GENERAL

FLIGHT OPERATION
WEIGHT AND BALANCE

When the following terms are used in the Training manual for General Weight and Balance, they have
the following meanings:

1. Actual ( Operational ) Landing Weight ( LDW ). Is the maximum landing weight authorized at
touchdown by applicable government regulations when this weight is subject to varying degree
of destination airport limitations such as wind, elevation, temperature, runway length, etc. The
operational landing weight must never exceed the maximum design landing weight.

2. Actual ( Operational ) Take Off Weight ( TOW ). Is the maximum weight authorized at take off
brake release by applicable government regulations when this weight is subject to varying degree
of departure, destination airport limitations and en route conditions. It excludes taxi and run up
fuel, unless otherwise stipulated, and must never exceed the maximum design take off weight.

3. Actual Zero Fuel Weight ( ZFW ). Is the airplane basic operating weight plus payload and must
never exceed the MZFW.

4. Aircraft. Any weight-carrying device designed to be supported by the air.

5. Airplane Basic Empty Weight. Is the weight of the structure, power-plant, furnishings, un-usable
fuel, oil, chemical toilet fuel fluid, basic emergency equipment, oxygen system, galley structure,
electronic equipment required by operator, and fluid are contained in closed system.

6. Airplane. An engine-driven fixed wing aircraft, heavier than air, that is supported in flight by the
dynamic reaction of the air against its wing.

7. Body Station ( B.S. ) Numbers. Are the numbers which represent the number of INCHIES that the
particular location is aft of the datum line.

8. Center Of Gravity. Is the point in an aircraft around which all weight is evenly distributed or
balanced. The point of balance in an aircraft is the center of gravity.

9. Datum Line. Is an imaginary reference line from which all calculations or measurements are
taken for weight and balance purposes.

10. Dry ( Basic ) Operating Weight ( DOW ). Is the aircraft basic empty weight plus the following
operational items: engine tank oil, food and beverages equipment ( trolleys ), washing and
drinking water, emergency equipment, cargo equipment and related airborne equipment.

11. Empty Weight Center Of Gravity. Is the center of gravity of the airplane in an empty weight
condition.

12. Fulcrum. A point on which a lever is supported ( or about which it turns ).

13. Gross Weight. Is the weight of an airplane after all items have been added.

14. Basic Arm ( B.A. ). Is the horizontal distance from the CG of an object to the datum line.

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15. LEMAC. Is the abbreviation of Leading Edge Main Aerodynamic Chord, to indicate the leading
edge of the MAC.

16. MAC, Is the abbreviation of Mean Aerodynamic Chord.

17. Maximum ( Designed ) Zero Fuel Weight ( MZWF ). Is the maximum airplane weight less usable
fuel, and other consumable propulsion agents. It may include usable fuel in specified tanks when
carried in lieu of payload. The addition of usable and consumable items to the MZFW must be in
accordance with applicable government regulations so that the airplane structure and
airworthiness requirement are not exceeded.

18. Maximum ( Designed ) Landing Weight ( MLW ). Is the maximum landing weight authorized at
touchdown by applicable government regulations. This maximum landing gross weight is
structural limitation.

19. Maximum ( Designed ) Ramp ( Taxi ) Weight ( MRW ). Is the maximum weight authorized for
ground maneuver by the applicable structural limitations and includes taxi and run-up fuel.

20. Maximum ( Designed ) Take Off Weight ( MTOW ). Is the maximum weight authorized at take off
brake release by applicable government regulations and excluded taxi and run-up fuel. This
weight is limited by structural limitations of the airplane.

21. Moment. Is the tendency, or the measurement of the tendency, to produce rotation about point
or axis. Moment can be determined by multiplying the weight of a mass by its horizontal distance
from the center of gravity.
On some jet aircraft, datum line may be forward of the airplane’s nose.

