thru 1877, 1879 thru 1885, and S/N 1005 thru 1147 if Cirrus Service Bulletin SB 20-01-

00 is complied with:

Forward Limits: 138.7 inches at 2110 lbs with a straight line taper to 141.0 inches at 2694 lbs, and 

144.1 inches at 3000 lbs. 

Aft Limits: 144.6 inches at 2110 lbs, with straight line taper to 147.4 inches at 2570 lbs, and to 

148.1 inches at 2900 lbs, and 148.0 inches at 3000 lbs. 

S/N 1878, 1886 and subsequent:

Forward Limits: 137.8 inches at 2100 lbs with a straight line taper to 139.1 inches at 2700 lbs, and 

to 140.7 inches at 3050 lbs 

Aft Limits: 148.1 inches at 2100 lbs, with straight line to 148.1 inches at 3050 lbs. 

 Empty Weight 

 C.G. Range None 

 Maximum Weight S/N 1005 thru 1147:

 Takeoff and Landing: 2900 lbs. 

 S/N 1148 thru 1877, 1879 thru 1885, and S/N 1005 thru 1147 if Cirrus Service Bulletin SB 20-01-

00 is complied with:

 Takeoff: 3000 lbs. 

 Landing: 2900 lbs. 

 Zero Fuel: 2900 lbs. 

 S/N 1878, 1886 and subsequent:

 Takeoff and Landing: 3050 lbs. 
2-9

Figure 2-15. Sample excerpt from TCDS A00009CH (continued).

Figure 2-15. Sample excerpt from TCDS A00009CH (continued).

-3- A00009CH 

 Minimum Crew One (1) Pilot 

 Number of Seats 4 (2 at 143.5 inches aft of datum, 2 at 180 inches aft of datum) 

 Maximum Baggage 130 Lbs. at 208 inches 

 Fuel Capacity Total: S/N 1005 thru 1877, 1879 thru 1885 

 60.5 gal at 153.75 inches 

 Usable: 56 gal (See Note 1) 

 S/N 1878, 1886 and subsequent:

 58.5 gal at 154.9 inches 

 Usable: 56 gal (See Note 1) 

 Oil Capacity 8 quarts at 76.2 inches 

 Maximum Operating 

 Altitude With a portable oxygen system, the aircraft is limited to 17,500 ft MSL. 

Oxygen must be provided as required by the operating rules. Only portable oxygen systems listed 

in the FAA Approved Airplane Flight Manual, document number 11934-002, or later FAA 

approved revisions, are allowed. 

 Control Surface 

 Movements Wing Flaps: Up 0°± 0.5° Down 50% 16°± 0.5° Down 100% 32°± 0.5° 

 Aileron: Up 12.5° ± 1.0° Down 12.5° ±1.0° 

 Elevator: Up 25.0° +0°/-1.0° Down 15° ± 1.0° 

 Elevator Trim: Up 17.0° Minimum Down 10.5° ± 1.0° 

 Rudder: Right 20.0° ± 1.0° Left 20.0° ± 1.0° 

 Additional Limitations: Airframe life limit: 12,000 flight hours 

 Design Data: The airplane shall be manufactured in accordance with the latest FAA approved revision of 

“Master Drawing List”, Document No. 13750, or other FAA approved data. NOTE: Document 

No. 12609 is the predecessor document to Document No. 13750. 

 Serial Nos. Eligible 1005 and on 

II - Model SR22, Normal Category, Approved November 30, 2000

 Engine Teledyne Continental IO-550-N, Type Certificate Data Sheet E3SO 

 Engine Limits Maximum Take-off 2700 RPM (310 hp) 

 Maximum Continuous Power 2700 RPM (310 hp) 

 Propeller and 

 Propeller limits 1. Hartzell Propeller Inc. P/N PHC-J3YF-1RF/F7694 or F7694B 

 TCDS P36EA Hartzell 

 Maximum Diameter: 78 inches 

 Minimum Diameter: 76 inches 

 Number of Blades: 3 

 Low Pitch: 14.1°+/-0.5° 

 High Pitch: 35.0°+/-1.0° 

 No operating limitations to 2700 RPM 

 Spinner: Hartzell P/N A-2295-1(P) NOTE: Spinner may be painted or polished. 

 2. McCauley Propeller Systems P/N D3A34C443/78CYA-0 

 TCDS P47GL McCauley 

 Maximum Diameter: 78 inches 

 Minimum Diameter: 76 inches 

 Number of Blades: 3 

-5- A00009CH 

 Minimum Crew One (1) Pilot 

 Number of Seats 4 (2 at 143.5 inches aft of datum, 2 at 180 inches aft of datum) 

 Maximum Baggage 130 Lbs. at 208 inches 

 Fuel Capacity Total: S/N 0002 thru 2333, 2335 thru 2419, and 2421 thru 2437

 84 gallon at 154.9 inches 

 Usable: 81 gallon (See Note 1) 

 S/N 2334, 2420, 2438 and subsequent

 94.5 gallon at 154.9 inches 

 Usable: 92.0 gallon (See Note 1) 

 or 

 58.5 gal at 154.9 inches 

 Usable: 56 gal (See Note 1) 

 Oil Capacity 8 quarts at 77.1 inches 

 Maximum Operating 

 Altitude With a portable oxygen system, the aircraft is limited to 17,500 ft MSL. 

Oxygen must be provided as required by the operating rules. Only portable oxygen systems listed 

in the FAA Approved Airplane Flight Manual, document number 13772-001, or later FAA 

approved revisions, are allowed. 

 Control Surface 

 Movements Wing Flaps: Up 0°±0.5° Down 50% 16° ± 0.5° Down 100% 32° ±0.5° 

 Aileron: Up 12.5° ± 1.0° Down 12.5° ± 1.0° 

 Aileron Trim: Up 6 ± 1.0° Down 6 ± 1.0° 

 Elevator: Up 25.0° +0°/-1.0° Down 15° ± 1.0° 

 Elevator Trim: Up 17.0° Minimum Down 10.5° ± 1.0° 

 Rudder: Right 20.0° ± 1.0° Left 20.0° ± 1.0° 

 Additional Limitations: Airframe life limit: 12,000 flight hours 

 Design Data: The airplane shall be manufactured in accordance with the latest FAA approved revision of 

“Master Drawing List”, Document No. 13750, or other FAA approved data. 

 Serial Nos. Eligible 0001 and on. 

Data Pertinent to All Models

 Reference Datum 100 inches in front of the forward face of firewall bulkhead 

 Leveling Means Door sill and leveling points as defined in AFM

 Certification Basis Model SR20: 14 CFR Part 23 of the Federal Aviation Regulations effective February 1, 1965, as 

amended by 23-1 through 23-47, except as follows: 

 14 CFR 23.573, 23.575, 23.611, 23.657, 23.673 through Amendment 23-48;

14 CFR 23.783, 23.785, 23.867, 23.1303, 23.1307, 23.1309, 23.1311, 23.1321, 23.1323, 

23.1329, 23,1361, 23.1383, 23.1401, 23.1431, 23.1435 through Amendment 23-49; 

14 CFR 23.3, 23.25, 23.143, 23.145, 23.155, 23.1325, 23.1521, 23.1543, 23.1555, 23.1559, 

23.1567, 23.1583, 23.1585, 23.1589 through Amendment 23-50; 

14 CFR 23.777, 23.779, 23.901, 23.907, 23.955, 23.959, 23.963, 23.965, 23.973, 23.975, 

23.1041, 23.1091, 23.1093, 23.1107, 23.1121, 23.1141, 23.1143, 23.1181, 23.1191, 23.1337 

through Amendment 23-51; 

14 CFR 23.1305 through Amendment 23-52 

Noise: 14 CFR Part 36 dated December 1, 1969 as amended by 36-1 through 36-21. 
2-10

Figure 2-16. Weight and balance report.

Item C.G. Arm (in) Moment (lb-in/1,000)

Standard empty weight

Optional equipment

Special installation

Paint

Unusable fuel

Basic empty weight

 1,876.0

 1.2

 6.2

 —

 30.0

 1,913.4

 67,798.6

 16.7

 257.3

 —

 1,380.0

 69,452.6

 36.1

 13.9

 41.5

 —

 46.0

Aircraft Serial No. 18259080 FAA Registration No. N42565 Date: 4-22-09

Weight (lb) × 

Figure 2-15. Sample excerpt from TCDS A00009CH (continued).

Note 2. All placards specified in the latest FAA approved revisions of the PILOTS OPERATING HANDBOOK AND 

FAA APPROVED AIRPLANE FLIGHT MANUAL FOR THE CIRRUS SR20", document numbers 11934-001, 

11934-002, 11934-003 or 11934-004 and the latest FAA approved revisions of the "PILOT'S OPERATING 

HANDBOOK AND FAA APPROVED AIRPLANE FLIGHT MANUAL FOR THE CIRRUS SR22" document 

numbers 13772-001 or 13772-002 must be displayed in the airplane in the appropriate locations. 

Note 3. FAA approved Airworthiness Limitations are included in Section 4 of the Airplane Maintenance Manual (AMM) 

Document No. 12137-001 for model SR20, and 13773-001 for model SR22. 

Note 4. Exterior colors are limited to those specified in the latest FAA accepted revision of the Airplane Maintenance 

Manual (AMM) Document No. 12137-001 for model SR20, and 13773-001 for model SR22. 

Note 5. Major structural repairs must be accomplished in accordance with FAA approved Cirrus 

Design repair methods or other methods approved by the FAA. 

Note 6. For Model SR22 S/N 0002 thru 2333, 2335 thru 2419, and 2421 thru 2437 a maximum landing weight exists 

along the line between 141.4 inches at 3210 lbs and 142.7 inches at 3400 lbs. 

-- END -- 

Data Pertinent to all Models

The location of the datum is specified and is described, for 

example, as: 100 inches in front of the forward face of the 

firewall bulkhead.

Manufacturer-Furnished Information

When an aircraft is initially certificated, its empty weight and 

EWCG are determined and recorded in the weight and balance 

record, such as the one in Figure 2-16. Notice in this figur 

that the moment is expressed as “Moment (lb-in/1,000).” This 

is a moment index, which means that the moment, a very 

large number, has been divided by 1,000 to make it more 

manageable. Chapter 4, Light Sport Aircraft—Weight and 

Balance Control, discusses moment indices in more detail.

The aircraft is furnished with an equipment list, specifies

all the required equipment and all equipment approved for 

installation in the aircraft. The weight and arm of all installed 

equipment are included on the list and checked prior to the 

aircraft leaving the factory.

When an aircraft mechanic or repairman adds or removes 

any item on the equipment list, he or she must change the 

weight and balance record to indicate the new empty weight 

and EWCG, and the equipment list is revised to show 

which equipment is actually installed. Figure 2-17 is from a 

comprehensive equipment list that includes all of the items 

of equipment approved for this particular model of aircraft. 

The pilot’s operating handbook (POH) for each individual 

aircraft includes an aircraft-specific equipment list of the 

items from this master list. When any item is added to or 

removed from the aircraft, its weight and arm are determined 

in the equipment list and used to update the weight and 

balance record. The POH and airplane flight manual (AFM) 

also contain CG moment envelopes and loading graphs. 

Examples of the use of these helpful graphs are given in 

Chapter 5, Single-Engine Aircraft Weight and Balance 

Computations.

In addition to the weight and balance report and equipment 

list, the manufacturer also provides a CG range chart. The 

CG range can be found in text form in the TCDS. The CG 

range chart is furnished in the POH.
2-11

Figure 2-17. Excerpt from a typical comprehensive equipment list.
2-12

Figure 2-17. Excerpt from a typical comprehensive equipment list (continued).
3-1

Introduction

Chapter 2, Weight and Balance Theory, explained the theory 

of weight and balance and gave examples of the way the 

center of gravity (CG) could be found for a lever loaded 

with several weights. In this chapter, the practical aspects 

of weighing an airplane and locating its CG are discussed. 

Formulas are introduced that allow the CG location to be 

measured in inches from various datum locations and in 

percentage of the mean aerodynamic chord (MAC).

Weighing the Aircraft and 

Determining the Empty Weight 

Center of Gravity

Chapter 3
3-2

Figure 3-1. Platform scales.

Requirements

Regulations do not require periodic weighing of privately 

owned and operated aircraft. Such aircraft are usually 

weighed when originally certificatedor after major alterations 

that can affect the weight and balance. The primary purpose 

of aircraft weight and balance control is safety. Manufacturers 

conduct extensive flight tests to establish loading limits 

for their aircraft because limit information is critical for 

safe flight. A secondary purpose is to aid efficiency during 

flight. Overloading of the aircraft is not the only concern; 

the distribution of the weight is important also. The aircraft 

has CG limits, and any loading that places the CG outside 

the established limits seriously impairs controllability of 

the aircraft.

Weight and balance is of such vital importance that each 

Federal Aviation Administration (FAA) certificated 

mechanic or repairman maintaining an aircraft must be fully 

aware of his or her responsibility to provide the pilot with 

current and accurate information for the actual weight of the 

aircraft and the location of the CG. The pilot in command 

(PIC) is responsible for knowing the weight of the load, CG, 

maximum allowable weight, and CG limits of the aircraft. 

The weight and balance report must include an equipment 

list showing weights and moment arms of all required and 

optional items of equipment included in the certificate 

empty weight.

Weight and balance records used in accounting for and 

correcting the CG location are reliable for only limited 

periods of time. For this reason, periodic aircraft weighing is 

desirable. An aircraft should be reweighed and a new weight 

and balance record should be started after the aircraft has 

undergone extensive repair or major alteration, when the pilot 

reports unsatisfactory flight characteristics (e.g., nose or tail 

heaviness), and when recorded weight and balance data are 

suspected to be in error.

Repairs and alterations are major sources of weight change. 

The airframe and powerplant (A&P) FAA-certificated 

mechanic or repairman who is responsible for making any 

repair or alteration must:

1. Establish by computation that the authorized weight 

and CG limits as shown in the type certificate data 

sheet (TCDS) and aircraft specifications are not 

exceeded, and

2. Record the new empty weight center of gravity 

(EWCG) data in the current approved aircraft flight

manual or issued operating limitations.

When an aircraft has undergone extensive repair or major 

alteration, it should be reweighed and a new weight and 

balance record started. The A&P FAA-certificated mechanic 

or repairman responsible for the work must provide the pilot 

with current and accurate aircraft weight information and 

location of the EWCG.

Equipment for Weighing

Weighing aircraft with accurately calibrated scales is the only 

sure method of obtaining an accurate empty weight and CG 

location. The two basic types of scales used to weigh aircraft 

are platform and load cell.

Platform scales [Figure 3-1] or ramp wheel scales 

[Figure 3-2] (usually a form or modified version of the 

platform scale) are low profi e, easy to handle, safe, and 

reliable. Tow or push the aircraft wheels or skids onto the 

scale pad at ground level. With one scale per wheel, each 

device should be capable of measuring up to at least 60,000 

pounds since the weight on each wheel rarely exceeds this 

figure

Load cell scales [Figure 3-3] are also a reliable means to 

weigh aircraft and are typically cheaper than the platform 

type.Using load cell scales allows for the aircraft to be set 

up and weighed in its level flight attitude. With this method, 

the aircraft is placed on jacks with electronic load cells 

placed between the jack and the jack pad on the aircraft. 

The aircraft is raised on the jacks until the wheels or skids 

are off the floor and the aircraft is in a level flight attitude. 

The weight measured by each load cell is indicated on the 

control panel. Jacking an aircraft off the ground from all 

load points can be an inconvenience, as well as a safety 

risk, which some operators would rather avoid by opting for 

more expensive—but simpler to use—platform equipment. 

In addition, weighing with platform scales typically takes 

only one-third of the time needed to weigh with load cells.
3-3

Figure 3-2. Ramp scales.

Figure 3-3. Load cell scales.

All scales for aviation use, manual or electronic, must be 

protected when stored or shipped, and they must be checked 

periodically for accuracy. The maximum recognized period 

between calibration checks is 12 months; however, this period 

may be reduced by an airworthiness authority dependent on 

the conditions of use. Scales in daily use may require a shorter 

interval and/or testing to determine the continued accuracy 

of the unit. Scales should be returned to the manufacturer for 

proper calibration and testing. 

Preparation for Weighing

In general, weight procedures may vary with the aircraft 

and types of weight equipment employed. The weighing 

procedure contained in the manufacturer’s maintenance 

manual should be followed for each particular aircraft. The 

major considerations in preparing an aircraft for weighing 

are described in the following paragraphs.

Scale Preparation

Mechanical and electronic scales shall be inspected prior to 

use and set to zero. This is done by adding and removing 

a weight, then rechecking for zero. This process should be 

repeated until a steady zero setting is obtained. The scales 

should be located in the same environment in which they 
3-4

Figure 3-4. Standard fuels and weights with temperatures of 32 °F 

and 59 °F.

Weighing Point

AVGAS (Aviation Gasoline)

JET A & A-1

Water

Oil

6.14

6.75

8.35

7.50

6.01

6.68

8.33

7.43

Pounds per U.S. Gallon

32 °F 59 °F

are to be used and allowed to come up to temperature at 

least 2 hours prior to use. Scales should not be used in 

temperature extremes below 40 °F or above 100 °F unless the 

scale is specifically designed for use in those temperatures. 

Electronic scales are very sensitive and, if subjected to 

freezing temperatures, the liquid displays may be damaged 

beyond use.

Weigh Clean Aircraft Inside Hangar

The aircraft should be weighed inside a hangar where wind 

cannot blow over the surface and cause fluctuating or false 

scale readings. The aircraft should be clean inside and out, 

with special attention paid to the bilge area to ensure that no 

water or debris is trapped there. The outside of the aircraft 

should be as free as possible of all mud and dirt.

Equipment List

All of the required equipment must be properly installed, and 

there should be no equipment installed that is not included in 

the equipment list. If such equipment is installed, the weight 

and balance record must be corrected to indicate it.

Ballast

All required permanent ballasts must be properly secured in 

place. All temporary ballasts must be removed.

Standard Weights

Standard weights are established weights for numerous items 

involved in weight and balance computations. These weights 

should not be used if actual weights are available. Some of 

the standard weights are listed in Figure 3-4.

Note the difference in weight as temperatures change. 

Although this change is a very small amount per gallon, it could 

end up in a significanttotal weight gain when dealing with large 

quantities of fluids,such as those found in commercial aircraft.

Draining the Fuel

Drain fuel from the tanks in the manner specified by the 

aircraft manufacturer. If there are no specific instructions, 

drain the fuel until the fuel quantity gauges read empty when 

the aircraft is in level-flight attitude. Any fuel remaining in 

the system is considered residual or unusable fuel and is part 

of the aircraft empty weight.

The amount of residual fuel and its arm are normally found 

in Note 1 in the section of the Type Certificate Data Sheets 

(TCDS), “Data pertaining to all Models.” For additional fuel 

capacity information, see Chapter 2, Weight and Balance 

Theory.

If it is not feasible to drain the fuel, the tanks can be topped 

off to be sure of the quantity they contain and the aircraft 

weighed with full fuel. After weighing is complete, the 

weight of the fuel and its moment are subtracted from those 

of the aircraft as weighed. To correct the empty weight for 

the residual fuel, add its weight and moment.

When computing the weight of the fuel (e.g., a tank full of 

jet fuel), measure its specific gravity (sg) with a hydrometer 

and multiply it by 8.345 (the nominal weight of 1 gallon of 

pure water whose sg is 1.0). If the ambient temperature is 

high and the jet fuel in the tank is hot enough for its specific

gravity to reach 0.81 rather than its nominal sg of 0.82, the 

fuel actually weighs 6.76 pounds per gallon rather than its 

normal weight of 6.84 pounds per gallon.

Oil

The empty weight for aircraft certificated under the Civilian 

Air Regulations (CAR) part 3 does not include the engine 

lubricating oil. The oil must either be drained before the 

aircraft is weighed, or its weight must be subtracted from 

the scale readings to determine the empty weight. To weigh 

an aircraft that does not include the engine lubricating oil 

as part of the empty weight, place it in level flight attitude, 

then open the drain valves and allow the oil to drain out. 

Any remaining is undrainable oil and is part of the empty 

weight. Aircraft certificated under Title 14 of the Code of 

Federal Regulations (14 CFR) parts 23 and 25 include full 

oil as part of the empty weight. If it is impractical to drain the 

oil, the reservoir can be filled to the specified level and the 

weight of the oil computed at 7.5 pounds per gallon. Then, 

its weight and moment are subtracted from the weight and 

moment of the aircraft as weighed. The amount and arm of 

the undrainable oil are found in Note 1 of the TCDS, and 

this must be added to the empty weight.

Other Fluids

The hydraulic fluid reservoir and all other reservoirs 

containing fluids required for normal operation of the aircraft 

should be full. Fluids not considered to be part of the empty 

weight of the aircraft are potable (drinkable) water, lavatory 

precharge water, and water for injection into the engines.
3-5

Figure 3-5. Locating the CG of an airplane relative to the datum.

Weighing point Scale reading (lb) TARE weight (lb) Net weight (lb) Arm (in) Moment (lb-in) CG

46.0

46.0

−32.0

32.8

Right side

Left side

Nose

Total

16

16

8

830

836

340

2,006

38,180

38,456

−10,880

65,756

846

852

348

Configuration of the Aircraf

Consult the aircraft service manual regarding position of 

the landing gear shock struts and the control surfaces for 

weighing. When weighing a helicopter, the main rotor must 

be in its correct position.

Jacking the Aircraft

Aircraft are often weighed by rolling them onto ramps in 

which load cells are embedded. This eliminates the problems 

associated with jacking the aircraft off the ground. However, 

many aircraft are weighed by jacking the aircraft up and then 

lowering them onto scales or load cells.

Extra care must be used when raising an aircraft on jacks for 

weighing. If the aircraft has spring steel landing gear and it 

is jacked at the wheel, the landing gear will slide inward as 

the weight is taken off of the tire, and care must be taken to 

prevent the jack from tipping over.

For some aircraft, stress panels or plates must be installed before 

the aircraft is raised with wing jacks to distribute the weight 

over the jack pad. Be sure to follow the recommendations 

of the aircraft manufacturer in detail anytime an aircraft is 

jacked. When using two wing jacks, take special care to raise 

them simultaneously, keeping the aircraft so it does not slip 

off the jacks. As the jacks are raised, keep the safety collars 

screwed down against the jack cylinder to prevent the aircraft 

from tilting if one of the jacks should lose hydraulic pressure.

Leveling the Aircraft

When an aircraft is weighed, it must be in its level fligh 

attitude so that all of the components are at their correct 

distance from the datum. This attitude is determined by 

information in the TCDS. Some aircraft require a plumb line 

to be dropped from a specified location so that the point of the 

weight (the bob) hangs directly above an identifiable point. 

Others specify that a spirit level be placed across two leveling 

lugs, often special screws on the outside of the fuselage. Other 

aircraft call for a spirit level to be placed on the upper door sill.

Lateral level is not specified for all light aircraft, but 

provisions are normally made on helicopters for determining 

both longitudinal and lateral level. This may be done by 

built-in leveling indicators, or by a plumb bob that shows the 

conditions of both longitudinal and lateral level.

The actual adjustments to level the aircraft using load cells 

are made with the jacks. When weighing from the wheels, 

leveling is normally done by adjusting the air pressure in the 

nosewheel shock strut.

Safety Considerations

Special precautions must be taken when raising an aircraft 

on jacks.

1. Stress plates must be installed under the jack pads if 

the manufacturer specifies them

2. If anyone is required to be in the aircraft while it is 

being jacked, there must be no movement.