22. Operating Center Of Gravity Range. Is the distance between the fore and aft CG limits.

23. Payload. Consist of the total weight of the passenger, baggage, cargo and mail.

24. TEMAC. Is the abbreviation of Trailing Edge Mean Aerodynamic Chord, to indicate the trailing
edge of MAC.

25. Unusable Fuel. Is the fuel remaining after a fuel run-out test has been completed in accordance
with applicable government regulations and is considered to be in two portions, drainable and
trapped. The drainable unusable fuel can only be drawn off from the sump jiffy drains.

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WEIGHT AND BALANCE

GENERAL

1. Weight.
Gravity is the pulling force that tends to draw all bodies to the center of the earth. The Center of
Gravity (CG) may be considered as a point at which all the weight of the airplane is concentrated. If
the airplane were supported at its exact center of gravity, it would balance in any attitude. It will be
noted that center of gravity is of major importance in an airplane, for its position has a great bearing
upon stability.
The location of the center of gravity is determined by the general design of each particular airplane.
The designers determine how far the Center of Pressure (CP) will travel. They then fix the center of
gravity forward of the center of pressure for the corresponding flight speed in order to provide an
adequate restoring moment to retain flight equilibrium.

2. Weight Control.
Weight is a major factor in airplane construction and operation, and it demands respect from all pilots
and particular diligence by all Aircraft & Power-plant ( A&P ) mechanics and repairmen.
Excessive weight reduces the efficiency of an aircraft and the safety margin available if an emergency
condition should arise.
When an aircraft is designed, it is made as light as the required structural strength will allow, 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 1,361Kg airplane must support 1,361 Kg 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 2.721 Kg.
Severe uncoordinated maneuvers or flight into turbulence can impose dynamic loads on the structure
great enough to cause failure. In accordance with of the Regulations , the structure of a normal
category airplane must be strong enough to sustain a load factor of 3.8 times its weight. That is, every
Kilogram of weight added to an aircraft requires that the structure be strong enough to support an
additional 3.8 Kg.
An aircraft operated in the utility category must sustain a load factor of 4.4, 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 the 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.

3. Effects of Weight.
Most modern aircraft are so designed that if all seats are occupied, all baggage allowed by the
baggage compartment is carried, and all of the fuel tanks are full, the aircraft will be grossly
overloaded. This type of design requires the pilot to give great consideration to the requirements of
the trip. 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.

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Some of the problems caused by overloading an aircraft are:

• the aircraft will need a higher takeoff speed, which results in a longer takeoff run;
• both the rate and angle of climb will be reduced;
• the service ceiling will be lowered;
• the cruising speed will be reduced;
• the cruising range will be shortened;
• maneuverability will be decreased;
• a longer landing roll will be required because the landing speed will be higher;
• excessive loads will be 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 the aircraft is loaded so the proposed flight can be safely made.

4. Fuel System.
General aspect of the aircraft fuel system such as fuel capacity, fuel control system as well as the
impact on CG position ( CG movement during refueling and CG movement in flight ).
The main part of the aircraft weight and most particularly the payload is located in the fuselage. In
flight, this weight is balanced by the lift created mainly on the wings. This distribution generates a
bending moment around the wing root. This has strong impact on the aircraft structure and leads to
define a Maximum Zero Fuel Weight ( MZFW ) in order to limit the stress at these locations.

Lift Lift
2 2

Important
structural
strain.
Weight
= mg

Figure 1. Zero Fuel Weight

On the one hand, the weight of fuel tanked in the wings balances the effect of the lift and reduces the
bending moment. On the other hand, the fuselage is an extra load the wing have to create lift for. So
fuel has to be kept in the wing as long as possible.
This is the reason why the wing tanks are the first tanks to be filled and the last one to be emptied and
this no matter the aircraft type.
Moreover, as the wing tanks are generally divided into outer and inner wing tanks, the outer tank is
filled first in order to bring the fuel vector further out. However, it is worth pointing out that this rule
is only applicable to certain extent: too much weight in the outer tank creates some structural strain
on the wing on the ground ( when not balanced by the lift ).

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5. Longitudinal Stability.
Longitudinal stability about the lateral axis is considered to be the most affected by certain variables
in various flight conditions.
Longitudinal stability is the quality that makes an aircraft stable about its lateral axis. It involves the
pitching motion as the aircraft’s nose moves up and down in flight. A longitudinally unstable aircraft
has a tendency to dive or climb progressively into a very steep dive or climb, or even a stall. Thus, an
aircraft with longitudinal instability becomes difficult and sometimes dangerous to fly. Static
longitudinal stability or instability in an aircraft, is dependent upon three factors:
1. Location of the wing with respect to the CG;
2. Location of the horizontal tail surfaces with respect to the CG;
3. Area or size of the tail surfaces ( horizontal stabilizer ).

In analyzing stability, it should be recalled that a body free to rotate always turns about its CG.
To obtain static longitudinal stability, the relation of the wing and tail moments must be such that, if
the moments are initially balanced and the aircraft is suddenly nose up, the wing moments and tail
moments change so that the sum of their forces provides an unbalanced but restoring moment
which, in turn, brings the nose down again. Similarly, if the aircraft is nose down, the resulting change
in moments brings the nose back up.

W
Figure 2. Longitudinal Stability.

The CP in most asymmetrical airfoils has a tendency to change its fore and aft positions with a change
in the Angle of Attatck ( AoA ). The CP tends to move forward with an increase in AoA and to move aft
with a decrease in AoA. This means that when the AoA of an airfoil is increased, the CP, by moving
forward, tends to lift the leading edge of the wing still more. This tendency gives the wing an inherent
quality of instability.
Most aircraft are designed so that the wing’s CP is to the rear of the CG. This makes the aircraft “nose
heavy” and requires that there be a slight downward force on the horizontal stabilizer in order to
balance the aircraft and keep the nose from continually pitching downward.

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T
W
L
CP

CG T
W
Figure 3. Nose down.

Compensation for this nose heaviness is provided by setting the horizontal stabilizer at a slight
negative AoA. The downward force thus produced holds the tail down, counterbalancing the “heavy”
nose. Itis as if the line CG-CL-T were a lever with an upward force at CL and two downward forces
balancing each other, one a strong force at the CG point and the other, a much lesser force, at point T
(downward air pressure on the stabilizer).

6. 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 can
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. With some helicopters,
they 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 will be 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 will
be noted by a placard, usually on the instrument panel, and they should be strictly adhered to.
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As an aircraft ages, its weight usually increases due to trash and dirt collecting in hard-to-reach
locations, and moisture absorbed in the cabin insulation. This growth in
weight is normally small, but it can only be determined 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, will probably cause the CG to shift and this must be computed and
annotated in the weight and balance record.
Repairs and alteration are the major sources of weight changes, and it is the responsibility of the A&P
mechanic or repairman making any repair or alteration to know the weight and location of these
changes, and to compute the CG and record the new empty weight and EWCG in the aircraft weight
and balance record.
If the newly calculated EWCG should happen to fall outside the EWCG range, it will be necessary to
perform adverse loading check. This will require 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. Adverse loading checks are a deliberate attempt to
load an aircraft in a manner that will create the most critical balance condition and still remain within
the design CG limits of the aircraft. If any of the checks fall outside the loaded

CG range, the aircraft must be reconfigured or placarded to prevent the pilot from loading the aircraft
improperly. It is sometimes possible to install fixed ballast in order for the aircraft to again operate
within the normal CG range.
The A&P 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 pilot in command to use the most current weight and balance data when
operating the aircraft.

7. 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 determines safety in flight.
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 so it will maintain straight and
level flight with hands off the controls. Longitudinal stability is maintained by ensuring the CG is
slightly ahead of the CP.
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.

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In sufficient
elevator to nose
up force

Figure 4. CG too far forward.

Longitudinal forces acting on an airplane in flight.