3. The jacks must be straight under the jack pads before 

beginning to raise the aircraft.

4. All jacks must be raised simultaneously and the safety 

devices are against the jack cylinder to prevent the 

aircraft tipping if any jack should lose pressure. Not 

all jacks have screw down collars, some use drop pins 

or friction locks.

Determining the CG

When the aircraft is in its level flight attitude, drop a plumb 

line from the datum and make a mark on the hangar floor

below the tip of the bob. Draw a chalk line through this point 

parallel to the longitudinal axis of the aircraft.

Then, draw lateral lines between the actual weighing 

points for the main wheels, and make a mark along the 

longitudinal line at the weighing point for the nosewheel 

or the tailwheel. These lines and marks on the floor allow 

accurate measurements between the datum and the weighting 

points to determine their arms.

Determine the CG by adding the weight and moment of 

each weighing point to determine the total weight and total 

moment. Then, divide the total moment by the total weight 

to determine the CG relative to the datum. As an example of 

locating the CG with respect to the datum, which in this case 

is the firewall, consider the tricycle landing gear airplane as 

detailed in the Figure 3-5 table and illustrated in Figure 3-6.

When the airplane is on the scales with the parking brakes 
3-6

Figure 3-9. Determining the CG with datum forward of an airplane 

with nosewheel landing gear.

CG 

F × L

W = D − 

= 128 − 

= 114.8

( )

340 × 78

2,006 ( )

Figure 3-8. The datum is 100 inches forward of the wing root 

leading edge.

D = 128.0

100.0 13.2

50

 114.8

Datum

L = 78.0

Figure 3-10. The datum is aft of the main wheels at the wing 

trailing edge.

153

D = 75

 88.2

Datum

L = 78.0

Figure 3-7. Determining the CG.

CG Total moment

Total weight =

65,756

2,006 =

= 32.8 inches behind the datum

Figure 3-6. The datum is located at the firewall.

32 46.0

 32.8

Datum

off, place chocks around the wheels to keep the airplane 

from rolling. Subtract the weight of the chocks, called tare 

weight, from the scale reading to determine the net weight 

at each weighing point. Multiply each net weight by its arm 

to determine its moment, and then determine the total weight 

and total moment. The CG is determined by dividing the total 

moment by the total weight. [Figure 3-7]

The airplane illustrated in Figures 3-5 and 3-6 has a net 

weight of 2,006 pounds, and its CG is 32.8 inches behind 

the datum.

EWCG Formulas

A chart such as the one in Figure 3-5 helps the pilot visualize 

the weights, arms, and moments when solving an EWCG 

problem, but it is quicker to determine the EWCG by using 

formulas and an electronic calculator. The use of a calculator 

for solving these problems is described in Chapter 8, Use of 

Computers in Weight and Balance Computations.

There are four possible conditions and their formulas that 

relate the location of CG to the datum. Notice that the 

formula for each condition first determines the moment of 

the nosewheel or tailwheel and then divides it by the total 

weight of the airplane. The arm thus determined is then added 

to or subtracted from the distance between the main wheels 

and the datum, distance D.

Datum Forward of the Airplane—Nosewheel 

Landing Gear

The datum of the airplane in Figure 3-8 is 100 inches forward 

of the leading edge of the wing root or 128 inches forward 

of the main-wheel weighing points. This is distance (D). The 

weight of the nosewheel (F) is 340 pounds, and the distance 

between main wheels and nosewheel (L) is 78 inches. The 

total weight of the airplane (W) is 2,006 pounds. Determine 

the CG by using the formula in Figure 3-9.

The CG is 114.8 inches aft of the datum. This is 13.2 inches 

forward of the main-wheel weighing points, which proves 

the location of the datum has no effect on the location of the 

CG if all measurements are made from the same location.

Datum Aft of the Main Wheels—Nosewheel 

Landing Gear

The datum of some aircraft may be located aft of the main 

wheels. The airplane in this example is the same one just 

discussed, but the datum is at the intersection of the trailing 

edge of the wing with the fuselage. The distance (D) between 

the datum of the airplane in Figure 3-10 and the main-wheel 
3-7

Figure 3-13. Determining the CG with datum forward of the main 

wheels in an airplane with tailwheel landing gear.

CG 

R × L

W = D + 

= 7.5 + 

= 19.7

( )

67 × 222

1,218 ( )

Figure 3-11. Determining the CG with datum aft of the main wheels 

of an airplane with nosewheel landing gear.

CG 

F × L

W = − D + 

= − 75 + 

= −88.2

( )

340 × 78

2,006 ( )

Figure 3-12. The datum of this tailwheel airplane is the wing root 

leading edge.

D = 7.5

Datum

229.5

L = 222.0

19.7

Datum

L = 222.0

D = −80 142

 67.8

Figure 3-14. The datum is aft of the main wheels, at the intersection 

of the wing trailing edge and the fuselage.

CG 

R × L

W = −D + 

= −80 + 

= −67.8

( )

67 × 222

1,218 ( )

Figure 3-16. Determining the CG with datum aft of the main wheels 

in an airplane with tailwheel landing gear.

weighing points is 75 inches, the weight of the nosewheel 

(F) is 340 pounds, and the distance between main wheels 

and nosewheel (L) is 78 inches. The total net weight of the 

airplane (W) is 2,006 pounds.

The location of the CG may be determined by using the 

formula in Figure 3-11.

The CG location is a negative value, which means it is 88.2 

inches forward of the datum. This places it 13.2 inches 

forward of the main wheels, exactly the same location as 

when it was measured from other datum locations.

Location of Datum

The location of the datum is not important, but all 

measurements must be made from the same location.

Datum Forward of the Main Wheels—Tailwheel 

Landing Gear

Locating the CG of a tailwheel airplane is done in the same 

way as locating it for a nosewheel airplane except the formula 

is .

The distance (D) between the datum of the airplane in 

Figure 3-12 and the main-gear weighing points is 7.5 inches, 

the weight of the tailwheel (R) is 67 pounds, and the distance 

(L) between the main-wheel and the tailwheel weighing points 

is 222 inches. The total weight of the airplane (W) is 1,218 

pounds. Determine the CG by using the formula in Figure 3-13. 

R × L

W

The CG is 19.7 inches behind the datum.

Datum Aft of the Main Wheels—Tailwheel 

Landing Gear

The datum of the airplane in Figure 3-14 is located at the 

intersection of the wing root trailing edge and the fuselage.

This places the arm of the main gear (D) at –80 inches. The 

net weight of the tailwheel (R) is 67 pounds, the distance 

between the main wheels and the tailwheel (L) is 222 inches, 

and the total net weight (W) of the airplane is 1,218 pounds.

Since the datum is aft of the main wheels, use the formula 

found in Figure 3-15. 

The CG is 67.8 inches forward of the datum or 12.2 inches 

aft of the main-gear weighing points. The CG is in exactly 

the same location relative to the main wheels, regardless of 
3-8

Figure 3-16. The MAC is the chord drawn through the geographic 

center of the plan area of the wing.

CG MAC

Trailing edge mean 

aerodynamic chord

Leading edge mean 

aerodynamic chord

TEMAC

LEMAC

Datum

where the datum is located.

Center of Gravity (CG) and Mean 

Aerodynamic Chord (MAC)

Center of Gravity (CG)

In addition to overloading or drastically reducing the 

aircraft’s weight, the distribution of weight is also a concern. 

When aircraft equipment is changed, the person making the 

equipment change must make an entry on the equipment list 

indicating items added, removed, or relocated; the date of 

the change; and the person’s name and certification number 

in the aircraft’s maintenance records.

For the purpose of weight and balance computations, the CG 

of an airplane is an imaginary point about which the noseheavy (–) moments and tail-heavy (+) moments are exactly 

equal in magnitude. If suspended from the CG point, the 

aircraft would have no tendency to rotate nose up or nose 

down. The CG of the loaded aircraft can range fore and aft 

within certain limits that are determined during the fl ght test 

for type certific tion. These limits are the most forward- and 

rearward-loaded CG positions at which the aircraft meets the 

performance and flight characteristics required by the FAA. 

Any loading that places the CG outside the limits for a 

particular aircraft seriously impairs the pilot’s ability to 

control the aircraft. For example, it is more difficult to take 

off and gain altitude in a nose-heavy aircraft, and the aircraft 

tends to drop its nose when the pilot reduces throttle. It also 

requires a higher speed to land safely. A tail-heavy aircraft is 

more susceptible to stalling at low speed, which is a concern 

during the landing approach.

On small airplanes and on all helicopters, the CG location 

is identified as being a specific number of inches from the 

datum. The CG range is identified in the same way. On larger 

airplanes, from private business jets to large jumbo jets, the 

CG and its range are typically identified in relation to the 

width of the wing.

Mean Aerodynamic Chord (MAC)

The width of the wing, or straight-line distance from the 

leading edge to the trailing edge, on an airplane is known as 

the chord. If the leading edge and the trailing edge of a wing 

are parallel, the chord is equal at all points along the entire 

length of the wing. The average length of the chord, or MAC, 

of a tapered wing is more complicated to define. The MAC, 

as seen in Figure 3-16, is the chord of an imaginary airfoil 

that has the same aerodynamic characteristics as the actual 

airfoil. You can also think of it as the chord drawn through 

the geographic center of the plan area of the wing.

Usually listed in the aircraft’s TCDS when it is required for 

weight and balance computations, the MAC is established 

by the manufacturer, defining its leading edge (LEMAC) 

and trailing edge (TEMAC) in terms of inches from the 

datum. [Figure 3-16] The CG location and various limits 

are then expressed in percentage of the chord or percent 

MAC. In order to relate the percent MAC to the datum, all 

weight and balance information includes two items: the MAC 

length in inches and the location of the LEMAC in inches 

from the datum. For the purpose of simplicity, most lightaircraft manufacturers express the CG range in inches from 

the datum; transport-category aircraft CGs are expressed in 

percent MAC.

The relative positions of the CG and the aerodynamic 

center of lift of the wing have critical effects on the flight

characteristics of the aircraft. Consequently, relating the 

CG location to the chord of the wing is convenient from 

a design and operations standpoint. Normally, an aircraft 

has acceptable fl ght characteristics if the CG is located 

somewhere near the 25 percent average chord point. This 

means the CG is located one-fourth of the distance back from 

the LEMAC to the TEMAC. Such a location places the CG 

forward of the aerodynamic center for most airfoils.

The weight and balance data of the airplane in Figure 3-17 states 

that the MAC is from stations 144 to 206 and the CG is 

located at station 161.

MAC = TEMAC – LEMAC

MAC = 206" – 144" 

MAC = 62"
3-9

Figure 3-18. Formula for determining the CG expressed in percent 

MAC.

CG inches % MAC = Distance aft of LEMAC × 100

MAC

=

= 27.4

17 × 100

62

Figure 3-17. Large aircraft weight and balance calculation diagram.

CG = 27.4%

CG = 161

MAC

Datum

LEMAC

TEMAC

CG = 27.4%

CG = 161

MAC

 Station 144" Station 206"

Figure 3-19. Formula for determining the CG in inches from the 

datum.

CG inches from datum = LEMAC + MAC × CG % MAC

100

= 144 + 

= 160.9

62 × 27.4

100

CG = 161"

In order to find the percent MAC, firs determine the distance 

of the CG from LEMAC.

CG – LEMAC = Distance from LEMAC

161" – 144" = 17"

The location of the CG expressed in percent MAC is 

determined using the formula found in Figure 3-18. The CG 

of the airplane is located at 27.4 percent MAC.

It is sometimes necessary to determine the location of the 

CG in inches from the datum when its location in percent 

MAC is known.

The CG of the airplane is located at 27.4 percent MAC.

MAC = 206 – 144 = 62

LEMAC = station 144

Determine the location of the CG in inches from the datum 

by using the formula found in Figure 3-19.

The CG of this airplane is located at station 160.9 inches aft 

of the datum. It is important for longitudinal stability that the 

CG be located ahead of the center of lift of a wing. Since the 

center of lift is expressed as percent MAC, the location of 

the CG is expressed in the same terms.
3-10
4-1

Introduction

This chapter discusses the weight and balance procedures 

for light sport aircraft (LSA) that differ from conventional 

aircraft, specifically weight-shift control (WSC) aircraft 

(also called trikes), powered parachutes, and amateur-built 

LSA. [Figure 4-1]

Light Sport Aircraft Weight 

and Balance Control

Chapter 4
4-2

Figure 4-1. Examples of light sport aircraft (from top left, clockwise): weight-shift control, powered parachute, glider, airplanes, hot 

air balloon, and amateur-built LSA.

LSA Definition of Term

LSA is a category of simple, very basic, small, light-weight, 

low-performance aircraft, other than a helicopter or poweredlift, and a classification of aircraft specific to the United 

States. The Federal Aviation Administration (FAA) defines

LSA as an aircraft with a maximum gross takeoff weight of 

not more than 1,320 pounds (600 kg) for aircraft not intended 

for operation over water, or 1,430 pounds (650 kg) for aircraft 

intended for operation over water; a maximum airspeed in 

level flight of 120 knots (220 kilometers per hour (km/h); 140 

miles per hour (mph)); a maximum stall speed of 45 knots 

(83 km/h; 52 mph); a maximum seating capacity of no more 

than two persons (including the pilot); fixed undercarriage 

and fixed-pitch or ground-adjustable propeller; and a single 

reciprocating engine (if powered). 

An aircraft that qualifies as LSA may be operated by the 

holder of a sport pilot certificate, whether registered as LSA 

or not. Pilots with a private, recreational, or higher pilot 

certificate may also fly LSA, even if their medical certificate 

have expired, as long as they have a valid driver’s license to 

prove that they are in good enough health to fly. LSA also 

have less restrictive maintenance requirements and may be 

maintained and inspected by traditionally certificated aircraft 

maintenance technicians (AMTs) or by individuals holding 
4-3

Figure 4-2. CG of a trike.

CG

Aircraft CG

• CG is under wing hang point for level flight.

• Center of lift is directly above CG, above the

 wing hang point for level flight.

Thrust line is typically 

designed to be at vertical 

CG.

Fuel tank

a Repairman: Light Sport certificate, and (in some cases) by 

their pilots and/or owners.

Weight and Balance

Aircraft such as balloons, powered parachutes, and WSC do 

not require weight and balance computations because the load 

is suspended below the lifting mechanism. The CG range in 

these types of aircraft is such that it is difficult to exceed CG 

limits. For example, the rear seat position and fuel of a WSC 

aircraft are as close as possible to the hang point with the 

aircraft in a suspended attitude. Thus, load variations have 

little effect on the CG. This also holds true for lighter-thanair aircraft, such as a balloon basket or gondola. While it is 

difficult to exceed CG limits in these aircraft, pilots should 

never overload an aircraft, as doing so may cause structural 

damage and/or failures. 

Weight affects performance; therefore, pilots should calculate 

weight and remain within the manufacturer’s established 

limits at all times.

WSC Aircraft

WSC aircraft are one- and two-place aircraft that exceed the 

criteria of an ultra-light vehicle but do meet the criteria of an 

LSA. The definition for WSC can be found in 14 CFR part 1. 

A WSC aircraft used for sport and private pilot flying must 

be registered with an FAA N-number, have an airworthiness 

certificate, a pilot’s operating handbook (POH), and/or 

limitations with a weight and loading document aboard. 

As mentioned earlier, WSC aircraft are commonly called 

trikes. These aircraft have few options for loading because 

they lack places to put useful load items. One-place trikes 

have only one seat and a fuel tank, which means the only 

variables for a flight are amount of fuel and weight of the 

pilot. Two-place trikes can accommodate a pilot and a 

passenger. This version may have a small storage bin in 

addition to the fuel tank.

The most significant factor affecting the weight and balance 

of a trike is the weight of the pilot and, if the aircraft has two 

seats, the weight of the passenger. The trike acts somewhat 

like a single, main-rotor helicopter because the weight of the 

aircraft hangs like a pendulum under the wing. Figure 4-2

shows a two-place trike, in which the mast and the nose strut 

come together slightly below the wing attach point. When 

the trike is in flight, the weight of the aircraft hangs from 

the wing attach point. The weight of the engine and fuel is 

behind this point, the passenger is almost directly below this 

point, and the pilot is forward of this point. The balance of 

the aircraft is determined by how all these weights compare. 

The wing attach point, with respect to the wing keel, is an 

adjustable location. The attach point is moved slightly forward 

or slightly aft, depending on the weight of the occupants. For 
4-4

Figure 4-3. Wing attach point.

Front collar

Trike mast

Keel nose plate Control frame apex

Distance from nose

plate to front edge 

of front collar

Rear collar adjusted to eliminate gap when 

pivot block assembly is against front collar

501/2" to 521/2"

example, if the aircraft is flown by a heavy person, the attach 

point can be moved farther aft, bringing the wing forward to 

compensate for the change in CG. Figure 4-3 shows a closeup of the wing attach point and the small amount of forward 

and aft movement that is available.

Similar to airplanes, sailplanes, and powered parachutes, 

increasing weight creates increases in speed and descent 

rate. However, the WSC aircraft has a unique characteristic. 

Adding weight to a WSC aircraft creates more twist in the 

wing because the outboard leading edges flex more. With 

less lift at the tips, a nose-up effect is created and the trim 

speed lowers. Therefore, adding weight can increase speed 

similar to other aircraft, but reduce the trim speed because 

of the increased twist unique to the WSC aircraft. Each 

manufacturer’s make/model has different effects depending 

on the specific design. For detailed weight and balance 

information, characteristics, and operating limitations, always 

reference the specific manufacturer’s manual or POH for the 

make and model. Figure 4-4 shows an example of a weight 

and loading sheet that would be issued with a WSC aircraft. 

Every aircraft has its own weight and loading data that should 

come from the manufacturer. The example in Figure 4-4

comes from Airborne, an Australian company, named 

Airborne XT WSC aircraft. For additional information, 

refer to the Weight-Shift Control Aircraft Flying Handbook 

(FAA-H-8083-5).

Powered Parachutes

Powered parachutes have many of the same characteristics 

as WSC aircraft when it comes to weight and balance. They 

have the same limited loading, with only one or two seats, and 

a fuel tank. A powered parachute acts like a pendulum with 

the weight of the aircraft hanging beneath the inflated wing 

(parachute). The point at which the inflated wing attaches 

to the structure of the aircraft is adjustable to compensate 

for pilots and passengers of varying weights. With a 

very heavy pilot, the wing attach point would be moved 

forward to prevent the aircraft from being too nose heavy. 

Figure 4-5 illustrates the structure of a powered parachute 

and the location of the wing attachment. 

A powered parachute used for sport and private flying must 

be registered with an FAA N-number, have an airworthiness 

certificate, a POH, and/or limitations with a weight and 

balance document aboard. The aircraft must be maintained 

properly by the aircraft owner, or other qualified personnel, 

and have the aircraft logbooks available for inspection. 

Always refer to the POH for weight and balance information 

specifi to the powered parachute being fl wn. For additional 

information, refer to the Powered Parachute Flying Handbook 

(FAA-H-8083-29).

Weight and Balance Computations 

(Amateur-Built Aircraft)

A good weight and balance calculation is the keystone 

of flight testing an amateur-built aircraft. Accurately 

determining the aircraft’s takeoff weight and ensuring that 

the CG is within the aircraft’s design for each flight is critical 

to conducting a safe flight test

The aircraft should be level when weighed, spanwise and 

fore and aft in accordance with the kit manufacturer’s 

instructions, and should be in the level flight position. It 

is highly recommended that the aircraft be weighed in an 

enclosed area using three calibrated scales. Bathroom scales 

are not recommended because they are not always accurate.
4-5

Figure 4-4. Weight and loading for WSC aircraft.

Hang point

Keel roller

Rear

Front

Keel roller

adjustment holes

U bracket

The effect of changing the hang position is to change the trim speed of the aircraft. Moving the hang point forward 

increases trim speed (shorter distances to the datum). The hang point position range may be used on the 

applicable wings for the entire weight range of the aircraft. CG limits on the trike base/gondola section of the 

aircraft limits are not critical. The defined hang point position and its limits are defined in this document.

The datum point is the forward bolt on 

the nose plate (a line between the 

leading edge pivot bolts on the nose 

plate) referenced to the hang bolt at 

the top of the mast for the trike base 

(gondola).

Moving hang point position is made by 

moving the keel roller forward or 

backward along the keel or by moving 

the hang point within the U-bracket. A 

sticker is used on the keel to show the 

standard position for the keel roller. It 

lists the keel tube holes used as 

standard trim points for each model 

wing. The table below shows the 

standard keel roller position and the 

allowable range of CG position/hang 

bolt position referenced to the nose 

plate datum.

Diagram of the wing hang point construction, showing 

keel, U-bracket, and the top of the control frame.

Wing model and 

maximum takeoff 

weight

Streak 3

MTO@ 450 kg

Streak 2B

MTOW 450 kg

Cruze

MTOW 450 kg

Wizard

MTOW 430 kg

Standard trim position and range 

Distance from the forward keel roller 

bolt to forward nose plate bolt

Keel hole 2, second from front

1,340 mm +0 – 20 mm

Keel hole 3, third from front

1,360 mm ±20 mm

Keel hole 3, third from front

1,360 mm ±20 mm

Keel hole 1, at front

1,545 mm +0 – 45 mm

Distance from the hang point to forward 

nose-plate bolt

Permissible rear and foremost positions

Permissible U-bracket holes

1,293.5 mm (54.9 in) rear limit

1,353.5 mm (53.3 in) forward limit

Middle U-bracket hole only

1,413.5 mm (55.6 in) rear limit

1,373.5 mm (54.1 in) forward limit

Front and middle U-bracket hole only

1,413.5 mm (55.6 in) rear limit

1,373.5 mm (54.1 in) forward limit

Middle U-bracket hole only

1,578.5 mm (61.2 in) rear limit

1,648.5 mm (64.9 in) forward limit

All U-bracket holes permitted

Date

Empty weight

Maximum takeoff weight

Registration

Wing model

Serial number

Make and (trike) base model Airborne XT series/X series aircraft

Applicable wing models Airborne streak 3/Streak 2B/Cruze/Wizard wing
4-6

Figure 4-8. Powered parachute.
5-1

Introduction

Weight and balance data allows the pilot to determine the 

loaded weight of the aircraft and determine whether or not 

the loaded center of gravity (CG) is within the allowable 

range for the weight. See Figure 5-1 for an example of the 

data necessary for these calculations.

Single-Engine Aircraft Weight 

and Balance Computations

Chapter 5
5-2

3,200

3,000

2,800

2,600

2,400

2,200

2,000

1,800 32 34 36 38 40 42 44 46 48

Fuselage station (in)

Loaded aircraft weight (lb)

Airplane basic empty weight 1,874.0 lb, EWCG +36.1

CG range (+40.9) to (+46.0) at 3,100 lb

(+33.0) to (+46.0) at 2,250 lb or less

Straight line variation between points 

given

Empty weight CG range None

Maximum weight 3,100 lb takeoff/flight

2,950 lb landing

No. of seats 4 (2 front at +37.0)

(2 rear at +74.0)

Maximum baggage 160 lb

Area A (100 lb at +97.0)

Area B (60 lb at +116.0)

Fuel capacity 92 gal (88 gal usable); two 46 gal

integral tanks in wings at +46.6

Oil capacity 12 qt (−15)

Figure 5-1. Weight and balance data needed to determine proper 

loading of a small airplane.

100 lb

max 60 lbmax

Datum Front seats +37 Rear seats +74 Baggage B +116

Fuel +46.6 Baggage A +97

88 gal usable

EWCG 36.1

Figure 5-2. Airplane loading diagram.

Determining the Loaded Weight and CG

An important part of preflight planning is determining that 

the aircraft is loaded so its weight and CG location are within 

the allowable limits. The methods of accomplishing this are 

the manual computational method using weights, arms, and 

moments; the chart method using weight and moment indexes 

[Figure 5-2]; and the loading graph method, which eliminates 

the need for some mathematical calculations.

Manual Computational Method

The manual computational method uses weights, arms, and 

moments. It relates the total weight and CG location to a CG 

limits chart similar to those included in the Type Certificate

Data Sheet (TCDS) and the Pilot’s Operating Handbook/

Aircraft Flight Manual (POH/AFM).

A worksheet, such as the one shown in Figure 5-3, provides 

a means to record and compute pertinent weights, arms, and 

moments for all onboard fuel, personnel, equipment, cargo, 

and baggage that is not included in the aircraft’s basic empty 

weight (BEW). Figure 5-4 is a sample of a typical equipment 

list where many of the pertinent weights and moment values 

can be found. 

As part of preflight planning, fill in the blanks in the 

worksheet with the specific data for the flight. The following 

weights were used to complete the sample weight and balance 

worksheet in Figure 5-3.

Pilot ................................................................120 lb

Front seat passenger .......................................180 lb

Rear seat passenger ........................................175 lb

Fuel (88 gal) ...................................................528 lb

Baggage A......................................................100 lb

Baggage B ........................................................50 lb

Multiply each item’s weight by its arm to determine the 

moment. Then, determine the total weight and the sum of 

the moments. Divide the total moment by the total weight to 

determine the CG in inches from the datum. For this example, 

the total weight is 3,027 pounds and the CG is 43.54 inches 

aft of the datum (a negative result would have indicated a 

CG forward of the datum). 

To determine whether or not the airplane is properly loaded 

for this flight, use the CG limits chart. [Figure 5-5] Draw a 

line vertically upward from the CG of 43.54 inches and one 

horizontally to the right from the loaded weight of 3,027 

pounds. These lines cross inside the envelope, which shows 

the airplane is properly loaded for takeoff, but 77 pounds 

overweight for landing. Note that for this sample chart, the 

envelope is defined by the solid black line that indicates 

CG limits at or below the maximum weight for takeoff 

and landing. There is an additional region identified by a 

segmented black line that includes weights suitable only for 

takeoff. It is important to note these subtle differences as they 

may or may not be found in every POH/AFM.
5-3

Item Weight (3,100 lb max.) Arm (in) Moment (lb-in) CG (inches from datum)

67,651.4

+43.54

Airplane

Front seats

Rear seats

Fuel (88 gal usable)

Baggage A (100 max.)

Baggage B (60 max.)

1,874 36.1

 37.0

 74.0

 46.6

 97.0

 116.0

300

175

528

100

50

3,027

11,100.0 

12,950.0 

 24,604.8 

9,700.0 

5,800.0 

131,806.2 

x =

Figure 5-3. Loading schedule chart derived from loading problem.

* Indicates total weight/arm for all subcomponents

Figure 5-4. Typical equipment list.
5-4

3,200

3,100

3,000

2,950

2,900

2,800

2,700

2,600

2,500

2,400

2,300

2,200

2,100

2,000

1,900

1,800

1,450

1,400

1,350

1,300

1,250

1,200

1,150

1,100

1,050

1,000

950

900

850

+32 +34 +36 +38 +40 +42 +44 +46

850 900 950 1,000 1,050 1,100 1,150 1,200

Aircraft CG location

Inches aft of datum (Sta. 0.0)

Loaded aircraft weight (lb)Loaded aircraft weight (kg)

Aircraft CG location

Millimeters aft of datum (Sta. 0.0)

CG limits

 Takeoff and landing

 Takeoff only

 If takeoff weight is more than 

 landing weight of 2,950 pounds, 

 allow flight time for fuel burn-off 

 to 2,950 pounds before landing

Figure 5-5. CG limits chart from a typical POH.

Loading Graph Method

The charts and graphs found in the POH/AFM can help 

simplify and expedite the preflight weight and balance 

computation process. Some use a loading graph and moment 

indexes rather than the arms and moments. These charts 

eliminate the need for calculating moments and make 

computations quicker and easier. [Figure 5-6]

Moment Indexes

Moments determined by multiplying the weight of each 

component by its arm result in large numbers that are awkward 

to calculate and can become a source of mathematical error. 

To eliminate these large numbers, moment indexes are 

sometimes used. The moment is divided by a reduction 

factor, such as 100 or 1,000, to get the moment index. The 

loading graph provides the moment index for each component 

to avoid mathematical calculations. The CG envelope uses 

moment indexes rather than arms and moments.

The CG limits envelope is the enclosed area on a graph of 

the airplane loaded weight and the CG location. If lines 

drawn from the weight and CG cross within this envelope, 

the airplane is properly loaded.

Loading Graph

Figure 5-6 is a typical loading graph taken from the POH 

of a modern four-place airplane. It is a graph of load weight 

and load moment indexes. Diagonal lines for each item 

relate the weight to the moment index without having to use 

mathematical calculations.

Compute Weight and Balance Using the Loading 

Graph

To compute the weight and balance using the loading graph 

in Figure 5-6, make a loading schedule chart like the one 

in Figure 5-7. In Figure 5-6, follow the horizontal line for 

300 pounds load weight to the right until it intersects the 

diagonal line for pilot and front passenger. From this point, 

drop a line vertically to the load moment index along the 

bottom to determine the load moment for the front seat 

occupants. This is 11.1 lb‐in divided by 1,000. Record it in

the loading schedule chart. Determine the load moment for 

the 175 pounds of rear seat occupants along the diagonal for 

second row passengers or cargo. This is 12.9; record it in the 

loading schedule chart. 

Determine the load moment for the fuel and the baggage 

in areas A and B in the same way and enter them all in the 

loading schedule chart. The maximum fuel is marked on 

the diagonal line for fuel in terms of gallons or liters. The 

maximum is 88 gallons of usable fuel. The total capacity is 92 

gallons, but in our example, 4 gallons are unusable and have 

already been included in the empty weight of the aircraft. The 

weight of 88 gallons of fuel is 528 pounds and its moment 

index is 24.6. The 100 pounds of baggage in area A has a 

moment index of 9.7 and the 50 pounds in area B has an 

index of 5.8. Enter all of these weights and moment indexes 

in the loading schedule chart and add all of the weights and 

moment indexes to determine the totals. 

Transfer totals to the CG moment envelope in Figure 5-8. 

The CG moment envelope is an enclosed area on a graph 

of the airplane loaded weight and loaded moment. If lines 

drawn from the weight and loaded moment cross within 

this envelope, the airplane is properly loaded. The loading 

schedule from the example in Figure 5-7 shows that the total 

weight of the loaded aircraft is 3,027 pounds, and the loaded 

airplane moment divided by 1,000 is 131.8.
5-5

550

500

450

400

350

300

250

200

150

100

50

0

250

225

200

175

150

125

100

75

50

25

0

0 5 10 15 20 25 30 35

0 50 100 150 200 250 300 350 400

Load moment/1,000 (lb-in)

Load weight (lb)Load weight (kg)

Load moment/1,000 (kg-mm)

Note: Line representing adjustable seats shows pilot and front seat passenger CG on adjustable seat positioned for an 

 average occupant. Refer to the Loading Arrangements diagram for forward and aft limits of occupant CG range. 

80 (302.8)

70 (265.0)

65 gal reduced

Loading graph

60 (227.1)

50 (189.3)

40 (151.4)

30 (113.6)

20 (75.7)

10 (37.9)

88 gal max (333.1liters)

Baggage (Area “B”) (60 lb max)

Baggage (Area “A”) (100 lb max)

Fuel (6 lb/gal)

Pilot & front passenger

2nd row passengers or cargo

Figure 5-6. Typical loading graph.

Figure 5-7. Loading schedule chart.

Item Weight Moment/1,000

Airplane (BEW) 67.7

Front seats

Rear seats

Fuel

Baggage A

Baggage B

Total

1,874

300

175

528

100

50

3,027

11.1

12.9

24.6

9.7

5.8

131.8
5-6

3,200

3,100

3,000

2,950

2,900

2,800

2,700

2,600

2,500

2,400

2,300

2,200

2,100

2,000

1,900

1,800

1,450

1,400

1,350

1,300

1,250

1,200

1,150

1,100

1,050

1,000

950

900

850

55 65 75 85 95 105 115 125 135 145 

650 850 1,050 1,250 1,450 1,650

 750 950 1,150 1,350 1,550 

Loaded aircraft moment/1,000 (lb-in)

Loaded aircraft weight (lb)Loaded aircraft weight (kg)

Loaded aircraft moment/1,000 (kg-mm)

Center of gravity

Moment envelope

 Takeoff and landing

 Takeoff only

 If takeoff weight is more than 

 landing weight of 2,950 pounds, 

 allow flight time for fuel burn-off 

 to 2,950 pounds before landing.

Figure 5-8. CG moment envelope.

Referring to Figure 5-8, draw a line vertically upward from 

131.8 on the horizontal index at the bottom of the chart and 

a horizontal line from 3,027 pounds in the left-vertical index. 

These lines intersect within the dashed area, which shows 

that the aircraft is loaded properly for takeoff, but it is too 

heavy for landing (similar to the previous example). Because 

of this, if the aircraft had to return for landing immediately 

after takeoff, it would need to fly long enough to burn 77 

pounds (slightly less than 13 gallons) of fuel to reduce its 

weight for landing.
6-1

Introduction

Weight and balance computations for small multiengine 

airplanes are similar to those discussed for single-engine 

airplanes. See Figure 6-1 for an example of weight and 

balance data for a typical light twin-engine airplane.

Multiengine Aircraft 

Weight and Balance 

Computations

Chapter 6
6-2

Figure 6-1. Typical weight and balance data for a light twin-engine 

airplane.

 

 Datum Forward face of fuselage bulkhead ahead of 

 rudder pedals

 Seats 2 at 37.0

 2 at 75.0

 1 at 113.0: 200 lb limit 

Fuel 213.4 gal (2 wing tanks, 105.0 gal each 103.0 gal

 usable at +61.0) 

 Undrainable fuel: 1.6 lb at +62

 Oil 24 quarts (12 quarts in each engine): −3.3

 Baggage Forward 100 lb limit : −15

 Aft 200 lb limit: +113

 CG Range (+38) to (+43.1) at 5,200 lb

 (+43.6) at 4,800 lb

 (+32) to (+43.6) at 4,300 lb or less

 Straight line variation between points given

 Engine 2 240-horsepower horizontally opposed engines

 Fuel burn: 24 gph for 65% cruise at 175 knots

 29 gph for 75% cruise at 180 knots 

Datum

Front seats +37 2nd seats +75 3rd seat or baggage limit 200 lb +113

Fuel +61

100 lb baggage −15 EWCG 35.28

32.0 43.6

CG limits

20.1 MAC = 61.6

Figure 6-2. Twin-engine airplane weight and balance diagram.

Item Weight (lb) (5,200 max.) Arm (in) Moment (lb-in) CG

120,093

51,240

11,840

23,250

−1,500

10,170

215,093 42.47

Airplane

Fuel (140 gal)

Front seat

Row 2 seats

Foward baggage

Aft baggage

Total

3,404

840

320

310

100

90

5,064

35.28

 61.0

 37.0

 75.0

 −15.0

 113.0

x =

Figure 6-3. Determining the loaded CG of the sample airplane in Figure 6-2.

The airplane in this example was weighed to determine its 

basic empty weight (BEW) and empty weight center of 

gravity (EWCG). The weighing conditions and results are:

Weight with fuel drained and oil full:

Right wheel scales........................1,084 lb, tare 8 lb

Left wheel scales..........................1,148 lb, tare 8 lb

Nose wheel scales.......................1,202 lb, tare 14 lb

Determine the Loaded CG

First, add the weights indicated by the individual scales and 

then subtract the tare weights to determine the BEW. Next, 

using the BEW and EWCG, the loaded weight and CG of the 

aircraft can be determined with data from Figure 6-2, using 

a chart such as the one in Figure 6-3.

The aircraft is loaded as shown:

Fuel (140 gal) ................................................ 840 lb

Front seats ..................................................... 320 lb

Row 2 seats.................................................... 310 lb

Forward baggage ........................................... 100 lb

Aft baggage ..................................................... 90 lb

Chart Method Using Weight, Arm, and Moments

Make a chart showing the weight, arm, and moments of the 

airplane and its load.
6-3

5,200

5,000

4,800

4,600

4,400

4,200

4,000

32 34 36 38 40 42 44

Inches from the datum

Weight (lb)

Figure 6-4. Sample CG range chart.

CG in % MAC CG in inches from LEMAC x 100

MAC =

22.37 x 100

61.6 =

= 36.3% MAC

Figure 6-5. Finding CG in percent MAC.

The loaded weight for this fl ght is 5,064 pounds, and the CG 

is located at 42.47 inches aft of the datum.

To determine that the weight and CG are within the allowable 

range, refer to the CG range chart in Figure 6-4. Draw a line 

vertically upward from 42.47 inches from the datum and one 

horizontally from 5,064 pounds. These lines cross inside the 

envelope, showing that the airplane is properly loaded.

Determining the CG in Percentage of Mean 

Aerodynamic Chord (MAC)

Refer again to Figures 6-2 and 6-3.

The loaded CG is 42.47 inches aft of the datum.

The MAC is 61.6 inches long.

The LEMAC is located at station 20.1.

The CG is 42.47 – 20.1 = 22.37 inches aft of LEMAC.

Use the formula in Figure 6-5 to findthe CG in percent MAC.

The loaded CG is located at 36.3 percent MAC.

The Chart Method Using Weight and Moment 

Indexes

As mentioned in the previous chapter, anything that can be 

done to make careful preflight planning easier makes flying

safer. Many manufacturers furnish charts in the Pilot’s 

Operating Handbook/Aircraft Flight Manual (POH/AFM) 

that use weight and moment indexes rather than weight, arm, 

and moments. The charts also help reduce errors by including 

tables of moment indexes for the various weights.

Consider the loading for this particular flight

Cruise fuel flow = 16 gallons per hou

Estimated time en route = 2 hours, 10 minutes

Reserve fuel = 45 minutes = 12 gallons

Total required fuel = 47 gallons

The pilot completes a chart like the one in Figure 6-6 using 

moment indexes from tables in Figures 6-7 and 6-8.

The moments divided by 100 in the index column are found 

in the charts in Figures 6-7 through 6-9. If the exact weight 

is not in the chart, interpolate between the weights that are 

included. When a weight is greater than any of those shown 

in the charts, add the moment indexes for a combination of 

weights to get that which is desired. For example, to get the 

moments divided by 100 for the 320 pounds in the front seats, 

add the moment index for 100 pounds (105) to that for 220 

pounds (231). This gives the moment index of 336 for 320 

pounds in the front seats. 

Use the moment limits versus weight envelope in Figure 6-10

to determine if the weight and balance conditions are 

within allowable limits for both takeoff and landing at the 

destination. The moment limits versus weight envelope is an 

enclosed area on a graph of three parameters. The diagonal 

line representing the moment divided by 100 crosses the 

horizontal line representing the weight at the vertical line 

representing the CG location in inches aft of the datum. When 

the lines cross inside the envelope, the aircraft is loaded 

within its weight and CG limits.

Takeoff: – 3,781 lb and 4,296 moment divided by 100

Landing: – 3,571 lb and 4,050 moment divided by 100
6-4

Weight and Balance Loading Form

Model _________________________________________________ Date ____________________________________________________ 

Serial Number___________________________________________ Reg. Number _____________________________________________ 

Item Pounds (3,900 max.) Index moment/100

Airplane basic empty weight

Front seat occupants

Row 2 seats

Baggage (200 lb max.)

Subtotal

 Zero fuel condition (3,500 max.)

 Fuel loading (gallons)

Subtotal

 Ramp condition

* Less fuel for start, taxi, and takeoff

Subtotal

Takeoff condition

 Less fuel to destination (gallons)

Landing condition

* Fuel for start, taxi, and takeoff is normally 24 pounds at a moment index of 28.

2,625

320

290

90

3,325

480

3,805

−24

3,781

−210

3,571

2,864

336

412

150

3,762

562

4,324

−28

4,296

−246

4,050

80

35

Figure 6-6. Typical weight and balance loading form.

Occupant Moments/100

Weight Front seats

Arm +105

Row 2 seats

Arm +142

100

110

120

130

140

150

160

170

180

190

200

210

220

230

240

250

105

116

126

137

147

158

168

179

189

200

210

221

231

242

252

263

142

156

170

185

199

213

227

241

256

270

284

298

312

327

341

355

Figure 6-7. Sample weight and moment index for occupants.

Locate the moment divided by 100 diagonal line for 4,296 and 

follow it down until it crosses the horizontal line for 3,781 

pounds. These lines cross inside the envelope at the vertical 

line for a CG location of 114 (113.6) inches aft of the datum.

The maximum allowable takeoff weight is 3,900 pounds, and 

this airplane weighs 3,781 pounds. The CG limits for 3,781 

pounds are 109.8 to 117.5. The CG of 114 (113.6) inches 

falls within these allowable limits.
6-5

Baggage Moments/100

Weight Arm +167

10

20

30

40

50

60

70

80

90

100

110

120

130

140

150

160

170

180

190

200

17

33

50

67

84

100

117

134

150

167

184

200

217

234

251

267

284

301

317

334

Figure 6-8. Sample weight and moment index for baggage.

Usable Fuel – Arm +117

Gallons Pounds Moment/100

10

20

30

40

50

60

70

80

90

100

60

120

180

240

300

360

420

480

540

600

70

140

211

281

351

421

491

562

632

702

Figure 6-9. Sample weight and moment index for fuel.
6-6

4,000

3,900

3,800

3,600

3,500

3,400

3,200

3,000

2,800

2,600

2,400

106 108 110 112 114 116 118

CG inches aft of datum

Weight (lb)

3,900

4,200

4,300

4,400

4,500

4,600

4,100

4,000

3,800

3,700

3,600

3,500

3,400

3,100

3,300

3,200

3,000

2,900

2,800

2,700

2,600

Maximum zero fuel weight

Moment/100

Envelope based on the following weight and CG limit data (landing gear down)

 Weight condition FWD CG limit Aft CG limit

3,900 pounds (max takeoff/landing) 110.6 117.5

3,250 pounds or less 106.6 117.5

Maximum takeoff

and landing weight

Figure 6-10. Sample moment limits versus weight envelope.
7-1

Introduction

The largest weight changes that occur during the lifetime of 

an aircraft are those caused by alterations and repairs. It is the 

responsibility of the FAA-certificated mechanic or repairman 

doing the work to accurately document the weight change 

and record it in both the maintenance records and the Pilot’s 

Operating Handbook/Aircraft Flight Manual (POH/AFM).

Center of Gravity Change After 

a Repair or Alteration

Chapter 7
7-2

Figure 7-1. A typical 14 CFR part 23 weight and balance record.

Weight and Balance Record

(Continuous history of changes in structure or equipment affecting weight and balance)

Airplane Model Serial Number Page Number

Date

Item No.

In Out

Description

of Article or

Modification

Weight Change Running

Basic Empty Added (+) Removed (−) Weight

Wt.

(lb)

Wt.

(lb)

Arm

(in)

Moment/

1,000

Wt.

(lb)

Arm

(in)

Moment/

1,000

Moment/

1,000

 As delivered 1,876 67.8

 8-7-09 Alteration per FAA Form 337 

 Dated 8-7-09 7.38 .346 1,883.4 68.1

 34-XX Turn coordinator –2.5 15.0 –.037 1,880.9 68.1

 34-XX Directional gyro –3.12 13.5 –.042 1,877.8 68.1

 22-XX Autopilot system 13.0 32.7 .425 1,890.8 68.5

Cessna 182L 18259080 1

Equipment List

The Federal Aviation Administration (FAA) considers 

the addition or removal of equipment included in the 

Comprehensive Equipment List to be a minor alteration. 

The weights and arms are included with the items in the 

equipment list, and these minor alterations can be done and the 

aircraft approved for return to service by an FAA-certificate 

mechanic or repairman. The only documentation required is an 

entry in the aircraft maintenance records and the appropriate 

change to the weight and balance record in the POH/AFM. 

[Figure 7-1]

Figure 7-2 is a comprehensive list of all Cessna equipment 

that is available for the Model 182S airplane. It should 

not be confused with the airplane-specific equipment list. 

An airplane-specific list is provided with each individual 

airplane at delivery and is typically found at the end of the 

Pilot’s Operating Handbook. The following comprehensive 

equipment list and the airplane-specific list have a similar 

order of listing.

The comprehensive equipment list provides the following 

information in column form:

In the Item No column, each item is assigned a coded number. 

The first two digits of the code represent the assignment 

of an equipment item within the ATA Specification 100 

breakdown (Chapter 11 for Placards, Chapter 21 for Air 

Conditioning, Chapter 77 for Engine Indicating, etc.). These 

assignments also correspond to the Maintenance Manual 

chapter breakdown for the airplane. Items receive a unique 

sequence number (01, 02, 03, etc.). After the sequence 

number (and hyphen), a suffix letter is assigned to identify 

an equipment item as required, standard, or optional. Suffix

letters are as follows:

–R = required item or equipment for FAA certificatio

–S = standard equipment item

–O = optional equipment item replacing required or 

standard item(s)

–A = optional equipment item that is in addition to 

required or standard items

In the Equipment List Description column, each item is 

assigned a descriptive name to help identify its function. 

In the Ref Drawing column, a drawing number is provided 

that corresponds to the item.

Note: Additional equipment must be installed in accordance 

with the reference drawing, service bulletin, or a separate 

FAA approval.

In the Wt and Arm columns, you find the weight in pounds 

and arm in inches of the equipment item.

Notes: Unless otherwise indicated, true values (not net 

change values) for the weight and arm are shown. Positive 

arms are distances aft of the airplane datum; negative 

arms are distances forward of the datum. Asterisks (*) in 

the weight and arm column indicate complete assembly 

installations. Some major components of the assembly are 

listed on the lines immediately following. The sum of these 

major components does not necessarily equal the complete 

assembly installation.
7-3

Figure 7-2. Typical comprehensive equipment list.

* Indicates total weight/arm for all subcomponents

Major Alteration and Repair

Within the following text, information concerning major 

repairs or major alterations does not apply to any aircraft 

within the light-sport category. This category of aircraft is 

not eligible for major repairs or alterations.

Any major alteration or repair requires the work to be done 

by an FAA-certificated mechanic or repairman or facility. 

The work must be checked for conformity to FAA-approved 

data and signed off by an aircraft mechanic holding an 

Inspection Authorization (IA) or by an authorized agent of 

an appropriately rated FAA-approved repair station. A Repair 

Station record and/or FAA form 337, Major Repair and 

Alteration, must be completed describing the work. (Reference 

Title 14 of the Code of Federal Regulations (14 CFR) part 43, 

appendix B, for the recording of major repair and alterations 

requirements.) A dated and signed revision to the weight and 

balance record is made and kept with the maintenance records, 

and the airplane’s new empty weight and empty weight arm 

or moment index are entered in the POH/AFM.

Weight and Balance Revision Record

Aircraft manufacturers use different formats for their weight 

and balance data, but Figure 7-3 is typical of a weight and 

balance revision record. All weight and balance records 

should be kept with the other aircraft records. Each revision 

record should be identified by the date, the aircraft make, 

model, and serial number. The pages should be signed by 

the person making the revision and his or her certificate type 

and number must be included.