If a rising air current should cause the nose to pitch up, the airplane will slow down and the
downward force on the tail will decrease. The weight concentrated at the CG will pull the nose back
down. If the nose should drop in flight, the airspeed will increase and the increased downward tail
load will bring the nose back up to level flight.
As long as the CG is maintained within the allowable limits for its weight, the airplane will have
adequate longitudinal stability and control. If the CG is too far aft, it will be too near the center of lift
and the airplane will be unstable, and difficult to recover from a stall.

In sufficient
W elevator to nose
down force

Figure 5. CG too far aft

If the unstable airplane should ever enter a spin, the spin could become flat and recovery would be
difficult or impossible.
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.
If the CG is too far forward, the downward tail load will have to be increased to maintain level flight.
This increased tail load has the same effect as carrying additional weight; the aircraft will have to fly at
a higher angle of attack, and drag will increase.

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A more serious problem caused by the CG being too far forward is the lack of sufficient elevator
authority. At slow takeoff speeds, the elevator might not produce enough nose-up force to rotate and
on landing there may not be enough elevator force to flare the airplane. Both takeoff and landing
runs will be lengthened if the CG is too far forward.
If the CG is too far forward, there will not be enough elevator nose-up force to flare the airplane for
landing.
The basic aircraft design assumes that lateral symmetry exists. For each item of weight added to the
left of the centerline 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 burn-off. The position of the lateral CG is
not normally computed for an airplane, but the pilot must be aware
of the adverse effects that will result from a laterally unbalanced condition.
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 inefficiently.
Lateral imbalance causes wing heaviness, which may be corrected by deflecting the aileron. The
additional lift causes additional drag and the airplane flies inefficiently.
Helicopters are affected by lateral imbalance more 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 to require large lateral displacement
of the cyclic control to maintain level flight, the fore-and-aft cyclic control effectiveness will be 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, and as it is used from the inboard tanks, the CG shifts aft. For
this reason, fuel-use scheduling in swept-wing airplanes operation is critical.

Outboard fuel
tail heavy

Inboard fuel
nose heavy

Figure 6. Fuel in tank swept-wing

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.

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Weight and Balance Theory.

1. Element in the Weight and Balance.

• The total weight of the aircraft must be no greater than the maximum weight allowed for the
particular make and model of the aircraft.
• The center of gravity, 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.

2. Aircraft Arms, Weight, and Moments.

The term arm, usually measured in inches, refers to the distance between the center of gravity of an
item or object and the datum. Arms ahead of, or to the left of the datum are negative(-), and those
behind, or to the right of the datum are positive(+). When the datum is ahead of the aircraft, all of the
arms are positive and computational errors are minimized. Weight is normally measured in pounds.
When weight is removed from an aircraft, it is negative(-), and when added, it is positive (+).
The manufacturer establishes the maximum weight and range allowed for the CG, as measured in
inches from the reference plane called the datum. Some manufacturers
specify this range as measured in percentage of the Mean Aerodynamic Chord (MAC), the leading
edge of which is located a specified distance from the datum.
The datum may be located anywhere the manufacturer chooses; it is often the leading edge of the
wing or some specific distance from an easily identified location. One popular location for the datum
is a specified distance forward of the aircraft, measured in inches from some point, such as the nose
of the aircraft, or the leading edge of the wing, or the engine firewall.
The datum of some helicopters is the center of the rotor mast, but this location causes some arms to
be positive and others negative. To simplify weight and balance computations, most modern
helicopters, like airplanes, have the datum located at the nose of the aircraft or a specified distance
ahead of it.
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 Kilogram-Meters (Kg-M) and may be either positive or negative.

WEIGHT ARM MOMENT ROTATION

+ + + Nose Up

+ - - Nose Down

- + - Nose Down

- - + Nose Up

Figure 7. Relationships between the algebraic signs of weight, arms, and

moments.

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3. The Law of the Lever.

The 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 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
algebraic sum of the moments about the fulcrum is zero.
50 Cm 25 Cm

A=50 KG B= 100 KG

Fulcrum
CG
Moment +
Moment -

Figure 8. The lever is balanced when the algebraic sum of the moments 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 counter-clockwise).