The computations for a weight and balance revision 

are included on a weight and balance revision form. 

Appropriate fore-and-aft extreme loading conditions should 

be investigated and the computations shown. The date those 

computations were made is shown in the upper right corner 

of Figure 7-3. When this work is superseded, a notation 

must be made on the new weight and balance revision form, 

including a statement that the new computations supersede 

the computations dated “MM/DD/YY.”
7-4

Figure 7-3. A typical airplane weight and balance revision record.

Weight & Balance 

Cessna 182L

N42565

S/N 18259080 

Removed the following equipment:

1. Turn coordinator P/N C661003-0201

2. Directional gyro P/N 0706000

 

 

Installed the following equipment:

1. S-TEC System 40 autopilot, includes

 turn coordinator and directional gyro

*REVISED LICENSED EMPTY WEIGHT

 NEW USEFUL LOAD 1,216.62

 Forward check (Limit +33.0)

 Wt. × Arm = Moment

A/C empty 

Fwd. seats

Aft seats

Fuel (min.)

Baggage A

Baggage B

Jeseph P. Kline

A&P 123456789

Supersedes computations of FAA Form

337 dated 10/02/07.

 Weight x Arm = Moment

 2.50 lb 15.0 37.50

 +3.12 lb 13.5 +42.12

 5.62 79.62

 1,876.00 36.14 67,798.64

 −5.62 −79.62

 1,870.38 36.20 67,719.02

 

 Weight x Arm = Moment

 13 lb 32.70 425.13

 1,870.38 36.20 67,719.02

 +13.00 +425.13

 1,883.38 36.18 68,144.15

 Rearward check (Limit +46.0)

 

A/C empty 

Fwd. seats

Aft seats

Fuel (min.)

Baggage A

Baggage B

1,883.38 36.18 68,144.15

 170.00 37.00 6,290.00

 115.00 48.00 5,520.00

2,168.38 +36.87 79,954.15

Total

Aircraft after removal:

 Wt. × Arm = Moment

 1,883.38 36.18 68,144.15

 170.00 37.00 6,290.00

 340.00 74.00 25,160.00

 528.00 48.20 25,449.60

 100.00 97.00 9,700.00

 60.00 116.00 6,960.00

3,081.38 45.98 141,703.75

Date: 08/07/09
7-5

Figure 7-4. Weight, arm, and moment changes caused by typical alteration or repair.

Item Weight (lb) Arm (in) Moment (lb-in) New CG

67,723.6

−192.8

−874.0

−432.0

+122.6

+229.5

+2,037.0

68,613.9 +36.4

Airplane

Radio removed

Power supply removed

ELT removed

Radio installed

ELT installed

Passenger seat installed

Total

1,876.0

−12.2

−9.2

−3.2

+8.4

+1.7

+21.0

1,882.5

36.1

 15.8

95.0

135.0

14.6

135.0

97.0

x =

Figure 7-5. Weight and moment index changes caused by a typical alteration or repair.

Item Weight (lb) Moment indexes (lb-in/100) New CG (inches from datum)

+677.2

−1.93

−8.74

−4.32

+1.23

+2.29

+20.37

 +686.1 +36.4

Airplane

Radio removed

Power supply removed

ELT removed

Radio installed

ELT installed

Passenger seat installed

Total

1,876.0

−12.2

−9.2

−3.2

+8.4

+1.7

+21.0

1,882.5

The weight and balance revision sheet should clearly show 

the revised empty weight, empty weight arm and/or moment 

index, and the new useful load. An example of these entries 

can be found at the bottom of Figure 7-3.

Weight Changes Caused by a Repair or 

Alteration

A typical alteration might consist of removing two pieces 

of radio equipment from the instrument panel and a power 

supply that was located in the baggage compartment behind 

the rear seat. In this example, these two pieces are replaced 

with a single lightweight, self-contained radio. At the 

same time, an old emergency locator transmitter (ELT) is 

removed from its mount near the tail, and a lighter weight 

unit is installed. A passenger seat is installed in the baggage 

compartment.

Computations Using Weight, Arm, and Moment

The first step in the weight and balance computation is to 

make a chart like the one in Figure 7-4, listing all of the 

items that are involved. The new center of gravity (CG) of 

36.4 inches aft of the datum is determined by dividing the 

new moment by the new weight.

Computations Using Weight and Moment Indexes

If the weight and balance data uses moment indexes rather 

than arms and moments, this same alteration can be computed 

using a chart such as the one in Figure 7-5. Subtract the 

weight and moment indexes of all the removed equipment 

from the empty weight and moment index of the airplane. 

Add the weight and moment indexes of all equipment 

installed to determine the total weight and the total moment 

index. To determine the position of the new CG in inches 

aft of the datum, multiply the total moment index by 100 to 

get the moment, and divide this by the total weight to get 

the new CG.
7-6

Figure 7-8. Formula for determining the CG in percent MAC.

CG in % MAC CG in inches from LEMAC x 100

MAC =

10.42 x 100

58.0 =

= 17.9% MAC

Figure 7-7. Loading conditions.

CG

range

Datum Front seats Rear seats +74 Baggage B +116

Fuel +48.2 Baggage A +97

Forward

limit

Aft limit 46.0

Figure 7-6. Weight and balance information.

3,200

3,000

2,800

2,600

2,400

2,200

2,000

1,800 32 34 36 38 40 42 44 46 48

Fuselage station (in)

Loaded aircraft weight (lb)

Airplane EW and EWCG 1,876.0 lb at +36.14

Engine METO horsepower 230

CG range (+40.9) to (+46.0) at 3,100 lb

(+33.0) to (+46.0) at 2,250 lb or less

Straight line variation between points given

Empty weight CG range None

Maximum weight 3,100 lb takeoff/flight

2,950 lb landing

Datum to LEMAC 25.98

MAC 58.00

No. of seats 4 (2 front at +34.0)

(2 rear at +74.0)

Fuel capacity 92 gal (88 gal usable)

two 46-gal integral tanks in wings at +48.2

See NOTE 1 for data on unusable fuel.

Minimum fuel (METO HP ÷ 2) 115 lb at +48

Maximum baggage 160 lb

Area A (100 lb at +97.0)

Area B (60 lb at +116.0 

Oil capacity 12 qt (−15) (6 qt usable)

See NOTE 1 for data on undrainable oil.

NOTE 1: The certificated empty weight and corresponding center of

gravity location must include unusable fuel of 30 lb (+46) and 

undrainable oil of 0 lb.

Determining the CG in Percentage of Mean 

Aerodynamic Chord (Percent MAC)

This procedure is the same as found in Chapter 5, SingleEngine Aircraft Weight and Balance Computations. 

Refer to the load conditions and CG information found in 

Figures 7-5, 7-6, and 7-7 to compute the CG in percent MAC:

The loaded CG is +36.4 inches aft of the datum.

The MAC is 58.0 inches long.

The leading edge mean aerodynamic chord (LEMAC) 

is located at station 25.98.

The CG is +36.4 – 25.98 = 10.42 inches aft of LEMAC.

Use the formula in Figure 7-8 to determine CG in MAC 

percentages.

The loaded CG after alteration or repair is located at 17.9 

percent MAC.

Empty Weight CG (EWCG) Range

The fuel tanks, seats, and baggage compartments of some 

aircraft are so located that changes in the fuel or occupant 

load have a very limited effect on the balance of the aircraft. 

Aircraft of such a configuration show an empty weight CG 

(EWCG) range in the Type Certificate Data Sheet (TCDS). 

If the EWCG is located within this range, it is impossible to 

legally load the aircraft so that its loaded CG falls outside 

its allowable range.

If the TCDS lists an EWCG range, and after the alteration is 

completed the EWCG falls within this range, then there is no 

need to compute a fore and aft check for adverse loading. But 

if the TCDS lists the EWCG range as “None” (and most of 

them do), a check must be made to determine whether or not 

it is possible by any combination of legal loading to cause the 

aircraft CG to move outside of either its forward or aft limits.

Adverse-Load CG Checks

Many modern aircraft have multiple rows of seats and often 

more than one baggage compartment. After any repair or 

alteration that changes the weight and balance, the Airframe 

and Powerplant (A&P) FAA-certificated mechanic or 

repairman must ensure that no legal condition of loading can 
7-7

Figure 7-9. Load conditions for forward adverse-load CG check.

Item Weight (lb) Arm (in) Moment (lb-in) Most forward CG +33.0

67,798.6

5,780.0

5,520.0

79,098.6 +36.6

Airplane (empty)

Pilot

Fuel (minimum)

Total

1,876.0

170.0

115.0

2,161.0

36.14

 34.0

48.0

x =

When rear row of seats is occupied, 120 pounds of 

baggage or ballast must be carried in forward baggage 

compartment. For additional loading instruction, 

see Weight and Balance Data.

Figure 7-11. Typical baggage compartment placard.

Figure 7-10. Load conditions for aft adverse-load CG check.

Item Weight (lb) Arm (in) Moment (lb-in) Most Aft CG +46.0

67,798.6

5,780.0

25,449.6

25,160.0

9,700.0

6,960.0

140,848.2 +45.8

Airplane (empty)

Pilot

Fuel (full tanks – 88 gal)

Rear seat occupants (2)

Baggage A

Baggage B

Total

1,876.0

170.0

528.0

340.0

100.0

60.0

3,074.0

 36.14

34.0

 48.2

 74.0

 97.0

 116.0

x =

move the CG outside of its allowable limits. To determine 

this, adverse-loaded CG checks must be performed and 

the results noted in the weight and balance revision sheet. 

[Figure 7-3]

Forward Adverse-Load CG Check

To conduct a forward CG check, make a chart that includes 

the airplane and any occupants and items of the load located 

in front of the forward CG limit. Include only those items 

behind the forward limit that are essential to flight: the pilot, 

and the minimum fuel.

In this example, the pilot, whose nominal weight is 170 

pounds, is behind the forward CG limit. The fuel is also 

behind the forward limit, so the minimum fuel is used. For 

weight and balance purposes, the minimum fuel is no more 

than the quantity needed for one-half hour of operation at 

rated maximum continuous power. This is considered to 

be 1⁄12 gallon for each maximum except takeoff (METO) 

horsepower. Because aviation gasoline weighs 6 pounds per 

gallon, determine the number of pounds of the minimum fuel 

by dividing the METO horsepower by two. In this example, 

minimum fuel is 115 pounds. The front and rear seats and the 

baggage are all behind the forward CG limit, so no passengers 

or baggage are considered.

Make a chart like the one in Figure 7-9 to determine the CG 

with the aircraft loaded for its most forward CG. With the 

load consisting of only a pilot and the minimum fuel, the CG 

is +36.6, which is behind the most forward allowable limit 

for this weight of +33.0.

Aft Adverse-Load CG Check

To conduct an aft or rearward CG check, make a chart that 

includes the empty weight and EWCG of the aircraft after 

the alteration and all occupants and items of the load behind 

the aft CG limit of 46.0. The pilot is in front of this limit but 

is essential for flight and must be included. In this example, 

only the pilot occupies the front seats. Since the CG of the 

fuel is behind the aft limit, full fuel is used, as well as the 

nominal weight (170 lb) for both rear seat passengers and 

the maximum allowable baggage.

Under these loading conditions, the CG is located at +45.8, 

which is ahead of the aft limit of +46.0. [Figure 7-10] With 

only the pilot in front of the aft CG limit and maximum of all 

items behind the aft limit, the CG is at +45.8 inches, which 

is ahead of the aft limit of +46.0 inches.

Ballast

It is possible to load most modern airplanes so the CG 

shifts outside of the allowable limit. Placards and loading 

instructions in the weight and balance data inform the pilot 

of the restrictions that prevent such a shift from occurring. 

A typical placard in the baggage compartment of an airplane 

is shown in Figure 7-11. When the CG of an aircraft falls 

outside of the limits, it can usually be brought back in by 

using ballast.
7-8

Figure 7-13. Formula for determining ballast.

Ballast weight Aircraft empty weight × Dist. out of limits

Distance between ballast and desired CG =

Temporary Ballast

Temporary ballast, in the form of lead bars or heavy canvas 

bags of sand or lead shot, is often carried in the baggage 

compartments to adjust the balance for certain flight 

conditions. The bags are marked “Ballast XX Pounds—

Removal Requires Weight and Balance Check.” Temporary 

ballast must be secured so it cannot shift its location in flight 

and the structural limits of the baggage compartment must not 

be exceeded. All temporary ballast must be removed before 

the aircraft is weighed.

Temporary Ballast Formula

The CG of a loaded airplane can be moved into its allowable 

range by shifting passengers or cargo or by adding temporary 

ballast.

Permanent Ballast

If a repair or alteration causes the aircraft CG to fall outside 

of its limit, permanent ballast can be installed. Usually 

permanent ballast is made of blocks of lead painted red and 

marked “Permanent Ballast—Do Not Remove.” It should be 

attached to the structure so that it does not interfere with any 

control action and attached rigidly enough that it cannot be 

dislodged by any flight maneuvers or rough landing

Two things must first be known to determine the amount of 

ballast needed to bring the CG within limits: the amount the 

CG is out of limits, and the distance between the location of 

the ballast and the limit that is affected.

If an airplane with an empty weight of 1,876 pounds has been 

altered so its EWCG is +32.2, and CG range for weights up 

to 2,250 pounds is +33.0 to +46.0, permanent ballast must be 

installed to move the EWCG from +32.2 to +33.0. There is 

a bulkhead at fuselage station 228 strong enough to support 

the ballast.

To determine the amount of ballast needed, use the formula 

in Figure 7-13.

 Ballast needed = 1.876 × 0.8

228 – 33

= 1500.8

 195

= 7.7 pounds

A block of lead weighing 7.7 pounds, attached to the bulkhead 

at fuselage station 228, moves the EWCG back to its proper 

forward limit of +33. This block should be painted red and 

marked “Permanent Ballast—Do Not Remove.”
8-1

Introduction

Weight and balance considerations of a helicopter are similar 

to those of an airplane, except they are far more critical, 

and the center of gravity (CG) range is much more limited. 

[Figures 8-1 and 8-2] The engineers who design a helicopter 

determine the amount of cyclic control authority that is 

available, and establish both the longitudinal and lateral CG 

envelopes that allow the pilot to load the helicopter so there 

is sufficient cyclic control for all flight conditions

Weight and Balance Control—

Helicopter

Chapter 8
8-2

Figure 8-1. Weight and balance data needed to determine proper 

loading of a helicopter.

Helicopter basic empty weight 1,545.0 lb, EWCG +101.4

(+0.2 lateral offset)

CG range (+92.0) to (+98.0) at 1,600 lb

(+92.0) to (+95.0) at 2,250 lb 

Empty weight CG range None

Maximum weight 2,250 lb 

Number of seats 2 at (+64.0) (L−13.5) (R+13.5)

lateral offset

Maximum baggage (100 lb) at +105.0 (0.0 lateral offset)

Fuel capacity 50 gal (48 gal usable); 

288 lb at (+96.0) (−8.4 lateral offset)

Oil capacity 4.3 qt at (+90) (0.0 lateral)

Figure 8-2. Typical helicopter datum, flight stations, and butt line 

locations.

Datum FS 0.0 FS 68.0 FS 254.5

BL 27.4 R

BL −38.7 L

BL 0

Butt 

line 

zero

Figure 8-3. Typical helicopter CG envelopes.

2,400

2,200

2,000

1,800

1,600 92 94 96 98 100

Longitudinal CG (inches aft of datum)

Gross Weight (lb)

Longitudinal CG Envelope

Approved Area

2,203

2,027

1,877

94.4 95.0 98.2

2,300

2,200

2,100

2,000

1,900

1,800

1,700

1,600

−5,000 −3,000 −1,000 0 1,000 3,000 5,000 

Left Lateral Offset Moment Right

(lb-in)

Gross Weight (lb)

Lateral Offset Moment Envelope

2,203

2,027

1,877

−1,705 −326 1,699

If the CG is ahead of the forward limit, the helicopter tilts 

and the rotor disk has a forward pull. To counteract this 

and maintain a stationary position, rearward cyclic stick 

displacement would be required. If the CG is too far forward, 

there may not be enough available cyclic authority to allow 

the helicopter to flare during landing, and it consequently 

requires an excessive landing distance.

If the CG is aft of the allowable limits, the helicopter flies

with a tail-low attitude and may need more forward cyclic 

stick displacement than is available to maintain a hover in a 

no-wind condition. There might not be enough cyclic travel 

to prevent the tail boom from striking the ground. If gusty 

winds should cause the helicopter to pitch up during high 

speed fl ght, there might not be enough forward cyclic control 

to safely lower the nose.

Helicopters are approved for a specifi maximum gross 

weight, but it is not safe to operate them at this weight under 

some conditions. A high density altitude decreases the safe 

maximum weight as it affects the hovering, takeoff, climb, 

autorotation, and landing performance.

The fuel tanks on some helicopters are behind the CG, 

causing it to shift forward as fuel is used. Under some flight

conditions, the balance may shift enough that there is not 

sufficient cyclic authority to flare for landing. For these 

helicopters, the loaded CG should be computed for both 

takeoff and landing weights.

Lateral balance of an airplane is usually of little concern 

and is not normally calculated. Some helicopters, especially 

those equipped for hoist operations, are sensitive to the lateral 

position of the CG and their Pilot’s Operating Handbook/

Rotorcraft Flight Manual (POH/RFM) include both 

longitudinal and lateral CG envelopes, as well as information 

on the maximum permissible hoist load. Figure 8-3 is an 

example of such CG envelopes.
8-3

Determining the Loaded CG of a 

Helicopter

The empty weight and empty weight center of gravity 

(EWCG) of a helicopter are determined in the same way 

as for an airplane. See Chapter 5, Single-Engine Aircraft 

Weight and Balance Computations. The weights recorded 

on the scales supporting the helicopter are added and their 

distances from the datum are used to compute the moments 

at each weighing point. The total moment is divided by the 

total weight to determine the location of the CG in inches 

from the datum. The datum of some helicopters is located at 

the center of the rotor mast, but since this causes some arms 

to be positive (behind the datum) and others negative (ahead 

of the datum), most modern helicopters have the datum 

located ahead of the aircraft, as do most modern airplanes. 

When the datum is ahead of the aircraft, all longitudinal 

arms are positive.

The lateral CG is determined in the same way as the 

longitudinal CG, except the distances between the scales and 

butt line zero (BL 0) are used as the arms. Arms to the right 

of BL 0 are positive and those to the left are negative. The 

butt line zero (or sometimes referred to as the buttock) is a 

line through the symmetrical center of an aircraft from nose 

to tail. It serves as the datum for measuring the arms used to 

find the lateral CG. Lateral moments that cause the aircraft 

to roll clockwise are positive (+), and those that cause it to 

roll counterclockwise are negative (–).

To determine whether or not a helicopter is within both 

longitudinal and lateral weight and balance limits, construct 

a table like the one in Figure 8-4, with the following data 

specific to the aircraft

Empty weight 1,545 lb

EWCG 101.4 inches aft of 

the datum

Lateral balance arm 0.2 inches right 

of BL 0

Maximum allowable gross weight 2,250 lb

Pilot 200 lb @ 64 inches 

aft of datum and 

13.5 inches right of 

BL 0

Passenger 170 lb @ 64 inches 

aft of datum and 

–13.5 in left of BL 0

Fuel (48 gal) 288 lb @ 96 inches 

aft of datum and 

–8.4 inches left of 

BL 0

Check the helicopter CG envelopes in Figure 8-3 to determine 

whether or not the CG is within limits both longitudinally 

and laterally.

In the longitudinal CG envelope, draw a line vertically 

upward from the CG of 94.4 inches aft of datum and a 

horizontal line from the weight of 2,203 pounds gross weight. 

These lines cross within the approved area.

In the lateral offset moment envelope, draw a line vertically 

upward from the –1,705 lb-in point (on the left side of the 

horizontal axis) and a line horizontally from 2,203 pounds on 

the gross weight index. These lines cross within the envelope, 

showing the lateral balance is also within limits.

Effects of Offloading Passengers an 

Using Fuel

Consider the helicopter in Figure 8-4. The first leg of the 

flight consumes 26 gallons of fuel, and at the end of this leg, 

the passenger deplanes. Is the helicopter still within allowable 

CG limits for takeoff? To find out, make a new chart like the 

one in Figure 8-5 to show the new loading conditions of the 

helicopter at the beginning of the second leg of the flight

Under these conditions, according to the helicopter CG 

envelopes in Figure 8-3, both the longitudinal CG and the 

lateral offset moment fall outside of the approved area of 

the envelope. The aircraft longitudinal CG is too far aft and 

the potential for excessive tail-low attitudes is very high. 

Under these conditions, it is possible that there will not be 

enough forward cyclic authority to maintain level flight 

The helicopter’s lateral offset moment is too far right and 

may lead to control issues, as well as an increased hazard of 

dynamic rollover. One possible option to bring the aircraft 

loading conditions within the approved envelope is to load 

either ballast or a passenger, as computed in Figure 8-6 and 

plotted in Figure 8-3.
8-4

Figure 8-6. Determining the longitudinal CG and the lateral offset moment for the second leg of the flight with ballast and/or a different 

passenger.

Item Weight Weight Longitude

Arm

Latitude

Arm

Longitude

Moment

Longitude

CG

Lateral Offset

Moment

95.0

Helicopter empty weight

Pilot

Ballast/Passenger

Fuel (22 gallons)

156,663

12,800

9,600

14,976

194,039

1,545

200

150

132

2,027

101.4

 64.0

64.0

96.0

309

+2,700

−2,025

−1,310

−731

1,545

200

150

132

2,027

+0.2

 +13.5

−13.5

−8.4

x = x =

Figure 8-5. Determining the longitudinal CG and the lateral offset moment for the second leg of the flight.

Item Weight Weight Longitude

Arm

Latitude

Arm

Longitude

Moment

Longitude

CG

Lateral Offset

Moment

98.2

Helicopter empty weight

Pilot

Fuel (22 gallons)

156,663

12,800

14,976

184,439

1,545

200

132

1,877

101.4

 64.0

96.0

309

+2,700

−1,310

+1,699

1,545

200

132

1,877

+0.2

 +13.5

−8.4

x = x =

Figure 8-4. Determining the longitudinal CG and the lateral offset moment.

Item Weight Weight Longitude

Arm

Latitude

Arm

Longitude

Moment

Longitude

CG

Lateral Offset

Moment

94.4

Helicopter empty weight

Pilot

Passenger

Fuel (48 gallons)

156,663

12,800

10,880

27,648

207,991

1,545

200

170

288

2,203

101.4

 64.0

64.0

96.0

309

+2,700

−2,295

−2,419

−1,705

1,545

200

170

288

2,203

+0.2

 +13.5

−13.5

−8.4

x = x =
9-1

Introduction

This chapter discusses general guidelines and procedures for 

weighing large fixed-wingaircraft exceeding a takeoff weight 

of 12,500 pounds. Several examples of center of gravity (CG) 

determination for various operational aspects of these aircraft 

are also included. Persons seeking approval for a weight and 

balance control program for aircraft operated under Title 

14 of the Code of Federal Regulations (14 CFR) part 91, 

subpart K, 121, 125, or 135 should consult with the Flight 

Standards District Office (FSDO) or Certificate Management 

Office (CMO) that has jurisdiction in their area. Additional 

information on weight and balance for large aircraft can be 

found in Federal Aviation Administration (FAA) Advisory 

Circular (AC) 120-27, Aircraft Weight and Balance Control, 

FAA Type Certificate Data Sheets (TCDS), and the aircraft 

flight and maintenance manuals for specific aircraf

Weight and Balance Control—

Commuter Category and 

Large Aircraft

Chapter 9
9-2

Figure 9-1. Determining the distance of CG.

Distance CG to LEMAC Datum to CG – Datum to LEMAC

CG

Distance weight is shifted

=

Figure 9-2. Determining the EWCG in percent MAC.

EWCG in % MAC CG in inches from LEMAC × 100

MAC =

Establishing the Initial Weight of an 

Aircraft 

Prior to being placed into service, each aircraft is weighed 

and the empty weight and CG location established. New 

aircraft are normally weighed at the factory and are eligible 

to be placed into operation without reweighing if the weight 

and balance records were adjusted for alterations and 

modificationsto the aircraft, such as interior reconfigurations

An aircraft transferred from one operator that has an approved 

weight and balance program to another operator with an 

approved program does not need to be weighed prior to use by 

the receiving operator unless more than 36 calendar months 

have elapsed since the last individual or fleet weighing, or 

unless some other modification to the aircraft warrants that 

the aircraft be weighed. Aircraft transferred, purchased, or 

leased from an operator without an approved weight and 

balance program, and that have not been modified or have 

been minimally modified, can be placed into service without 

being reweighed if the last weighing was accomplished by an 

acceptable method (for example, manufacturer’s instructions 

or AC 43.13-2, Acceptable Methods, Techniques, and 

Practices—Aircraft Alterations) within the last 12 calendar 

months and a weight and balance change record was 

maintained by the operator. It is potentially unsafe to fail to 

reweigh an aircraft after it has been modified

When weighing large aircraft, compliance with the relevant 

manuals, operations specifications, or management 

specification is required to ensure that weight and balance 

requirements specified in the Aircraft Flight Manual (AFM) 

are met in accordance with approved limits. This provides 

information to the flight crew that allows the maximum 

payload to be carried safely. 

The aircraft should be weighed in still air or an enclosed 

building after the aircraft has been cleaned. Ensure that the 

aircraft is in a configuration for weighing with regard to 

flight controls, unusable fuel, ballast, oil and other operating 

fluids, and equipment as required by the controlling weight 

and balance procedure.

Large aircraft are not usually raised off the floor on jacks 

for weighing; they are weighed on ramp-type scales. The 

scales must be properly calibrated, zeroed, and used in 

accordance with the manufacturer’s instructions. Each scale 

should be periodically checked for accuracy as recommended 

in the manufacturer’s calibration schedule, either by the 

manufacturer or by a recognized facility, such as a civil 

department of weights and measures. If no manufacturer’s 

schedule is available, the period between calibrations should 

not exceed 12 months.

Determining the Empty Weight and 

Empty Weight CG (EWCG)

When the aircraft is properly prepared for weighing, roll it 

onto the scales, and level it. The weights are measured at 

three weighing points: the two main wheel points and the 

nosewheel point. The empty weight and empty weight CG 

(EWCG) are determined by using the following steps with 

the results recorded in the weight and balance record for use 

in all future weight and balance computations.

1. Determine the moment index of each of the mainwheel points by multiplying the net weight (scale 

reading minus tare weight), in pounds, at these points 

by the distance from the datum, in inches. Divide these 

numbers by the appropriate reduction factor.

2. Determine the moment index of the nosewheel 

weighing point by multiplying its net weight, in 

pounds, by its distance from the datum, in inches. 

Divide this by the reduction factor.

3. Determine the total weight by adding the net weight of 

the three weighing points and the total moment index 

by adding the moment indexes of each point.

4. Divide the total moment index by the total weight and 

multiply the result by the reduction factor. This gives 

the CG in inches from the datum.

5. Determine the distance of the CG behind the leading 

edge of the mean aerodynamic chord (LEMAC) 

by subtracting the distance between the datum and 

LEMAC from the distance between the datum and 

the CG. [Figure 9-1]

6. Determine the EWCG in percentage of MAC (percent 

MAC) by using the formula in Figure 9-2.
9-3

Figure 9-3. Incremental weight changes that should be recorded in a weight and balance change record.

In the weight change record of a—

Large cabin aircraft

Medium cabin aircraft

Small cabin aircraft

An operator should record any weight changes of—

± 10 lb or greater

± 5 lb or greater

± 1 lb or greater

Figure 9-4. Loading schedule.

Item Weight (lb) Moment/1,000

92,837

1,781

16,602

1,020

2,915

10,451

10,451

25,589

161,646

BOW

Passengers Fwd station

Passengers Aft station

Fwd cargo

Aft cargo

Fuel tank 1

Fuel tank 3

Fuel tank 2

105,500

3,060

16,150

1,500

2,500

10,500

10,500

28,000

177,710

Documenting Changes to an Aircraft’s 

Weight and Balance 

The weight and balance system should include methods by 

which a complete, current, and continuous record of the 

weight and CG of each aircraft is maintained, such as a log, 

ledger, or other equivalent electronic means. Alterations and 

changes affecting the weight and/or balance of the aircraft 

should be recorded in this log. Changes in the weight or 

location of weight in or on the aircraft should be recorded 

whenever the weight change is at or exceeds the weights 

listed in Figure 9-3.

Determining the Loaded CG of the 

Airplane in Percent MAC

A loading schedule is used to document compliance with the 

certificated weight and balance limitations contained in the 

manufacturer’s AFM and weight and balance manual. The 

basic operating weight (BOW) and the operating index are 

entered into a loading schedule like the one in Figure 9-4, and 

the variables for a specific flight are entered as appropriate 

to determine the loaded weight and CG.

Use the data in this example:

Basic operating weight ..................................105,500 lb

Basic operating index (total moment/1,000) .... 98,837.0

MAC..................................................................180.9 in

LEMAC ................................................................. 860.5

Figure 9-5 illustrates passenger, cargo, and fuel loading 

tables. Using these tables, determine the moment indexes 

for the passengers (PAX), cargo, and fuel.

The airplane is loaded in this way:

Passengers (nominal weight—170 pounds each)

Forward compartment................................................. 18

Aft compartment......................................................... 95

Cargo

Forward hold......................................................1,500 lb

Aft hold..............................................................2,500 lb

Fuel

Tanks 1 and 3...........................................10,500 lb each

Tank 2..............................................................28,000 lb

The formula in Figure 9-6 can be used to determine the 

location of the CG in inches aft of the datum.
9-4

Figure 9-5. Loading schedule for determining weight and CG.

 # of passengers Weight (lb) Moment/1,000

 Forward compartment centroid—582.0

 Aft compartment centroid—1,028.0

 5 850 495

 10 1,700 989

 15 2,550 1,484

 20 3,400 1,979

 25 4,250 2,473

 29 4,930 2,869

 10 1,700 1,748

 20 3,400 3,495

 30 5,100 5,243

 40 6,800 6,990

 50 8,500 8,738

 60 10,200 10,486

 70 11,900 12,233

 80 13,600 13,980

 90 15,300 15,728

 100 17,000 17,476

 110 18,700 19,223

 120 20,400 20,971

 133 22,610 23,243

 

Weight (lb) Forward hold arm—680.0 Aft hold arm—1,166.0

 6,000 6,966

 5,000 3,400 5,830

 4,000 2,720 4,664

 3,000 2,040 3,498

 2,000 1,360 2,332

 1,000 680 1,166

 900 612 1,049

 800 544 933

 700 476 816

 600 408 700

 500 340 583

 400 272 466

 300 204 350

 200 136 233

 100 68 117 

Fuel Loading Table

Weight (lb) Arm Moment/1,000

 8,500 992.1 8,433

 9,000 993.0 8,937

 9,500 993.9 9,442

 10,000 994.7 9,947

 10,500 995.4 10,451

 11,000 996.1 10,957

 11,500 996.8 11,463

 12,000 997.5 11,970

Weight (lb) Arm Moment/1,000

 8,500 917.5 7,799

 9,000 917.2 8,255

 9,500 917.0 8,711

 10,000 916.8 9,168

 10,500 916.6 9,624

 11,000 916.5 10,082

 11,500 916.3 10.537

 12,000 916.1 10,993

Weight (lb) Arm Moment/1,000

 22,500 914.5 20,576

 23,000 914.5 21,034

 23,500 914.4 21,488

 24.000 914.3 21,943

 24,500 914.3 22,400

 25,000 914.2 22,855

 25,500 914.2 23,312

 26,000 914.1 23,767

 26,500 914.1 24,244

 27,000 914.0 24,678

 27,500 913.9 25,132

 28,000 913.9 25,589

 28,500 913.8 26,043

 29,000 913.7 26,497

 29,500 913.7 26,954

 30,000 913.6 27,408

 Full capacity **(see note at lower left)

 Tanks 1 and 3 (each) Tank 2 (3 cells)

Passenger Loading Table Cargo Loading Table

Moment/1,000

 18,500 915.1 16,929

 19,000 915.0 17,385

 19,500 914.9 17,841

 20,000 914.9 18,298

 20,500 914.8 18,753

 21,000 914.7 19,209

 21,500 914.6 19,664

 22,000 914.6 20,121

** Note:

 Computations for Tank 2 weights for

 12,500 lb to 18,000 lb have been

 purposely omitted.

 Full capacity
9-5

Figure 9-6. Determining the location of the CG in inches aft of 

the datum.

CG inches aft of datum = × 1,000 ( ) Total moment index

Total weight

= × 1,000 ( ) 161,646

177,710

= 909.6 inches

Figure 9-7. Determining the distance from the CG to the LEMAC.

Distance CG to LEMAC = Datum to CG – Datum to LEMAC

= 49.1 inches

= 909.6 – 860.5

Figure 9-8. Determining the location of the CG in percent MAC.

CG % MAC = × 100 ( ) Distance CG to LEMAC

MAC

= × 100 ( ) 49.1

180.9

= 27.1%

Determine the distance from the CG to the LEMAC by 

subtracting the distance between the datum and LEMAC from 

the distance between the datum and the CG. [Figure 9-7]

The location of the CG in percent MAC must be known in 

order to set the stabilizer trim takeoff. [Figure 9-8]

Operational Empty Weight (OEW)

Operational empty weight (OEW) is the basic empty weight 

or fl et empty weight plus operational items. The operator has 

two choices for maintaining OEW. The loading schedule may 

be utilized to compute the operational weight and balance of 

an individual aircraft, or the operator may choose to establish 

fleet empty weights for a fleet or group of aircraf

Reestablishing the OEW

The OEW and CG position of each aircraft should be 

reestablished at the reweighing. In addition, it should be 

reestablished through calculation whenever the cumulative 

change to the weight and balance log is more than plus or 

minus one-half of 1 percent (0.5 percent) of the maximum 

landing weight, or whenever the cumulative change in the 

CG position exceeds one-half of 1 percent (0.5 percent) of the 

MAC. In the case of rotorcraft and aircraft that do not have a 

MAC-based CG envelope (e.g., canard equipped airplane), 

whenever the cumulative change in the CG position exceeds 

one-half of 1 percent (0.5 percent) of the total CG range, the 

weight and balance should be reestablished. 

When reestablishing the aircraft OEW between reweighing 

periods, the weight changes may be computed provided the 

weight and CG location of the modifications are known; 

otherwise, the aircraft must be reweighed. 

Fleet Operating Empty Weights (FOEW) 

An operator may choose to use one weight for a fleet or group 

of aircraft if the weight and CG of each aircraft is within the 

limits stated above for establishment of OEW. When the 

cumulative changes to an aircraft weight and balance log 

exceed the weight or CG limits for the established fleetweight, 

the empty weight for that aircraft should be reestablished. 

This may be done by moving the aircraft to another group, or 

reestablishing new fleet operating empty weights (FOEWs)

Onboard Aircraft Weighing System

Some large transport airplanes have an onboard aircraft 

weighing system (OBAWS) that, when the aircraft is on the 

ground, gives the flight crew a continuous indication of the 

aircraft total weight and the location of the CG in percent 

MAC. Procedures are required to ensure the onboard weight 

and balance system equipment is periodically calibrated in 

accordance with the manufacturer’s instructions. 

An operator may use an onboard weight and balance 

system to measure an aircraft’s weight and balance as a 

primary means to dispatch an aircraft, provided the FAA 

has certified the system and approved the system for use in 

an operator’s weight and balance control program. As part 

of the approval process, the onboard weight and balance 

system must maintain its certificated accuracy. The accuracy 

demonstration test is provided in the maintenance manual 

portion of the Supplemental Type Certificate (STC) or type 

certificate of the onboard weight and balance system.

The system consists of strain-sensing transducers in each 

main wheel and nosewheel axle, a weight and balance 

computer, and indicators that show the gross weight, the 

CG location in percent MAC, and an indicator of the ground 

attitude of the aircraft.

The strain sensors measure the amount each axle defl cts and 

sends this data into the computer, where signals from all of 

the transducers and the ground attitude sensor are integrated. 

The results are displayed on the indicators for the flight crew. 

Using an onboard weight and balance system does not relieve 

an operator from the requirement to complete and maintain 

a load manifest. 
9-6

Figure 9-9. Stabilizer trim setting in ANU units.

Stabilizer Trim Setting—Units Airplane Nose Up

6

8

10

12

14

16

18

20

22

24

26

28

30

32

8

73/4

71/2

7

63/4

61/4

53/4

51/2

5

41/2

4

31/2

3

21/2

CG Flaps (all)

Figure 9-10. Determining the distance from CG to the LEMAC.

Distance CG to LEMAC = Datum to CG – Datum to LEMAC

= 10.7 inches

= 635.7 – 625

Figure 9-11. Determining the location of CG in percent MAC.

CG in % MAC = × 100 ( ) Distance CG to LEMAC

MAC

= × 100 ( ) 10.7

134.0

= 8.0 % MAC

Figure 9-12. Determining the location of CG in inches before cargo 

is removed.

CG (inches aft of LEMAC) = × MAC ( ) CG in % MAC

100

= × 141.5 ( ) 22.5

100

= 31.84 inches

Determining the Correct Stabilizer Trim Setting

It is important before takeoff to set the stabilizer trim for the 

existing CG location. There are two ways the stabilizer trim 

setting systems may be calibrated: in percent MAC and in 

units airplane nose up (ANU). 

If the stabilizer trim is calibrated in percent MAC, determine 

the CG location in percent MAC as has just been described, 

then set the stabilizer trim on the percentage figure thus 

determined. Some aircraft give the stabilizer trim setting in 

units of ANU that correspond with the location of the CG 

in percent MAC. When preparing for takeoff in an aircraft 

equipped with this system, first determine the CG in percent 

MAC in the way described above, then refer to the stabilizer 

trim setting chart on the takeoff performance page of the 

pertinent AFM. Figure 9-9 is an excerpt from the AFM chart 

on the takeoff performance of a Boeing 737.

Consider an airplane with these specifications

CG location................................................ station 635.7

LEMAC ........................................................ station 625

MAC..................................................................134.0 in

1. Determine the distance from the CG to the LEMAC 

by using the formula in Figure 9-10.

2. Determine the location of the CG in percent MAC by 

using the formula found in Figure 9-11.

Refer to Figure 9-9 for all flap settings and a CG located 

at 8 percent MAC; the stabilizer setting is 73

⁄4 units ANU.

Determining CG Changes Caused by 

Modifying the Cargo

Since large aircraft can carry substantial cargo, adding, 

subtracting, or moving any of the cargo from one hold to 

another can cause large shifts in the CG. 

Effects of Loading or Offloading Cargo

Both the weight and CG of an aircraft are changed when 

cargo is loaded or offloaded. In the following example, the 

new weight and CG are calculated after 2,500 pounds of 

cargo is offloaded from the forward cargo hold

Aircraft specifications are

Loaded weight .................................................90,000 lb

Loaded CG........................................ 22.5 percent MAC

Weight change ...................................................2,500 lb

Forward cargo hold centroid...................... station 352.1

MAC..................................................................141.5 in

LEMAC ................................................... station 549.13

1. Determine the CG location in inches from the 

datum before the cargo is removed. Do this by first

determining the distance of the CG aft of the LEMAC. 

[Figure 9-12]
9-7

Figure 9-13. Determining the distance between CG and the datum.

CG (inches from datum) = CG inches aft of LEMAC

 + Datum to LEMAC

= 580.97 inches

= 31.84 + 549.13

Figure 9-14. Determining the moment/1,000 for the original weight.

Moment/1,000 = Weight × Arm

1,000

90,000 × 580.97

1,000

= 52,287.3

=

Figure 9-15. Determining the moment/1,000 of the removed weight.

Moment/1,000 = Weight × Arm

1,000

2,500 × 352.1

1,000

= 880.25

=

Figure 9-16. New weights and CG.

Weight (lb) Moment/1,000 CG (inches from datum) CG (percent MAC)

580.97

587.51

22.5

27.12

Original data 

Changes

New data

90,000

– 2,500

87,500

52,287.3

– 880.3

51,407.0

Figure 9-17. Determining the location of new CG.

CG = × 1,000 Total moment/1,000

Total weight

51,407.0

87,500 = × 1,000

= 587.51 inches behind the datum

Figure 9-18. Determining the distance between the CG and LEMAC.

CG (inches aft of LEMAC) =

CG (inches from datum) – LEMAC

= 38.38 inches

= 587.51 – 549.13

Figure 9-19. Determining the new CG in percent MAC.

CG % MAC = × 100 ( ) Distance CG to LEMA

MAC

= × 100 ( ) 38.38

141.5

= 27.12% MAC

2. Determine the distance between the CG and the datum 

by adding the CG in inches aft of LEMAC to the 

distance from the datum to LEMAC. [Figure 9-13]

3. Determine the moment/1,000 for the original weight. 

[Figure 9-14]

4. Determine the new weight and new CG by first 

determining the moment/1,000 of the removed weight. 

Multiply the weight removed (2,500 pounds) by the 

centroid of the forward cargo hold (352.1 inches), and 

then divide the result by 1,000. [Figure 9-15]

5. Subtract the removed weight from the original weight 

and subtract the moment/1,000 of the removed weight 

from the original moment/1,000. [Figure 9-16]

6. Determine the location of the new CG by dividing 

the total moment/1,000 by the total weight and 

multiplying this by the reduction factor of 1,000. 

[Figure 9-17]

7. Convert the new CG location to percent MAC. First, 

determine the distance between the CG location and 

LEMAC. [Figure 9-18]

8. Then, determine the new CG in percent MAC. 

[Figure 9-19]

Loading 3,000 pounds of cargo into the forward cargo hold 

moves the CG forward 5.51 inches, from 27.12 percent MAC 

to 21.59 percent MAC.
9-8

Figure 9-24. Determining the change in CG caused by shifting 

2,500 pounds of cargo.

CG = × 100 Weight shifted × Distance shifted

Total weight

= 2,500 × (227.9 + 144.9)

90,000

= 2,500 × 372.8

90,000

= 10.36 inches

Figure 9-21. Determining the new CG after shifting cargo weight.

New CG = Old CG ± CG

= 591.33 inches

= 580.97 + 10.36

Figure 9-22. Converting the location of CG to percent MAC.

CG % MAC = × 100 ( ) CG inches

MAC

= × 100 ( ) 10.36

141.5

= 7.32% MAC

Figure 9-23. Determining the new CG in percent MAC.

New CG % MAC = Old CG ± CG

= 29.82% MAC

= 22.5% + 7.32%

Figure 9-20. Calculating the change in CG, using index arms.

CG = Weight shifted × Distance shifted

Total weight

= 2,500 × (724.9 – 352)

90,000

= 2,500 × 372.9

90,000

= 10.36 inches

Effects of Shifting Cargo From One Hold to 

Another

When cargo is shifted from one cargo hold to another, the CG 

changes, but the total weight of the aircraft remains the same.

For example, use the following data:

Loaded weight ................................................ 90,000 lb

Loaded CG............. station 580.97 (22.5 percent MAC)

Forward cargo hold centroid........................ station 352

Aft cargo hold centroid............................. station 724.9

MAC..................................................................141.5 in

LEMAC ....................................................... station 549

To determine the change in CG (∆CG) caused by shifting

2,500 pounds of cargo from the forward cargo hold to the 

aft cargo hold, use the formula in Figure 9-20.

Since the weight was shifted aft, the CG moved aft and the 

CG change is positive. If the shift were forward, the CG 

change would be negative.

Before the cargo was shifted, the CG was located at station 

580.97, which is 22.5 percent of MAC. The CG moved aft 

10.36 inches, so the new CG is found using the formula from 

Figure 9-21.

Convert the location of the CG in inches aft of the datum to 

percent MAC by using the formula in Figure 9-22. 

The new CG in percent MAC caused by shifting the cargo is 

the sum of the old CG plus the change in CG. [Figure 9-23]

Some AFMs locate the CG relative to an index point rather 

than the datum or the MAC. An index point is a location 

specified by the aircraft manufacturer from which arms used 

in weight and balance computations are measured. Arms 

measured from the index point are called index arms, and 

objects ahead of the index point have negative index arms, 

while those behind the index point have positive index arms. 

Use the same data as in the previous example, except for 

these changes:

Loaded CG.......................... index arm of 0.97, which is 

22.5 percent of MAC

Index point...................................fuselage station 580.0

Forward cargo hold centroid...............–227.9 index arm

Aft cargo hold centroid.......................+144.9 index arm

MAC..................................................................141.5 in

LEMAC ..............................................–30.87 index arm

The weight was shifted 372.8 inches (–227.9 + Δ = +144.9, 

Δ =372.8).

The change in CG can be calculated by using this formula 

found in Figure 9-24.
9-9

Figure 9-25. Determining the new CG, moved aft 10.36 inches.

New CG = Old CG ± CG

= 11.33 index arm

= 0.97 + 10.36

Figure 9-26. The change in the CG in percent MAC.

New CG % MAC = Old CG ± CG

= 29.82% MAC

= 22.5% + 7.32%

Figure 9-27. The new CG in percent MAC.

CG % MAC = × 100 ( ) CG inches

MAC

= × 100 ( ) 10.36

141.5

= 7.32% MAC

Figure 9-28. Determining pallet area in square feet.

Area (sq. ft.) = Length (inches) × Width (inches)

144 square inches/square foot

= 48.5 × 33.5

144

= 1,624.7

144

= 11.28 square feet

Since the weight was shifted aft, the CG moved aft, and the 

CG change is positive. If the shift were forward, the CG 

change would be negative. Before the cargo was shifted, 

the CG was located at 0.97 index arm, which is 22.5 percent 

MAC. The CG moved aft 10.36 inches, and the new CG is 

shown using the formula in Figure 9-25.

The change in the CG in percent MAC is determined by using 

the formula in Figure 9-26.

The new CG in percent MAC is the sum of the old CG plus 

the change in CG. [Figure 9-27]

Notice that the new CG is in the same location whether the 

distances are measured from the datum or from the index 

point.

Determining Cargo Pallet Loads and Floor 

Loading Limits

Each cargo hold has a structural floor loading limit based on 

the weight of the load and the area over which this weight is 

distributed. To determine the maximum weight of a loaded 

cargo pallet that can be carried in a cargo hold, divide its total 

weight, which includes the weight of the empty pallet and 

its tie down devices, by its area in square feet. This load per 

square foot must be equal to or less than the floor load limit.

In this example, determine the maximum load that can be 

placed on this pallet without exceeding the floorloading limit.