ITEM WEIGHT ( Kg) ARM (Cm) MOMENT

(Kg-Cm)

A 50 - 50 - 2500

B 100 + 25 + 2500

Total 150 - 0

Figure 9. When a lever is in balance, the sum of the moments is zero.

4. Determining the CG.

One of the easiest ways to understand weight and balance is to consider a board with weights placed
at various locations. We can determine the CG of the board and observe the way the CG changes as
the weights are moved.
The CG of a board like the one in may be determined by using these four steps:
1. Measure the arm of each weight in Meter from the datum.
2. Multiply each arm by its weight in Kg to determine the moment in Kg-M of each weight.
3. Determine the total of all weights and of all the moments. Disregard the weight of the board.
4. Divide the total moment by the total weight to determine the CG in M from the datum.

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90
45
20

A=50 KG A=200 KG A=100 KG

54.3
CG

ITEM WEIGHT ( Kg) ARM (Cm) MOMENT

(Kg-Cm)

A 50 20 1000

B 200 45 9000

C 100 90 9000

Total 350 19000

Figure 10. Determining CG position.

Determine the CG by dividing the total moment by the total weight.:

19 000
= 54.3 Kg/Cm
350

To prove this is the correct CG, move the datum to a location 54.3 to the right of the original datum
and determine the arm of each weight from this new datum. If the CG is correct, the sum of the
moments will be zero.
All of the measurements must be made from the same datum location.

5. Location with Respect to the Mean Aerodynamic Chord.

The aircraft mechanic or repairman is primarily concerned with the location of the CG relative to the
datum, an identifiable physical location from which measurements can be made. But because the
physical chord of a wing that does not have a strictly rectangular plan form is difficult to measure,
wings such as tapered wings express the allowable CG range in percentage of mean aerodynamic
chord (MAC). The allowable CG range is expressed in percentages of the MAC.

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Datum

CG LEMAC

MAC

Figure 11. The MAC is the chord drawn through the geographic center of the plan
area of the wing.

The MAC, as seen in Figure 11, is the chord of an imaginary airfoil that has all of the aerodynamic
characteristics of the actual airfoil. It can also be thought of as the chord drawn through the
geographic center of the plan area of the wing.

Wing Area
MAC =
Wing Span

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.

LEMAC TEMAC

MAC = 100%

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Consequently, relating the CG location to the chord of the wing is convenient from a design and
operations standpoint. Normally, an aircraft will have acceptable flight characteristics if the CG is
located somewhere near the 25 percent average chord point. This means the CG is located one-fourth
of the total distance back from the leading edge of the wing section. Such a location will place the CG
forward of the aerodynamic center for most airfoils.
It is sometimes necessary to determine the location of the CG in Metres from the datum when its
location in %MAC is known.
Determine the location of the CG in inches from the datum by using this formula:
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 a percentage of the MAC, the location of the CG is expressed in
the same terms.

6. Body Station ( BS ).
BS is item CG position.

a. Individual Adjustment.
For Item CG position relatively fix (permanent) e.g. passenger chair position.

Datum

Figure 13. Determining BS Individual Adjustment.

b. Area Adjustment.
For CG position item of baggage, cargo especially in bulk system.

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Datum

BS BS BS BS
Aft Forward Aft Forward
Arm

Centroid 1
Arm

Centroid 2
Figure 14. BS Area Adjustment.

BS Aft – BS Forward
Centroid ( BS )=
2
Moment Forward Cargo = Weight x Arm Centroid 1;

Moment Aft cargo Compartment = Weight x Arm Centroid 2.

7. Shifting the CG.

One common weight and balance problem involves moving passengers from one seat to another or
shifting baggage or cargo from one compartment to another to move the CG to a desired location.
This also can be visualized by using a board with three weights and then working out the problem the
way it is actually done on an airplane.