Pallet dimensions..........................................36 by 48 in

Empty pallet weight................................................47 lb

Tie down devices....................................................33 lb

Floor load limit ............................169 lb per square foot

The pallet has an area of 36 inches (3 feet) by 48 inches 

(4 feet), which equals 12 square feet, and the floor has a 

load limit of 169 pounds per square foot. Therefore, the total 

weight of the loaded pallet can be 169 × 12 = 2,028 pounds. 

Subtracting the weight of the pallet and the tie down devices 

gives an allowable load of 1,948 pounds (2,028 – [47 + 33]).

Determine the floor loading limit that is needed to carry a 

loaded cargo pallet having the following dimensions and 

weights:

Pallet dimensions ...................................48.5 by 33.5 in

Pallet weight ..........................................................44 lb

Tiedown devices ....................................................27 lb

Cargo weight .....................................................786.5 lb

First, determine the number of square feet of pallet area as 

shown in Figure 9-28.

Then, determine the total weight of the loaded pallet:

Pallet ....................................................................44.0 lb

Tiedown devices .................................................27.0 lb

Cargo .................................................................786.5 lb

Total...................................................................857.5 lb

Determine the load imposed on the floor by the loaded pallet. 

[Figure 9-29] The floor must have a minimum loading limit 

of 76 pounds per square foot.
9-10

Figure 9-29. Determining the load imposed on the floor by the 

loaded pallet.

Floor Load = Loaded weight

Pallet area

857.5

11.28

= 76.0 pounds/square foot

=

Figure 9-30. Finding the maximum takeoff weight.

Max limit

142,000

184,200

Landing weight

+ trip fuel

Takeoff weight

Trip limit 

142,000

+ 40,000

182,000

Figure 9-31. Determining zero fuel weight with lower trip limits.

Max limit

184,200

138,000

Landing weight

– fuel load

Zero fuel weight

Trip limit 

182,000

– 54,000

128,000

Figure 9-32. Finding maximum payload with lower trip limits.

Max limit

138,000 Zero fuel weight

– BOW

Payload (pounds)

Trip limit 

128,000

–100,500

27,500

Determining the Maximum Amount of Payload 

That Can Be Carried

The primary function of a transport or cargo aircraft is to carry 

payload, which is the portion of the useful load, passengers, 

or cargo that produces revenue. To determine the maximum 

amount of payload that can be carried, both the maximum 

limits for the aircraft and the trip limits imposed by the 

particular trip must be considered. In each of the following 

steps, the trip limit must be less than the maximum limit. If 

it is not, the maximum limit must be used.

These are the specifications for the aircraft in this example

Basic operating weight (BOW) .....................100,500 lb

Maximum zero fuel weight............................138,000 lb

Maximum landing weight..............................142,000 lb

Maximum takeoff weight ..............................184,200 lb

Fuel tank load .................................................54,000 lb

Estimated fuel burn en route............................40,000 lb

1. Compute the maximum takeoff weight for this trip. 

This is the maximum landing weight plus the trip fuel. 

[Figure 9-30]

2. The trip limit is lower than the maximum takeoff 

weight, so it is used to determine the zero fuel weight. 

[Figure 9-31]

3. The trip limit is again lower than the maximum takeoff 

weight, so use it to compute the maximum payload 

for this trip. [Figure 9-32]

Under these conditions, 27,500 pounds of payload may be 

carried.

Determining the Landing Weight

It is important to know the landing weight of the aircraft in 

order to set up the landing parameters and to be certain the 

aircraft is able to land safely at the intended destination.

In this example of a four-engine turboprop airplane, 

determine the airplane weight at the end of 4.0 hours of cruise 

under these conditions:

Takeoff weight...............................................140,000 lb

Pressure altitude during cruise..........................16,000 ft

Ambient temperature during cruise .....................–32 °C

Fuel burned during descent and landing............1,350 lb

Refer to the U.S. Standard Atmosphere Table in 

Figure 9-33 and the gross weight table in Figure 9-34 when 

completing the following steps:

1. Use the U.S. Standard Atmosphere Table to 

determine the standard temperature for 16,000 feet 

(–16.7 °C).

2. The ambient temperature is –32 °C, which is a 

deviation from standard of 15.3 °C. (–32° – (–16.7°) 

= –15.3°). It is below standard.

3. In the gross weight table, follow the vertical line 

representing 140,000 pounds gross weight upward 

until it intersects the diagonal line for 16,000 feet 

pressure altitude.

4. From this intersection, draw a horizontal line to the left 

to the temperature deviation index (0 °C deviation).

5. Draw a diagonal line parallel to the dashed lines for 

Below Standard from the intersection of the horizontal 

line and the Temperature Deviation Index.

6. Draw a vertical line upward from the 15.3 °C 

Temperature Deviation From Standard.
9-11

Figure 9-33. Standard atmosphere table.

Feet inHg mmHg PSI °C °F

0

2,000

4,000

6,000

8,000

10,000

12,000

14,000

16,000

18,000

20,000

22,000

24,000

26,000

28,000

30,000

32,000

34,000

36,000

38,000

40,000

42,000

44,000

46,000

48,000

50,000

55,000

60,000

29.92

27.82

25.84

23.98

22.23

20.58

19.03

17.58

16.22

14.95

13.76

12.65

11.61

10.64

9.74

8.90

8.12

7.40

6.73

6.12

5.56

5.05

4.59

4.17

3.79

3.44

2.71

2.14

760.0

706.7

656.3

609.1

564.6

522.7

483.4

446.5

412.0

379.7

349.5

321.3

294.9

270.3

237.4

226.1

206.3

188.0

171.0

155.5

141.2

128.3

116.6

105.9

96.3

87.4

68.6

54.4

14.7

13.66

12.69

11.78

10.92

10.11

9.35

8.63

7.96

7.34

6.75

6.21

5.70

5.22

4.78

4.37

3.98

3.63

3.30

2.99

2.72

2.47

2.24

2.02

1.85

1.68

1.32

1.04

15.0

11.0

7.1

3.1

−0.8

−0.4

−8.8

−12.7

−16.7

−20.7

−24.6

−28.6

−32.5

−36.5

−40.4

−44.4

−48.4

−52.4

−56.3

−56.5

−56.5

−56.5

−56.5

−56.5

−56.5

−56.5

59.0

51.9

44.7

37.6

30.5

23.3

16.2

9.1

1.9

−5.2

−12.3

−19.5

−26.6

−33.7

−40.9

−48.0

−55.1

−62.3

−69.4

−69.7

−69.7

−69.7

−69.7

−69.7

−69.7

−69.7

Temperature

remains constant

Figure 9-35. Determining the total fuel burn for a 4-hour cruise.

Total fuel burn =

Lb/hr/engine × No. engines × Hours flight duration

= 18,160 pounds

= 1,135 × 4 × 4

Figure 9-34. Gross weight table. Fuel flow—100 lb per hour per engine

Moment of the loaded aeroplane (lb in x 100)

20 15.3 10 0 80 90 100 110 120 130 140 150 160 170 180

16

15

14

13

12

11

10

9

8

7

6

11.35

Pressure altitude—1,000 feet

SL

4

8

12

16

20

36 32

28

24

TT limit

Gross weight—1,000 lb

7. Draw a horizontal line to the left from the intersection 

of the Below Standard diagonal and the 15.3 °C 

temperature deviation vertical line. This line crosses 

the fuel flow–100 pounds per hour per engine index 

at 11.35 and indicates that each of the four engines 

burns 1,135 (100 × 11.35) pounds of fuel per hour. 

The total fuel burn for the 4-hour cruise is shown in 

Figure 9-35.

The airplane gross weight was 140,000 pounds at takeoff 

with 18,160 pounds of fuel burned during cruise and 1,350 

pounds burned during the approach and landing phase. This 

leaves a landing weight of 140,000 – (18,160 + 1,350) = 

120,490 pounds.
9-12

Figure 9-36. Determining the amount of weight the aircraft must 

lose to reach the maximum allowable landing weight.

 171,000 lb cruise weight

– 142,500 lb maximum landing weight

 28,500 lb required reduction

Figure 9-37. Determining the amount of fuel burned from the 

beginning of the dump to touchdown.

Fuel flow = 3,170 lb/hr/engine

60

= 52.83 lb/min engine

Figure 9-38. Determining the amount of fuel needed to dump.

 28,500.00 lb required weight reduction

 – 2,007.54 lb fuel burned after start of dumping

 26,492.46 lb fuel to be dumped

Figure 9-39. Determine the time needed to dump fuel.

= 11.52 minutes 26,492.46

2,300

Determining Fuel Dump Time in Minutes

Most large aircraft are approved for a greater weight for 

takeoff than for landing. To make it possible for them to 

return to landing soon after takeoff, a fuel jettison system is 

sometimes installed. It is important in an emergency situation 

that the flight crew be able to dump enough fuel to lower the 

weight to its allowed landing weight. This is done by timing 

the dumping process.

In this example, the aircraft has two engines operating and 

these specifications apply

Cruise weight.................................................171,000 lb

Maximum landing weight..............................142,500 lb

Time from start of dump to landing.............. 19 minutes

Average fuel flow during

Dumping and descent .............................3,170 lb/hr/eng

Fuel dump rate .......................................2,300 lb/minute

To calculate the fuel dump time in minutes:

1. Determine the amount of weight the aircraft must 

lose to reach the maximum allowable landing weight. 

[Figure 9-36]

2. Determine the amount of fuel burned from the 

beginning of the dump to touchdown. [Figure 9-37]

For both engines, this is 52.83 × 2 = 105.66 lb/minute.

The engines burn 105.66 lbs of fuel per min for 19 minutes 

(the duration of the dump), which calculates to 2007.54 

pounds of fuel burned between the beginning of the dump 

and touchdown.

3. Determine the amount of fuel needed to dump by 

subtracting the amount of fuel burned during the dump 

from the required weight reduction. [Figure 9-38]

4. Determine the time needed to dump this amount of 

fuel by dividing the number of pounds of fuel to dump 

by the dump rate. [Figure 9-39]

Weight and Balance of Commuter Category 

Airplanes

The Beech 1900 is a typical commuter category airplane that 

can be configured to carry passengers or cargo. Figure 9-40

shows the loading data of this type of airplane in the 

passenger configuration

Determining the Loaded Weight and CG

As this airplane is prepared for flight, a manifest is prepared.

[Figure 9-41]
9-13

Figure 9-40. Loading data for passenger configuration. Row 1 F.S. 210 Crew F.S. 129

Fwd. press BHD F.S. 84.0

Aft press BHD F.S. 557.5

Row 2 F.S. 230Row 3 F.S. 260Row 4 F.S. 290Row 5 F.S. 320Row 6 F.S. 350Row 7 F.S. 380Row 8 F.S. 410Row 9 F.S. 440

Loading Data Standard Seating

F.S. 150.6F.S. 456.0F.S. 509.5

F.S. 43.0

F.S. 5177.35

F.S. 150.6F.S. 175.6F.S. 453.5F.S. 513.5

Aft baggage/cargo compt. (Aft section)

Aft baggage/cargo compt. (Fwd. section)

Aft partition

Baggage web

Fwd. cabin baggage compt.

Nose baggage compt.

Baggage/cargo location Maximum Weight Arm

Nose baggage compartment ................... ............................................150....................................................................................................65.5

Forward cabin baggage compartment.250................................. 163.5

 (includes up to 100 pounds that may be 

 suspended from clothes rod)

Aft baggage/cargo compartment Fwd. section ...................880....................................................................................................483.5

 Aft section.......................630....................................................................................................533.0

Note

For compartment loading that results in only partial utilization of total compartment volume, load items must be

distributed or secured in a manner that precludes shifting under normally anticipated operating conditions.

Figure 9-41. Determining the loaded weight and CG of a Beech 1900 in the passenger configuration.

Item Weight Arm Moment/100 CG

25,823

439

600

575

494

493

608

1,190

722

164

967

3,198

7,866

43,139 292.9

Airplane basic EW

Crew

Passengers

 Row 1

 Row 2

 Row 3

 Row 4

 Row 5

 Row 6

 Row 7

 Row 8

 Row 9

Baggage

 Nose

 Forward cabin

 Aft (forward section)

 Aft (aft section)

Fuel jet A @ +25 °C

 390 gallons 

9,226

340

300

250

190

170

190

340

190

100

200

600

2,633

14,729

129

210

230

260

290

320

350

380

410

440

65.5

163.6

483.5

533.0
9-14

Figure 9-42. Weight and moments—occupants.

Weight

 Moment/100

80

90

100

110

120

130

140

150

160

170

180

190

200

210

220

230

240

250

103

116

129

142

155

168

181

194

206

219

232

245

258

271

284

297

310

323

160

180

200

220

240

260

280

300

320

340

360

380

400

420

440

460

480

500

184

207

230

253

276

299

322

345

368

391

414

437

460

483

506

529

552

575

208

234

260

286

312

338

364

390

416

442

468

494

520

546

572

598

624

650

232

261

290

319

348

377

406

435

464

493

522

551

680

609

638

667

696

725

256

288

320

352

384

416

448

480

512

544

576

608

640

672

704

736

768

800

280

315

350

385

420

455

490

525

560

595

630

665

700

735

770

805

840

875

304

342

380

418

456

494

532

570

608

646

684

722

760

798

836

874

912

950

328

369

410

451

492

533

574

615

656

697

738

779

820

861

902

943

984

1025

352

396

440

484

528

572

616

660

704

748

792

836

880

924

968

1012

1056

1100

Useful Load Weights and Moments—Occupants

Crew Cabin seats

F.S.

129

F.S.

200

F.S.

230

F.S.

260

F.S.

290

F.S.

320

F.S.

350

F.S.

380

F.S.

410

F.S.

440

Figure 9-43. Weight and moments—baggage.

Weight

10

20

30

40

50

60

70

80

90

100

150

200

250

300

350

400

450

500

550

600

630

650

700

750

800

850

880

7

13

20

26

33

39

46

52

59

66

98

16

33

49

65

82

98

115

131

147

164

245

327

409

48

97

145

193

242

290

338

387

435

484

725

967

1209

1450

1692

1934

2176

2418

2659

2901

3046

3143

3384

3626

3868

4110

4255

53

107

160

213

266

320

373

426

480

533

800

1066

1332

1599

1866

2132

2398

2665

2932

3198

3358

 Moment/100

Useful Load Weights and Moments—Baggage

Nose 

baggage

compartment

F.S. 65.5

Forward

cabin

baggage

compartment

F.S. 163.6

AFT baggage/

cargo

compartment

(Forward

section)

F.S. 483.5

AFT baggage/

cargo

compartment

(Aft

section)

F.S. 533.0

Figure 9-44. Density variation of aviation fuel.

−40 −30 −20 −10 0 10 20 30 40

7.5

7.0

6.5

6.0

5.5

Aviation kerosene jet A & Jet A1

Jet B (JP-4)

Aviation gasoline grade 100/130

NOTE: The fuel quantity indicator 

is calibrated for correct indication 

when using aviation kerosene jet 

A and jet A1. When using other 

fuels, multiply the indicated fuel 

quantity in pounds by 0.99 for jet 

B (JP-4) or by aviation gasoline 

(100/130) to obtain actual fuel 

quantity in pounds.

Aviation 

kerosene 

Jet A and 

Jet A1 0.812

Jet B (JP-4) 0.785

AV gas grade 

100/130 0.703

Average 

specific

gravity

at 15 °C 

(59 °F)

Fuel 

Specific weight (lb/US gal)

Temperature (°C)

 1. The crew weight and the weight of each passenger is 

entered into the manifest. The moment/100 for each 

occupant is determined by multiplying the weight by 

the arm and dividing by 100. This data is available in 

the AFM and is shown in the Weight and Moments— 

Occupants table. [Figure 9-42]

2. The weight of the baggage in each compartment used 

is entered with its moment/100. This is determined 

in the Weights and Moments—Baggage table. 

[Figure 9-43]

3. Determine the weight of the fuel. Jet A fuel has 

a nominal specific gravity at +15 °C of 0.812

and weighs 6.8 pounds per gallon, but at +25 °C, 

according to the Density Variation of Aviation Fuel 

Chart [Figure 9-44], it weighs 6.75 lb/gal. Using this 
9-15

Figure 9-45. Weights and moments—usable fuel.

Gallons

 Weight Moment

100

6.5 lb/gal

 Weight Moment

100

6.6 lb/gal

 Weight Moment

100

6.7 lb/gal

 Weight Moment

100

6.8 lb/gal

10

20

30

40

50

60

70

80

90

100

110

120

130

140

150

160

170

180

190

200

210

220

230

240

250

260

270

280

290

300

310

320

330

340

350

360

370

380

390

400

410

420

425

65

130

195

260

325

390

455

520

585

650

715

780

845

910

975

1,040

1,106

1,170

1,235

1,300

1,365

1,430

1,495

1,560

1,625

1,690

1,755

1,820

1,885

1,950

2,015

2,080

2,145

2,210

2,275

2,340

2,405

2,470

2,535

2,600

2,665

2,730

2,763

197

394

592

789

987

1,185

1,383

1,581

1,779

1,977

2,175

2,372

2,569

2,765

2,962

3,157

3,351

3,545

3,739

3,932

4,124

4,315

4,507

4,698

4,889

5,080

5,271

5,462

5,651

5,842

6,032

6,225

6,417

6,610

6,802

6,995

7,188

7,381

7,575

7,768

7,962

8,156

8,259

66

132

198

264

330

396

462

528

594

660

726

792

858

924

990

1,056

1,122

1,188

1,254

1,320

1,386

1,452

1,518

1,584

1,650

1,716

1,782

1,848

1,914

1,980

2,046

2,112

1,278

2,244

2,310

2,376

2,442

2,508

2,574

2,640

2,706

2,772

2,805

200

401

601

802

1,002

1,203

1,404

1,605

1,806

2,007

2,208

2,409

2,608

2,808

3,007

3,205

3,403

3,600

3,797

3,992

4,187

4,382

4,576

4,770

4,964

5,158

5,352

5,546

5,738

5,932

6,125

6,321

6,516

6,711

6,907

7,103

7,299

7,495

7,691

7,888

8,085

8,282

8,386

67

134

201

268

335

402

469

536

603

670

737

804

871

938

1,005

1,072

1,139

1,206

1,273

1,340

1,407

1,474

1,541

1,608

1,675

1,742

1,809

1,876

1,943

2,010

2,077

2,144

2,211

2,278

2,345

2,412

2,479

2,546

2,613

2,680

2,747

2,814

2,848

203

407

610

814

1,018

1,222

1,426

1,630

1,834

2,038

2,242

2,445

2,648

2,850

3,053

3,254

3,454

3,654

3,854

4,053

4,250

4,448

4,646

4,843

5,040

5,236

5,433

5,630

5,825

6,022

6,218

6,416

6,615

6,813

7,011

7,210

7,409

7,609

7,808

8,007

8,207

8,407

8,513

68

136

204

272

340

408

476

544

612

680

748

816

884

952

1,020

1,088

1,156

1,224

1,292

1,360

1,428

1,496

1,564

1,632

1,700

1,768

1,836

1,904

1,972

2,040

2,108

2,176

2,244

2,312

2,380

2,448

2,516

2,584

2,652

2,720

2,788

2,856

2,890

206

413

619

826

1,033

1,240

1,447

1,654

1,861

2,068

2,275

2,482

2,687

2,893

3,099

3,303

3,506

3,709

3,912

4,113

4,314

4,514

4,715

4,915

5,115

5,315

5,514

5,714

5,912

6,112

6,311

6,512

6,713

6,915

7,116

7,318

7,520

7,722

7,924

8,127

8,330

8,532

8,640

Useful Load Weights and Moments—Usable Fuel

chart, determine the weights and moment/100 for 390 

gallons of Jet A fuel by interpolating between those 

for 6.7 lb/gal and 6.8 lb/gal. The 390 gallons of fuel at 

this temperature weighs 2,633 pounds, and its moment 

index is 7,866 lb-in/100.

4. Add all of the weights and all of the moment indexes. 

Divide the total moment index by the total weight, and 

multiply this by the reduction factor of 100. The total 

weight is 14,729 pounds; the total moment index is 

43,139 lb-in/100. The CG is located at fuselage station 

292.9. [Figure 9-45]

5. Check to determine that the CG is within limits for 

this weight. Refer to the Weight and Balance Diagram. 

[Figure 9-46] Draw a horizontal line across the 

envelope at 14,729 pounds of weight and a vertical 

line from the CG of 292.9 inches aft of the datum. 

These lines cross inside the envelope, verifying the 

CG is within limits for this weight.
9-16

Figure 9-46. Weight and balance diagram.

17,000

16,000

15,000

14,000

13,000

12,000

11,000

10,000

9,000

272 276 280 284 288 292 296 300 304

CG (inches aft of datum)

Weight (lb)

Maximum zero fuel weight

Maximum takeoff weight 16,600

Maximum landing weight 16,000

Maximum zero fuel weight 14,000

Maximum takeoff weight

Determining the Changes in CG When 

Passengers Are Shifted

Using the loaded weight and CG of the Beech 1900, calculate 

the change in CG when the passengers in rows 1 and 2 are 

moved to rows 8 and 9. [Figure 9-47] Note that there is no 

weight change, but the moment index has been increased by 

1,155 pound-inches/100 to 44,294. The new CG is at fuselage 

station 300.7. [Figure 9-48]

This type of problem is usually solved by using the following 

two formulas. The total amount of weight shifted is 550 

pounds (300 + 250) and both rows of passengers have moved 

aft by 210 inches (410 – 200 and 440 – 230). The CG has 

been shifted aft by 7.8 inches, and the new CG is at station 

300.7. [Figure 9-49]
9-17

Figure 9-47. Changes in CG caused by shifting passenger seats.

Item Weight Arm Moment/100 CG

(−) 600

(−) 575

(+) 1,230

(+) 1,100

43,139

(+) 1,155

44,294 300.7

 Row 1

 Row 2

 Row 8

 Row 9

Original conditions

Changes

New conditions

(−) 300

(−) 250

(+) 300

(+) 250

14,729

0

14,729

200

230

410

440

Figure 9-48. Determining the new CG at fuselage station.

CG = × 100 ( ) 43,139 + 1,155

14,729

= 300.7

Figure 9-49. Determining the new CG at station after CG has 

shifted aft.

CG = Weight shifted × Distance shifted

Total weight

= 550 × 210

14,729

= 7.8 inches

= 292.9 + 7.8

CG = Original CG + CG

= 300.7 inches aft of datum

Figure 9-50. One passenger configuration of a B737-800.

Mixed class:

 12 first class seats at 36-inch pitch

 148 economy class seats at 32-inch pitch

G

C

C

G

A

A

G

L

G

A

A L

A

In a large cabin aircraft with high-density seating such as 

the B737-800, the operator must account for the seating of 

passengers in the cabin [Figure 9-50]. If assigned seating 

is used to determine passenger location, the operator 

must implement procedures to ensure the assignment of 

passenger seating is incorporated into the loading procedure. 

It is recommended that the operator take into account the 

possibility that some passengers may not sit in their assigned 

seats. 

If the actual seating location of each passenger is not known, 

the operator may assume that all passengers are seated 

uniformly throughout the cabin or a specified subsection of 

the cabin. Reasonable assumptions can be made about the 

manner in which people distribute themselves throughout 

the cabin. For example, window seats are occupied first 

followed by aisle seats, followed by the remaining seats 

(window-aisle-remaining seating). Both forward and rear 

loading conditions should be considered. The passengers may 

fill up the window, aisle, and remaining seats from the front 

of the aircraft to the back, or the back to the front. 

If necessary, the operator may divide the passenger cabin 

into subsections or zones and manage the loading of each 

zone individually. It can be assumed that passengers will be 

sitting uniformly throughout each zone. 

Another consideration is the inflightmovement of passengers, 

crew, and equipment. It is assumed that all passengers, crew, 

and equipment are secured when the aircraft is in the takeoff 

or landing configuration. Standard operating procedures 
9-18

Figure 9-51. Loading data for cargo configuration.

Section A

Section B

Section C

Section D

Section E

Section F

Section G

Section H

Section I

Section K

Sec

J

F.S. 210

F.S. 240

F.S. 270

F.S. 300

F.S. 330

F.S. 360

F.S. 390

F.S. 420

F.S. 480

F.S. 513.5

F.S. 450

A 600 F.S. 225.0

B 600 F.S. 255.0

C 600 F.S. 285.0

D 600 F.S. 315.0

E 600 F.S. 345.0

F 600 F.S. 375.0

G 600 F.S. 405.0

H 600 F.S. 435.0

I 600 F.S. 465.0

J 250 F.S. 499.5

K 565 F.S. 533.0

Section

Notes:

1. All cargo in sections A 

 through J must be 

 supported on the seat

 tracks and secured to the

 seat tracks and side seat

 rails by an FAA-approved 

 system.

2. Concentrated cargo loads

 in sections A through L

 must not exceed 100 lb

 per square foot.

3. Cargo in sections K and L

 must be retained by 

 baggage webs and

 partitions provided as part 

 of standard airplane.

4. Any exception to the

 above procedures 

 requires approval by a local 

 FAA office. 

Centroid

arm

Maximum

structural

capacity

Loading cargo configuration

should be taken into account. Examples of items that can 

move during flight are:

• Flight deck crew members moving to the lavatory. 

• Flight attendants moving throughout the cabin. 

• Service carts moving throughout the cabin.

• Passengers moving throughout the cabin.

• Passengers moving to the lavatory.

Determining Changes in Weight and CG 

When the Aircraft Is Operated in Its Cargo 

Configuratio

To determine changes in weight and CG when the aircraft 

is operated in its cargo configuration, the Beech 1900 is 

used as an example. Figure 9-51 illustrates the airplane 

configuration. Notice that the arm of each cargo section is 

the centroid of that section.

The flight manifest of the Beech 1900 in the cargo 

confi uration is illustrated in Figure 9-52. The BOW includes 

the pilots and their baggage and there is no separate item 

for them. 

At the standard temperature of 15 °C, the fuel weighs 6.8 

pounds per gallon. Refer to Figure 9-45 to determine the 

weight and moment index of 370 gallons of Jet A fuel. The 

CG under these loading conditions is located at station 296.2.

Determining the CG Shift When Cargo Is Moved 

From One Section to Another

To calculate the CG when cargo is shifted from one section 

to another, use the formula found in Figure 9-53. If the cargo 

is moved forward, the CG is subtracted from the original CG. 

If the cargo is shifted aft, add the CG to the original.

Determining the CG Shift When Cargo Is Added 

or Removed

To calculate the CG when cargo is added or removed, add or 

subtract the weight and moment index of the affected cargo to 

the original loading chart. Determine the new CG by dividing 

the new moment index by the new total weight, and multiply 

this by the reduction factor. [Figure 9-54]

Determining Which Limits Are Exceeded

When preparing an aircraft for flight, consider all parameters 

and check to determine that no limits have been exceeded. 

Consider the parameters below, and determine which limit, 

if any, has been exceeded.

• The aircraft in this example has a basic empty weight 

of 9,005 pounds and a moment index of 25,934 pound 

inches/100.

• The crew weight is 340 pounds and its moment/100 

is 439.

• The passengers and baggage have a weight of 3,950 

pounds and a moment/100 of 13,221.

• The fuel is computed at 6.8 lb/gal. The ramp load is 

340 gallons or 2,312 pounds. Fuel used for start and 

taxi is 20 gallons, or 136 pounds. Fuel remaining at 

landing is 100 gallons, or 680 pounds.

• Maximum takeoff weight is 16,600 pounds.

• Maximum zero fuel weight is 14,000 pounds.

• Maximum landing weight is 16,000 pounds.
9-19

Figure 9-53. Shifting cargo from one section to another.

 CG = Weight shifted x Distance shifted

Total weight

Figure 9-52. Flight manifest of a Beech 1900 in the cargo configuration.

Item Weight Arm Moment/100 CG

25,934

675

1,020

1,283

1,890

2,070

2,250

810

7,520

43,452 296.2

BOW

Cargo section A

Cargo section B

Cargo section C

Cargo section D

Cargo section E

Cargo section F

Cargo section G

Cargo section H

Cargo section I

Cargo section J

Cargo section K

Fuel jet A at +15 °C

 gallons 370

9,005

300

400

450

600

600

600

200

2,516

14,671

225.0

255.0

285.0

315.0

345.0

375.0

405.0

435.0

465.0

499.5

533.0

Figure 9-54. Determining the new CG by dividing the new moment 

index by the new total weight, multiplied by the reduction factor.

 CG = x Reduction factor Total moment index

Total weight

Figure 9-55. Determining the zero fuel weight.

Item Weight (lb) Moment CG

Basic empty weight

Crew

Passengers & baggage

Zero fuel weight

9,005

340

3,950

13,295

25,934

439

13,221

39,594

Figure 9-56. Determining the takeoff weight and CG.

Item Weight (lb) Moment CG

Zero fuel weight

Takeoff fuel 320 gal

Ramp load−fuel for 

start & taxi

340 − 20 = 320 gal

Takeoff weight

13,295

2,176

15,471

39,594

6,512

46,106 298.0

Figure 9-57. Determining the landing weight and CG.

Item Weight (lb) Moment CG

Zero fuel weight

Fuel at landing 100 gal

Landing weight

13,295

680

13,975

39,594

1,977

41,571 297.5

Take these steps to determine which limit, if any, is exceeded:

1. Determine the zero fuel weight, which is the weight 

of the aircraft with all of the useful load except the 

fuel onboard. [Figure 9-55]

The zero fuel weight of 13,295 pounds is less than 

the maximum of 14,000 pounds, so this parameter is 

acceptable.

2. Determine the takeoff weight and CG. The takeoff 

weight is the zero fuel weight plus the weight of the 

ramp load of fuel, minus the weight of the fuel used 

for start and taxi. The takeoff CG is the moment/100 

divided by the weight, and then the result multiplied 

by 100. The takeoff weight of 15,471 pounds is below 

the maximum takeoff weight of 16,600 pounds, and a 

check of the weight and balance diagram shows that the 

CG at station 298.0 is also within limits. [Figure 9-56]

3. Determine the landing weight and CG. This is the zero fuel 

weight plus the weight of fuel at landing. [Figure 9-57]

The landing weight of 13,975 pounds is less than the 

maximum landing weight of 14,000 to 16,000 pounds. 

According to the weight and balance diagram, the 

landing CG at station 297.5 is also within limits.
9-20
10-1

Introduction

Almost all weight and balance problems involve only simple 

math. This allows slide rules and hand-held electronic 

calculators to relieve much of the tedium involved with these 

problems. This chapter compares the methods of determining 

the center of gravity (CG) of an airplane while it is being 

weighed. First, it shows how to determine the CG using a 

simple electronic calculator, then solves the same problem 

using an E6-B flight computer. Finally, it shows how to solve 

it using a dedicated electronic flight computer

Use of Computer for Weight 

and Balance Computations

Chapter 10
10-2

Figure 10-2. Weight and balance data of a typical nosewheel 

airplane.

Weighing Point Weight (lb) Arm (in)

0

0

–78

830

836

340

2,006

Right side

Left side

Nose

Total

Figure 10-1. A typical electronic calculator is useful for solving 

most types of weight and balance problems.

Examples of typical weight and balance problems (solved 

with an electronic calculator) that pilots and airframe and 

powerplant (A&P) maintenance technicians encounter 

throughout their aviation endeavors are shown later in this 

chapter.

Electronic Calculator

Determining the CG of an airplane in inches for the mainwheel weighing points can be done with any simple electronic 

calculator that has addition (+), subtraction (–), multiplication 

(×), and division (÷) functions. [Figure 10-1] Scientifi 

calculators with such additional functions as memory (M), 

parentheses (( )), plus or minus (+/–), exponential (yx), 

reciprocal (1/x), and percentage (%) functions allow more 

complex problems to be solved or simple problems to be 

solved using fewer steps.

According to Figure 10-2, the weight of the nosewheel (F) is 

340 pounds, the distance between main wheels and nosewheel 

(L) is –78 inches, and the total weight (W) of the airplane is 

2,006 pounds. L is negative because the nosewheel is ahead 

of the main wheels.

To determine the CG, use this formula:

CG = F × L

W

 = 340 × –78

 2,006

Key the data into the calculator and press the equal (=) key.

The answer should read as follows:

(340)(×)(78)(+/–)(÷)(2006)(=) –13.2

The arm of the nosewheel is negative, so the CG is –13.2 or 

13.2 inches ahead of the main wheel weighing points.

E6-B Flight Computer

The E6-B uses a special kind of slide rule. Instead of its scales 

going from 1 to 10, as on a normal slide rule, both scales 

go from 10 to 100. The E6-B cannot be used for addition or 

subtraction, but it is useful for making calculations involving 

multiplication and division. Its accuracy is limited, but it is 

sufficiently accurate for most weight and balance problems.

The same problem that was just solved with the electronic 

calculator can be solved on an E6-B by following these steps:

CG = F × L

 W

 = 340 × –78

 2,006

First, multiply 340 by 78 (disregard the minus sign) 

[Figure 10-3]: 

• Step 1—place 10 on the inner scale (this is the index 

opposite 34 on the outer scale that represents 340).

• Step 2—opposite 78 on the inner scale, read 26.5 on 

the outer scale.

Determine the value of these digits by estimating: 

300 × 80 = 24,000, so 340 × 78 = 26,500.

Then, divide 26,500 by 2,006 [Figure 10-4]: 
10-3

Figure 10-3. E6-B computer set up to multiply 340 by 78.

Figure 10-4. E6-B computer set up to divide 26,500 by 2,006.

Step 2

Step 1

Step 4

Step 3

Figure 10-5. Dedicated electronic flight computers are programmed 

to solve weight and balance problems, as well as flight problems.

• Step 3—on the inner scale, place 20, which represents 

2,006 opposite 26.5 on the outer scale (26.5 represents 

26,500).

• Step 4—opposite the index, 10, on the inner scale read 

13.2 on the outer scale.

Determine the value of 13.2 by estimating: 

20,000 ÷ 2000 = 10, so 26,500 ÷ 2,006 = 13.2.

The arm (–78) is negative, so the CG is also negative.

The CG is –13.2 inches or 13.2 inches ahead of the datum.

Dedicated Electronic Flight Computer

Dedicated electronic flight computers, like the one in 

Figure 10-5, are programmed to solve many flight problems 

such as wind correction, heading and ground speed, 

endurance, and true airspeed (TAS), as well as weight and 

balance problems.

The problem just solved with an electronic calculator and 

an E6-B can also be solved with a dedicated flight computer 

using the information shown in Figure 10-2. Each fligh 

computer handles the problems in a slightly different way, 

but all are programmed with prompts that solicit the required 

data to be inputted so there is no need to memorize any 

formulas. Weight and arms are inputted as called for, and a 

running total of the weight, moment, and CG are displayed.
10-4

Typical Weight and Balance Problems

A hand-held electronic calculator like the one in Figure 10-5

is a valuable tool for solving weight and balance problems. It 

can be used for a variety of problems and has a high degree 

of accuracy. The examples given here are solved with a 

calculator using only the (×),(÷),(+),( – ), and (+/–) functions. 

If other functions are available on your calculator, some of 

the steps may be simplified

Determining CG in Inches From the Datum

This type of problem is solved by firs determining the 

location of the CG in inches from the main wheel weighing 

points, then measuring this location in inches from the datum. 

There are four types of problems involving the location of 

the CG relative to the datum.

Nosewheel Airplane With Datum Ahead of the 

Main Wheels

The datum (D) is 128 inches ahead of the main wheel 

weighing points; the weight of the nosewheel (F) is 340 

pounds, and the distance between main wheels and nosewheel 

(L) is 78 inches. The total weight (W) of the airplane is 2,006 

pounds. Refer to Figure 3-5 on page 3-5.

Use this formula:

CG = D – F × L

 W 

1. Determine the CG in inches from the main wheel:

(340)(×)(78)(÷)(2006)(=) 13.2

2. Determine the CG in inches form the datum:

(128)(–)(13.2)(=) 114.8

The CG is 114.8 inches behind the datum.

Nosewheel Airplane With Datum Behind the Main 

Wheels

The datum (D) is 75 inches behind the main wheel weighing 

points, the weight of the nosewheel (F) is 340 pounds, and 

the distance between main wheels and nosewheel (L) is 78 

inches. The total weight (W) of the airplane is 2,006 pounds. 

Refer to Figure 3-6 on page 3-5.

Use this formula:

CG = – D + F × L

 W 

1. Determine the CG in inches from the main wheels:

(340)(×)(78)(÷)(2006)(=) 13.2

2. Determine the CG in inches from the datum:

(75)(+)(13.2)(=) 88.2

The minus sign before the parenthesis in the formula 

means the answer is negative. The CG is 88.2 inches 

ahead of the datum (–88.2).

Tailwheel Airplane With Datum Ahead of the Main 

Wheels

The datum (D) is 7.5 inches ahead of the main wheel 

weighing points, the weight of the tailwheel (R) is 67 pounds, 

and the distance between main wheels and tailwheel (L) is 

222 inches. The total weight (W) of the airplane is 1,218 

pounds. Refer to Figure 3-7 on page 3-6.

Use this formula:

CG = D + R × L

 W 

1. Determine the CG in inches from the main wheels.

(67)(×)(222)(÷)(1218)(=) 12.2

2. Determine the CG in inches from the datum:

(7.5)(+)(12.2)(=) 19.7

The CG is 19.7 inches behind the datum.

Tailwheel Airplane With Datum Behind the Main 

Wheels

The datum (D) is 80 inches behind the main wheel weighing 

points, the weight of the tailwheel (R) is 67 pounds, and the 

distance between main wheels and tailwheel (L) is 222 inches. 

The total weight (W) of the airplane is 1,218 pounds. Refer 

to Figure 3-8 on page 3-6.

Use this formula:

CG = –D + R × L

 W 

1. Determine the CG in inches from the main wheels:

(67)(×)(222)(÷)(1218)(=) 12.2

2. Determine the CG in inches from the datum:

(80)(+/–)(+)(12.2)(=) –67.8

The CG is 67.8 inches ahead of the datum.

( )

( )

( )

( )
10-5

Figure 10-6. Specifications for determining the CG of an airplane using weight and arm.

Figure 10-7. Specifications for determining the CG of an airplane 

using weights and moment indexes.

Weight Point Weight (lb) Arm (in) Moment (lb-in) CG

106,240

107,008

17,000

230,248 114.8

Right side

Left side

Nose

Total

830

836

340

2,006

128

128

50

Weighing Point Weight (lb) Moment/100 CG

114.8

Right side

Left side

Nose

Total

830

836

340

2,006

1,062.4

1,070.1

170

2,302.5

Determining CG, Given Weights, and Arms

Some weight and balance problems involve weights and arms 

to determine the moments. Divide the total moment by the 

total weight to determine the CG. Figure 10-6 contains the 

specificationsfor determining the CG using weights and arms.

Determine the CG by using the data in Figure 10-6 and 

following these steps:

1. Determine the total weight and record this number:

(830)(+)(836)(+)(340)(=) 2,006

2. Determine the moment of each weighing point and 

record them:

(830)(×)(128)(=) 106,240

(836)(×)(128)(=) 107,008

(340)(×)(50)(=) 17,000

3. Determine the total moment and divide this by the 

total weight:

(106240)(+)(107008)(+)(17000)(=)(÷)(2006)(=) 

114.8

This airplane weighs 2,006 pounds and its CG is 114.8 inches 

from the datum.

Determining CG, Given Weights, and Moment 

Indexes

Other weight and balance problems involve weights and 

moment indexes, such as moment/100 or moment/1,000. To 

determine the CG, add all the weights and all the moment 

indexes. Then, divide the total moment index by the total 

weight and multiply the answer by the reduction factor. 

Figure 10-7 contains the specifications for determining the 

CG using weights and moments indexes.

Determine the CG by using the data in Figure 10-7 and 

following these steps:

1. Determine the total weight and record this number:

(830)(+)(836)(+)(340)(=) 2,006

2. Determine the total moment index, divide this by the 

total weight, and multiply it by the reduction factor 

of 100:

(1062.4)(+)(1070.1)(+)(170)(=)(2302.5)(÷)(2006)(=)

(1.148)(×)(100)(=) 114.8

This airplane weighs 2,006 pounds and its CG is 114.8 inches 

from the datum.

Determining CG in Percent Mean Aerodynamic 

Chord (MAC)

• The loaded CG is 42.47 inches aft of the datum.

• MAC is 61.6 inches long.

• LEMAC is at station 20.1.

1. Determine the distance between the CG and LEMAC:

(42.47)(–)(20.1)(=) 22.37

2. Then, use this formula:

(22.37)(×)(100)(÷)(61.6)(=) 36.3

The CG of this airplane is located at 36.3 percent MAC.

Determining Lateral CG of a Helicopter

For a helicopter, it is often necessary to determine not only 

the longitudinal CG, but the lateral CG as well. Lateral CG is 

measured from butt line zero (BL 0). All items and moments 

to the left of BL 0 are negative, and all those to the right of 

BL 0 are positive. Figure 10-8 contains the specifications for 

determining the lateral CG of a typical helicopter.
10-6

Figure 10-8. Specifications for determining the lateral CG of a helicopter.

Item Weight (lb) Lateral Arm (in) Lateral Offset Moment (lb-in) Lateral CG

309

−2,295

2,700

−2,419

−1,705 –0.77

Helicopter empty weight

Pilot

Passenger

Fuel (48 gal)

Total

1,545

170

200

288

2,203

+0.2

 –13.5

+13.5

135.0

–8.4

x =

Determine the lateral CG by using the data in Figure 10-8 and 

following these steps:

1. Add all of the weights:

(1545)(+)(170)(+)(200)(+)(288)(=) 2,203

2. Multiply the lateral arm (the distance between butt 

line zero and the CG of each item) by its weight to 

get the lateral offset moment of each item. Moments 

to the right of BL 0 are positive and those to the left 

are negative.

(1,545)(×)(.2)(=) 309

(170)(×)(13.5)(+/–)(=) –2,295

(200)(×)(13.5)(=) 2,700

(288)(×)(8.4)(+/–)(=) –2,419

3. Determine the algebraic sum of the lateral offset 

moments.

(309)(+)(2295)(+/–)(+)(2700)(+)(2419)(+/–)(=) 

–1,705

4. Divide the sum of the moments by the total weight to 

determine the lateral CG.

(1705)(+/–)(÷)(2203)(=) –0.77

The lateral CG is 0.77 inch to the left of BL0.

Determining ΔCG Caused by Shifting Weights

Fifty pounds of baggage is shifted from the aft baggage 

compartment at station 246 to the forward compartment at 

station 118. The total airplane weight is 4,709 pounds. How 

much does the CG shift?

1. Determine the number of inches the baggage is shifted:

(246)(–)(118)(=) 128

2. Use this formula:

  CG = Weight shifted × Distance weight is shifte

 Total weight

(50)(×)(128)(÷)(4709)(=) 1.36

The CG is shifted forward 1.36 inches.

Determining Weight Shifted to Cause Specified

ΔCG

How much weight must be shifted from the aft baggage 

compartment at station 246 to the forward compartment 

at station 118 to move the CG forward 2 inches? The total 

weight of the airplane is 4,709 pounds.

1. Determine the number of inches the baggage is shifted:

 Weight shifted =

 Total weight shifted ×  C

 Distance weight is shifted

(246)(–)(118)(=) 128

2. Use this formula:

(2)(×)(4709)(÷)(128)(=) 73.6

Moving 73.6 pounds of baggage from the aft compartment 

to forward compartment shifts the CG forward 2 inches.

Determining Distance Weight Is Shifted to Move 

CG a Specific Distanc

How many inches aft does a 56 pound battery need to be 

moved to shift the CG aft by 1.5 inches? The total weight of 

the airplane is 4,026 pounds.

Use this formula:

Distance weight is shifted = Total weight ×  C

 Weight shifted

(1.5)(×)(4026)(÷)(56)(=) 107.8

Moving the battery aft by 107.8 inches shifts the CG aft 1.5 

inches.
10-7

Determining Total Weight of an Aircraft With a 

Specified ΔCG When Cargo Is Move

What is the total weight of an airplane if moving 500 pounds 

of cargo 96 inches forward shifts the CG 2.0 inches?

Use this formula:

Total weight = Weight shifted × Distance weight is shifted

  C

(500)(×)(96)(÷)(2)(=) 24,000

Moving 500 pounds of cargo 96 inches forward causes a 

2.0-inch shift in CG of a 24,000-pound airplane.

Determining Amount of Ballast Needed to Move 

CG to a Desired Location

How much ballast must be mounted at station 228 to move 

the CG to its forward limit of +33? The airplane weighs 

1,876 pounds and the CG is at +32.2, a distance of 0.8 inch 

out of limit.

Use this formula:

Ballast weight = Aircraft empty weight × Dist. out of limits 

 Distance ballast to desired CG

(1876)(×)(.8)(÷)(195)(=) 7.7

Attaching 7.7 pounds of ballast to the bulkhead at station 

228 moves the CG to +33.0.
10-8
A-1

FAA Publications 

(check for most current revision)

• Advisory Circular (AC) 20-161, Aircraft Onboard Weight and Balance Systems

• AC 23-21, Airworthiness Compliance Checklists Used to Substantiate Major Alterations for Small Airplanes

• AC 43.13-1, Acceptable Methods, Techniques, and Practices—Aircraft Inspection and Repair

• AC 90-89, Amateur-Built Aircraft and Ultralight Flight Testing Handbook

• AC 120-27, Aircraft Weight and Balance Control

• FAA-H-8083-30, Aviation Maintenance Technician Handbook—General 

• Order 8130.2, Airworthiness Certification of Aircraft and Related Product

• www.faa.gov, Light Sport Aircraft (multiple electronic documents)

Appendix A

Supplemental Study Materials for Aircraft Weight and Balance
A-2
B-1

Station 0 Station 50

Station 0 Station 70

10 lb

MAC

(+)

Arm

(–)

Arm

(+)

Arm

(+) Arm

70"

LEMAC

TEMAC

Aft limit

Forward limit

Datum

ation 50

MAC

TEM

LEM

CG

CG

range

(–)

A

(+)

A

Aft li

m

Forward

Multiengine aircraft

Single-engine aircraft

Datum

Datum

Weight

Moment = 700 lb–in 

Appendix B

Key Weight and Balance Locations on an Aircraft
B-2
G-1

A note on glossary terms: over the years there has been a 

proliferation of aircraft weight and balance terms. This is 

the result of many factors, such as the Federal Aviation 

Administration (FAA) certification regulation under which 

an aircraft was constructed; the FAA regulation under which 

the aircraft is being operated; manufacturer standardization 

agreements or a combination of these and others. Examples 

of such terms are empty weight, licensed empty weight, basic 

empty weight, operational empty weight, etc.

Many of these terms may have similar meanings or sound 

similar. Pilots and aircraft mechanics must ensure they 

understand the terminology and are applying the correct 

values based on the procedure and situation dictating the 

calculations undertaken.