CG =
Weight Shifted x Distance it is shifted

Total Weight

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8. Ballast.
It is possible to load most modern airplanes so the center of gravity shifts outside of the allowable
limit. Placards and loading instructions in the Weight and Balance Data
inform the pilot of the restrictions that will prevent such a shift from occurring. A typical placard in
the baggage compartment of an airplane might read:
When rear row of seats is occupied, a certain pounds of baggage or ballast must be carried in forward
baggage compartment. For additional loading instructions, see Weight and Balance Data.
When the CG of an aircraft falls outside of the limits, it can usually be brought back in by using ballast.

8.1. 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.
To determine the amount of temporary ballast needed, use this formula:

8.2. 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.
Load conditions for aft adverse-loaded CG check.
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.
This block should be painted red and marked “Permanent Ballast - Do Not Remove.”

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WEIGHT AND BALANCE

1. Light Aircraft Operations.

Weight and balance data allows the pilot to determine the loaded weight of the aircraft and
determine whether or not the loaded CG is within the allowable range for the weight.
An important part of preflight planning is to determine that the aircraft is loaded so its weight and CG
location are within the allowable limits. There are two ways of doing this: by the computational
method using weight, arms, and moments; and by the loading graph method, using weight and
moment indexes.
The 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 POH/AFM.
A worksheet such as the one in Figure 15 provides space for all of the pertinent weight, CG, and
moment along with the arms of the seats, fuel, and baggage areas.

ITEM WEIGHT ARM MOMENT CG

Airplane
Front seat
Rear seat
Fuel
Baggage A
Baggage B

Figure 15. Weight and Balance worksheet.

Determine the moment of each item by multiplying its weight by its arm. 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.
To determine that the airplane is properly loaded for this
flight, use the CG limits envelope in Figure 16 (which is typical of those found in the POH/AFM). Draw
a line vertically upward from the CG, and one horizontally to the right from the loaded weight . These
lines cross inside the envelope, which shows the airplane is properly loaded for takeo

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Figure 16. Center of

Gravity limit.

1.1. Loading Graph Method.

Everything possible is done to make flying safe, and one expedient method is the use of charts and
graphs from the POH/AFM to simplify and speed up the preflight weight
and balance computation. Some use a loading graph and moment indexes rather than the arms and
moments. These charts eliminate the need for calculating the moments and thus make computations
quicker and easier. [Figure 16].
Figure 17 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.

1.2. Moment Indexes.

Moments determined by multiplying the weight of each component by its arm result in large numbers
that are awkward to handle and can become a source of mathematical error. To eliminate these large
numbers, moment indexes are 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, so you can avoid
mathematical calculations. The CG envelope uses moment indexes rather than arms and moments.

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.

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Figure 17. Loading Graph.

From the figure 16 draw a line vertically upward from the value total moment on the horizontal index
at the bottom of the chart, and a horizontal line from the total weight in the left-hand vertical index.
These lines intersect within the dashed area, which shows that the aircraft is loaded properly for
takeoff.

2. Large Aircraft Operations.

2.1. Determining Empty Weight.

The empty weight and EWCG are determined by using the following steps, and the results are
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 main-wheel points by multiplying the net weight (scale
reading less 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 nose wheel 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 this by the reduction factor. This
gives the CG in inches, from the datum.

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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.
Distance CG to LEMAC = Datum to CG – Datum to LEMAC.
6. Determine the EWCG in % MAC by using this formula:

EWCG in % MAC = ( CG distance from LEMAC ) : MAC x 100% =

2.2. Determining the Loaded CG of the Airplane in Percent MAC.

The basic operating weight (BOW) and the operating index are entered into a loading schedule like
the one in Figure 18 and the variables for the specific flight are entered as are appropriate to
determine the loaded weight and CG.