Occasionally, the acronym “GAMA” follows the glossary 

term. This indicates that it is part of the standardized format 

adopted by the General Aviation Manufacturers Association 

in 1976, known as GAMA Specification No.1. These aircraft 

are usually manufactured under Title 14 of the Code of 

Federal Regulations (14 CFR) part 23 and differ from aircraft 

manufactured under the earlier certification rule Civil Air 

Regulation (CAR) part 3 for weight and balance purposes 

in the condition under which empty weight was established.

Aircraft certific ted under 14 CFR parts 23, 25, 27, and 29 

establish their certificated empty weight as required in the 

appropriate section of these regulations, which states that the 

empty weight and corresponding center of gravity (CG) must 

be determined by weighting the aircraft with:

• Fixed ballast

• Unusable fuel

• Full operating flu d, including oil, hydraulic flu d, 

and other fluids required for normal operation of 

the aircraft systems, except potable water, lavatory 

precharge water, and water intended for injection in 

the engine(s).

In contrast, aircraft certificated under the older CARs 

established certifi ated empty weight under similar conditions 

Glossary

as the newer aircraft, with the important exception that the 

aircraft weight did not include full oil, only undrainable oil. 

Mechanics and repair stations should consult the appropriate 

certification rule when reestablishing empty weight

14 CFR part 121. The Federal regulations governing 

domestic, flag, and supplemental operations

14 CFR part 135. The Federal regulations governing 

commuter and on-demand operations.

Adverse loaded CG check. A weight and balance check to 

determine that no condition of legal loading of an aircraft 

can move the CG outside of its allowable limits.

Aircraft specifications. Documentation containing the 

pertinent specifications for aircraft certificated under the 

CARs.

Airplane Flight manual (AFM). An FAA-approved 

document, prepared by the holder of a type certificate for an 

aircraft, that specifies the operating limitations and contains 

the required markings and placards and other information 

applicable to the regulations under which the aircraft was 

certificated

Approved type certificate. A certificate of approval issued 

by the FAA for the design of an aircraft, engine, or propeller.

Arm. (GAMA) The horizontal distance from the reference 

datum to the CG of an item. The algebraic sign is plus (+) if 

measured aft of the datum or to the right side of the center 

line when considering a lateral calculation. The algebraic 

sign is minus (–) if measured forward of the datum or the left 

side of the center line when considering a lateral calculation.

Balanced laterally. Being balanced in such a way that the 

aircraft wings tend to remain level.

Ballast. A weight installed or carried in an aircraft to move 

the center of gravity to a location within its allowable limits.
G-2

Permanent ballast (fixed ballast). A weight permanently 

installed in an aircraft to bring its CG into allowable limits. 

Permanent ballast is part of the aircraft empty weight.

Temporary ballast. Weights that can be carried in a cargo 

compartment of an aircraft to move the location of CG for a 

specifi fl ght condition. Temporary ballast must be removed 

when the aircraft is weighed.

Basic empty weight. (GAMA) Standard empty weight plus 

optional equipment.

Basic operating index. The moment of the airplane at its 

basic operating weight divided by the appropriate reduction 

factor.

Basic operating weight (BOW). The empty weight of the 

aircraft plus the weight of the required crew, their baggage, 

and other standard items, such as meals and potable water.

Bilge area. The lowest part of an aircraft structure in which 

water and contaminants collect.

Butt (or buttock) line zero. A line through the symmetrical 

center of an aircraft from nose to tail. It serves as the datum 

for measuring the arms used to determine the lateral CG. 

Lateral moments that cause the aircraft to rotate clockwise are 

positive (+) , and those that cause it to rotate counterclockwise 

are negative (–).

Calendar month. A time period used by the FAA for 

certification and currency purposes. A calendar month 

extends from a given day until midnight of the last day of 

that month.

Civil Air Regulations (CAR). The predecessor to the Federal 

Aviation Regulations.

CAM. The acronym for the manuals containing the 

certification rules under the Civil Air Regulations

Center of gravity (CG). (GAMA) The point at which an 

airplane would balance if suspended. Its distance from the 

reference datum is determined by dividing the total moment 

by the total weight of the airplane. It is the mass center of the 

aircraft, or the theoretical point at which the entire weight of 

the aircraft is assumed to be concentrated. It may be expressed 

in percent of MAC (mean aerodynamic cord) or in inches 

from the reference datum.

Center of lift. The location along the chord line of an airfoil at 

which all the lift forces produced by the airfoil are considered 

to be concentrated.

Centroid. The distance in inches aft of the datum of the 

center of a compartment or a fuel tank for weight and balance 

purposes.

CG arm. (GAMA) The arm obtained by adding the airplane’s 

individual moments and dividing the sum by the total weight.

CG limits. (GAMA) The extreme CG locations within which 

the aircraft must be operated at a given weight. These limits 

are indicated on pertinent FAA aircraft type certificate data 

sheets, specifications, or weight and balance records

CG limits envelope. An enclosed area on a graph of the 

airplane loaded weight and the CG location. If lines drawn 

from the weight and CG cross within this envelope, the 

airplane is properly loaded.

CG moment envelope. An enclosed area on a graph of 

the airplane loaded weight and loaded moment. If lines 

drawn from the weight and loaded moment cross within this 

envelope, the airplane is properly loaded.

Chord. A straight line distance across a wing from leading 

edge to trailing edge.

Curtailment. An operator created and FAA-approved 

operational loading envelope that is more restrictive than 

the manufacturer’s CG envelope. It ensures that the aircraft 

will be operated within limits during all phases of flight 

Curtailment typically accounts for, but is not limited to, inflight movement of passengers and crew, service equipment, 

cargo variation, seating variation, etc.

Datum. An imaginary vertical plane or line from which all 

measurements of arms are taken. The datum is established 

by the manufacturer. Once the datum has been selected, all 

moment arms and the location of CG range are measured 

from this point.

Delta (Δ). The Greek symbol “Δ” means a change in 

something. ΔCG means a change in the center of gravity 

location.

Dynamic load. The actual weight of the aircraft multiplied 

by the load factor, or the increase in weight caused by 

acceleration.

Empty weight. The weight of the airframe, engines, 

all permanently installed equipment, and unusable fuel. 

Depending upon the part of the federal regulations under 

which the aircraft was certificated, either the undrainable oil 

or full reservoir of oil is included.
G-3

Empty weight center of gravity (EWCG). This is the center 

of gravity of the aircraft in the empty condition, containing 

only the items specified in the aircraft empty weight. This 

CG is an essential part of the weight and balance record of 

the aircraft.

Empty weight center of gravity range. The distance 

between the allowable forward and aft empty-weight CG 

limits.

Equipment list. A list of items approved by the FAA for 

installation in a particular aircraft. The list includes the name, 

part number, weight, and arm of the component. Installation 

or removal of an item in the equipment list is considered to 

be a minor alteration.

Fleet weight. An average weight accepted by the FAA for 

aircraft of identical make and model that have the same 

equipment installed. When a fleet weight control program 

is in effect, the fleet weight of the aircraft can be used rather 

than requiring every individual aircraft to be weighed.

Fuel jettison system. A fuel subsystem that allows the fligh 

crew to dump fuel in an emergency to lower the weight of an 

aircraft to the maximum landing weight if a return to landing 

is required before sufficient fuel is burned off. This system 

must allow enough fuel to be jettisoned that the aircraft can 

still meet the climb requirements specified in 14 CFR part 25.

Fulcrum. The point about which a lever balances.

Index point. A location specifiedby the aircraft manufacturer 

from which arms used in weight and balance computations 

are measured. Arms measured from the index point are called 

index arms.

Interpolate. Determine a value in a range between two 

known values.

Landing weight. The takeoff weight of an aircraft minus the 

fuel burned and/or dumped en route.

Large aircraft (14 CFR part 1). An aircraft of more than 

12,500 pounds maximum certificated takeoff weight

Lateral balance. The balance around the roll, or longitudinal, 

axis.

Lateral offset moment. The moment, in pound inches, of a 

force that tends to rotate a helicopter about its longitudinal 

axis. The lateral offset moment is the product of the weight of 

the object and its distance from butt line zero. Lateral offset 

moments that tend to rotate the aircraft clockwise are positive, 

and those that tend to rotate it counterclockwise are negative.

LEMAC. Leading edge of the mean aerodynamic chord. A 

reference point for measurements, and specified in inches 

from the datum to allow computations to relate percent MAC 

to the datum.

Load cell. A component in an electronic weighing system that 

is placed between the jack and the jack pad on the aircraft. 

The load cell contains strain gauges whose resistance changes 

with the weight on the cell.

Load factor. The ratio of the maximum load an aircraft can 

sustain to the total weight of the aircraft. Normal category 

aircraft must have a load factor of a least 3.8; utility category 

aircraft, 4.4; and acrobatic category aircraft, 6.0.

Loading graph. A graph of load weight and load moment 

indexes. Diagonal lines for each item relate the weight to the 

moment index, eliminating the need for calculations.

Loading schedule. A method of calculating and documenting 

aircraft weight and balance prior to taxiing to ensure the 

aircraft will remain within all required weight and balance 

limitations throughout the flight

Longitudinal axis. An imaginary line through an aircraft 

from nose to tail, passing through its center of gravity.

Longitudinal balance. The balance around the pitch, or 

lateral, axis.

MAC. Mean aerodynamic chord.

Major alteration. An alteration not listed in the aircraft, 

aircraft engine, or propeller specifications (1) that might 

appreciably affect weight, balance, structural strength, 

performance, powerplant operation, flight characteristics, 

or other qualities affecting airworthiness; or (2) that is not 

done according to accepted practices or cannot be done by 

elementary operations.

Maximum landing weight. (GAMA) The maximum weight 

approved for the landing touchdown.

Maximum permissible hoist load. The maximum external 

load that is permitted for a helicopter to carry. This load is 

specified in the POH

Maximum ramp weight. (GAMA) The maximum weight 

approved for ground maneuver. It includes weight of start, 

taxi, and runup fuel.

Maximum takeoff weight. (GAMA) The maximum weight 

approved for the start of the takeoff run.
G-4

Maximum taxi weight. The maximum weight approved 

for ground maneuvers. This is the same as maximum ramp 

weight.

Maximum weight. The maximum authorized weight of 

the aircraft and all of its equipment as specified in the Type 

Certificate Data Sheets (TCDS) for the aircraft

Maximum zero fuel weight. The maximum authorized 

weight of an aircraft without fuel. This is the total weight 

for a particular flight minus the fuel. It includes the aircraft 

and everything that is carried on the flight except the weight 

of the fuel.

Mean aerodynamic chord (MAC). The average distance 

from the leading edge to the trailing edge of the wing.

METO horsepower (maximum except takeoff HP). The 

maximum power allowed to be produced continuously by an 

engine. Takeoff power is usually limited to a given amount 

of time, such as 1 minute or 5 minutes.

Minimum fuel. The amount of fuel necessary for one-half 

hour of operation at the rated maximum-continuous power 

setting of the engine, which, for weight and balance purposes, 

is 1

⁄12 gallon per maximum-except-takeoff (METO) horsepower. It is the maximum amount of fuel that could be used 

in weight and balance computations when low fuel might 

adversely affect the most critical balance conditions. To 

determine the weight of the minimum fuel in pounds, divide 

the METO horsepower by two.

Minor alteration. An alteration other than a major alteration. 

This includes alterations that are listed in the aircraft, aircraft 

engine, or propeller specifications

Moment. A force that causes or tends to cause an object to 

rotate. It is indicated by the product of the weight of an item 

multiplied by its arm.

Moment. (GAMA) The product of the weight of an item 

multiplied by its arm. (Moment divided by a constant is used 

to simplify balance calculations by reducing the number of 

digits; see reduction factor.)

Moment index. The moment (weight times arm) divided by 

a reduction factor.

Moment limits versus weight envelope. An enclosed area on 

a graph of three parameters. The diagonal line, representing 

the moment/100, crosses the horizontal line, representing the 

weight at the vertical line, representing the CG location in 

inches aft of the datum. If the lines cross inside the envelope, 

the aircraft is loaded within its weight and CG limits.

Net weight. The weight of the aircraft minus the weight of 

chocks or other devices used to hold the aircraft on the scales.

Normal category. A category of aircraft certificat d under 

14 CFR part 23 and CAR part 3 that allows the maximum 

weight and CG range while restricting the maneuvers that 

are permitted.

PAX. The abbreviation of passengers.

Payload. (GAMA) The weight of occupants, cargo, and 

baggage.

Percent MAC. The distance in inches of the CG from 

LEMAC divided by the MAC. It is a good standard for CG 

location in airplanes because it permits a standard weight and 

balance program for different types of airplanes.

Pilot’s operating handbook (POH). An FAA-approved 

document published by the airframe manufacturer that lists 

the operating conditions for a particular model of aircraft 

and its engine(s).

Potable water. The water carried in an aircraft for the 

purpose of drinking.

Ramp weight. The zero fuel weight plus all of the usable 

fuel on board.

Reference datum. (GAMA) An imaginary vertical plane 

from which all horizontal distances are measured for balance 

purposes.

Reduction factor. A number, usually 100 or 1,000, by which 

a moment is divided to produce a smaller number that is less 

likely to cause mathematical errors when computing the 

center of gravity.

Residual fuel. The fuel that remains trapped in the system 

after draining the fuel from the aircraft with the aircraft in 

level flightattitude. The weight of this residual fuel is counted 

as part of the empty weight of the aircraft.

Service ceiling. The highest altitude at which an aircraft 

can maintain a steady rate of climb of 100 feet per minute.

Small aircraft (14 CFR part 1). An aircraft weighing 12,500 

pounds or less maximum certificated takeoff weight

Standard empty weight. (GAMA) The weight of a standard 

airplane including unusable fuel, full operating fluids, and 

full oil.
G-5

Static load. The load imposed on an aircraft structure due 

to the weight of the aircraft and its contents.

Station. (GAMA) A location along the airplane fuselage 

usually given in terms of distance from the reference datum.

Strain sensor. A device that converts a physical phenomenon 

into an electrical signal. Strain sensors in a wheel axle sense 

the magnitude of the force the axle deflects and create an 

electrical signal that is proportional to the force that caused 

the deflection

Structural station. A location in the aircraft, such as a 

bulkhead, that is identifi d by a number designating its 

distance in inches or percent MAC from the datum. The 

datum is, therefore, identified as station zero. The stations 

and arms are identical. An item located at station +50 would 

have an arm of 50 inches.

Takeoff weight. The weight of an aircraft just before 

beginning the takeoff roll. It is the ramp weight minus the 

weight of the fuel burned during start and taxi.

Tare weight. The weight of any chocks or devices that are 

used to hold an aircraft on the scales when it is weighed. The 

tare weight must be subtracted from the scale reading to get 

the net weight of the aircraft.

TEMAC. Trailing edge of the mean aerodynamic chord.

Type certificate data sheets (TCDS). The official 

specifications issued by the FAA for an aircraft, engine, or 

propeller.

Undrainable oil. The oil that does not drain from an engine 

lubricating system when the aircraft is in the normal ground 

attitude and the drain valve is left open. The weight of the 

undrainable oil is part of the empty weight of the aircraft.

Unusable fuel. (GAMA) The fuel remaining after a test has 

been completed in accordance with governmental regulations.

Usable fuel. (GAMA) The fuel available for fl ght planning.

Useful load. (GAMA) The difference between takeoff 

weight, or ramp weight if applicable, and basic empty weight.

Utility category. A category of aircraft certificated under 14 

CFR part 23 and CAR part 3 that permits limited acrobatic 

maneuvers but restricts the weight and the CG range.

Wing chord. A straight line distance across a wing from 

leading edge to trailing edge.

Zero fuel weight. The weight of an aircraft without fuel.
G-6
A

adverse-loaded CG checks.............................................7-6

aft adverse-load CG check.............................................7-7

aircraft category .............................................................4-3

aircraft certificatio ........................................................4-3

aircraft classificatio ......................................................4-3

airplanes.........................................................................4-5

airships...........................................................................4-5

airworthiness certificat ......................................... 4-4, 4-8

alteration ................................................................ 7-1, 7-5

amateur-built aircraft ...................................... 4-3, 4-6, 4-8

amateur-built LSA .........................................................4-1

arm ................................................................ 2-2, 7-5, 10-5

B

balance ...........................................................................1-4

control.........................................................................1-4

ballast ............................................................ 3-4, 7-7, 10-7

balloon............................................................. 1-5, 1-6, 4-5

basic empty weight ................................................ 5-2, 6-2

basic weight and balance equation.................................2-4

C

cargo.............................................................................9-18

center of gravity ............................. 2-2, 3-1, 5-1, 9-1, 10-1

change.........................................................................7-1

configuratio .............................................................9-18

determining....................................................... 3-5, 10-4

determining changes.................................................. 9-6

determining lateral CG of a helicopter.....................10-5

determining the loaded CG.........................................9-3

limits envelope ...........................................................5-4

range ...........................................................................2-6

range chart................................................................2-10

certificates and document .............................................4-8

chart method........................................................... 6-2, 6-3

Civil Aeronautical Administration.................................2-6

civil air regulations ................................................ 2-6, 3-4

commuter category.........................................................9-1

commuter category airplanes.......................................9-12

comprehensive equipment list......................................2-10

computer ......................................................................10-1

consensus standard........................................................ 4-3

cyclic control................................................................. 8-1

D

data pertinent to all model..............................................2-6

datum..............................................................................3-7

aft of the main wheels ................................................3-7

forward of the airplane ...............................................3-6

dedicated electronic flight compute ............................10-3

determining cargo pallet loads...................................... 9-9

determining new balance point ......................................2-3

determining weight ......................................................10-6

documenting changes.....................................................9-3

draining the fuel .............................................................3-4

E

E6-B flight compute ........................................... 10-1,10-2

effects of weight.............................................................1-3

electronic calculator.....................................................10-2

electronic load cells........................................................3-2

empty weight..................................................................9-2

empty weight center of gravity ........ 3-2, 6-2, 7-6, 8-3, 9-2

empty weight CG range .................................................2-6

equipment for weighing .................................................3-2

equipment list.................................................. 3-4, 5-2, 7-2

EWCG formulas.............................................................3-6

experimental airworthiness certificat ...........................4-6

extensive repair..............................................................3-2

F

FAA designee.................................................................4-4

finding balance poin ......................................................2-3

fixed-wing airplan ........................................................1-5

fleet operating empty weights (FOEW ........................ 9-5

floor loading limit .........................................................9-9

forward adverse-load CG check.....................................7-7

fuel capacity...................................................................2-6

fuel dump .....................................................................9-12

Index
I-2

G

gliders..................................................................... 4-1, 4-5

H

helicopter................................................................ 1-5, 8-1

I

initial weight ..................................................................9-2

J

jacking the aircraft .........................................................3-5

jacks ...............................................................................3-5

K

kit-built aircraft ..............................................................4-6

L

landing gear shock struts................................................3-5

landing weight..............................................................9-10

large aircraft...................................................................9-1

lateral balance ........................................................ 1-4, 8-2

lateral CG.......................................................................8-3

lateral level.....................................................................3-5

law of the lever....................................................... 2-1, 2-2

leading edge ...................................................................3-8

leading edge mean aerodynamic chord (LEMAC) 7-6, 9-2

leveling the aircraft ........................................................3-5

lift...................................................................................1-2

light aircraft....................................................................1-5

lighter-than-air aircraft........................................... 4-1, 4-5

light sport aircraft (LSA) .........................4-1, 4-2, 4-3, 4-6

light-sport category ................................................ 4-3, 4-4

light-sport eligible kit.....................................................4-3

load cell scales ...............................................................3-2

loaded CG ......................................................................6-2

of a helicopter.............................................................8-3

loaded weight.................................................................6-3

loading graph method ....................................................5-4

locating balance point. ........................................... 2-2, 2-3

location of CG................................................................3-8

longitudinal CG..............................................................8-3

M

major alteration ...................................................... 3-2, 7-3

manual computational method.......................................5-2

manufacturer-furnished information............................2-10

maximum amount of payload ......................................9-10

maximum baggage .........................................................2-6

maximum gross weight..................................................1-6

maximum weights..........................................................2-6

mean aerodynamic chord ................................ 2-5, 3-1, 3-8

modifying the cargo .......................................................9-6

moment ........................................................... 2-2, 5-4, 7-5

index ...................................................5-4, 7-5, 9-2, 10-5

limits...........................................................................6-3

multiengine aircraft........................................................6-1

N

nosewheel airplane with datum ahead of the main

wheels ..........................................................................10-4

nosewheel airplane with datum behind the main

wheels ..........................................................................10-4

nosewheel landing gear.......................................... 3-6, 3-7

number of seats..............................................................2-6

O

offloading carg .............................................................9-6

offloading passenger .....................................................8-3

oil ...................................................................................3-4

capacity.......................................................................2-6

onboard aircraft weighing system..................................9-5

operational empty weight (OEW)..................................9-5

other fluid .....................................................................3-5

overloading ....................................................................1-3

P

passengers....................................................................9-16

percent mean aerodynamic chord (MAC)...... 6-3, 7-6, 9-3, 

10-5

permanent ballast ...........................................................7-8

pilot in command ...........................................................3-2

pilot’s operating handbook ..........................................2-10

platform scales...............................................................3-2

powered parachute .............1-5, 1-6, 4-1, 4-3, 4-5, 4-6, 4-8

preflight plannin ...........................................................5-2

R

ramp wheel scales..........................................................3-2

reestablishing the OEW .................................................9-5

registration certificat .....................................................4-8

repair ............................................................... 7-1, 7-3, 7-5

S 

safety...................................................................... 3-2, 3-5

scale preparation ............................................................3-3

shifting cargo .................................................................9-8

shifting the balance point...............................................2-4

shifting weight ..................................................... 2-4, 10-6

single-engine aircraft .....................................................5-1

solution by chart.............................................................2-4
I-3

special airworthiness certificat ............................. 4-4, 4-6

special light-sport aircraft (SLSA).................................4-5

stability...........................................................................1-4

standard airworthiness certificat ...................................4-4

standard weights.............................................................3-4

statement of compliance ........................................ 4-3, 4-5

stress panels ...................................................................3-5

stress plates ....................................................................3-5

Swept-wing airplanes.....................................................1-5

T

tailwheel airplane with datum ahead of the main ........10-4

tailwheel airplane with datum behind the main ...........10-4

tailwheel landing gear....................................................3-7

temperatures change.......................................................3-4

temporary ballast............................................................7-8

formula .......................................................................7-8

total weight.....................................................................9-2

trailing edge ...................................................................3-8

trim setting .....................................................................9-6

twin-engine airplane.......................................................6-1

type certificate data sheet .......................2-6, 3-8, 5-2, 7-6

typical placard................................................................7-7

U

using fuel........................................................................8-3

W

weighing aircraft ............................................................3-2

weighing procedures......................................................9-2

weight.............................................................................1-2

changes............................................................... 1-3, 7-5

envelope......................................................................6-3

procedures ..................................................................3-3

weight and balance.......................................... 4-6, 4-8, 6-1

computations ..............................................................4-8

control................................................................. 1-2, 3-2

problems...................................................................10-4

records........................................................................3-2

record..........................................................................7-3

theory..........................................................................2-2

weight and loading sheet................................................4-8

weight and moment indexes...........................................6-3

weights.........................................................................10-5

weight-shift control (WSC).................................... 4-1, 4-6

weight-shift control aircraft ............................ 1-5, 4-5, 4-6

wheels ..........................................................................10-4
I-4