Use the data in this example:

Basic operating Weight........................... 105,500 lbs.
Basic operating index (total moment/1,000) 98,837.0....
MAC.............................................................. 180.9 in
LEMAC............................................................. 860.5

ITEM WEIGHT ( Lbs ) MOMENT / 1000

BOW 105 500 92 387
Pax forward 18 3 060 1 781
Pax aft 95 16 150 16 602
Forward cargo 1 500 1 020
Aft cargo 2 500 2 915
Fuel tank 1 10 500 10 451
Fuel tank 2 10 500 10 451
Fuel tank 3 28 000 25 589
Total 177 710 161 646

CG position = ( 161 646 : 177 710 ) x 1000 = 909.6 Inch

CG MAC % = {( 909.6 – 860.5 ) : 180.9} x 100 = 27.1%.

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Figure 18. Loading schedule.

2.3. 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 % MAC, and in Units ANU (Airplane Nose Up).

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2.4. Stabilizer Trim Setting in Percent of MAC.

If the stabilizer trim is calibrated in units of % MAC, determine the CG location in % MAC as has just
been described, then set the stabilizer trim on the percentage figure thus determined.

2.5. Stabilizer Trim Setting in Percent of ANU (Airplane Nose Up)

Some aircraft give the stabilizer trim setting in Units ANU that correspond with the location of the CG
in % MAC.
When preparing for takeoff in an aircraft equipped with this system, first determine the CG in % MAC
in the way described above, then refer to the Stabilizer Trim Setting Chart on the Takeoff Performance
page of the AFM.

Figure 19. Stabilizer trim setting.

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 GENERAL
FLIGHT OPERATION
WEIGHT AND BALANCE

1. Structural Limitation and Floor Panel Limitation.

1.1. Structural Limitation.

1.1.1. Running ( Linear ) Load Limitation.

The Running Load Limitation is the maximum load acceptable on any given fuselage length of an
aircraft floor.

Weight of the piece ( W )

Running Load =
Length of the piece in the flight direction ( L )

Unit : Kg / M or Lb / Ft.

Example 1.

Assume a Maximum Running Load of 2000 Kg/M

870 KG Flight
Direction

1.2 M

Figure 20. Running Load wide contact area.

Running load = 870 : 2 = 725 Kg/M < 2000 Kg/M.

Maximum weight = 2000 x 1.2 = 2400 Kg > 870.
Running Load not exceeded.

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Flight
870 KG Direction

0.2 M 0.2 M
Figure 21. Running Load narrow contact area.

Running Load = 870 : ( 2 x 0.2 ) = 2175 Kg/M > 2000 Kg/M.

Maximum Weight = 2000 x ( 2x 0.2 ) = 800 Kg < 870 Kg.
Running Load Exceeded.

1.1.2. Area Load Limitation.

The Area Load Limitation is the maximum load acceptable on any surface unit of an aircraft floor.
It prevents the load from exceeding the capability of the aircraft structure ( floor beam, floor post,
floor panels and frames ).

Weight of the piece ( W )

Area Load =
Contour Area ( S )

2 2.
Unit : Kg / M or Lb / Ft .

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2
Assume a Maximum Area Load of 2000 Kg/M .

0.1 M

4 Contact
Point

870 KG 0.5 M

0.1 M

1.2 M

Figure 22. Area Load Limitation

2 2
Area Load = 870 : ( 1.2 x 0.5 ) = 1450 Kg/M < 2000 Kg/M .
Maximum Weight = 2000 x ( 1.2 x 0.5 ) = 1200 Kg > 870 Kg.
Area Load not exceeded.

1.1.3. Compartment Load Limitation.

The Compartment Load Limitation is the maximum load acceptable in an entire compartment.
This limitation applies to the whole load located in a given compartment.

Unit : Kg or Lb.

1.1.4. Cumulative Load Limitation.

The Cumulative Load Limitation is the maximum weight can be carried forward or aft of a given
section.
This limitation prevents the weight loaded in the forward and aft fuselage section is exceeded the
capability of the frames and skin stringers.
Unit : Kg or Lb.

1.2. Floor Panel Limitation.

1.2.1. Contact Load Limitation.

The Contact Load Limitation is the maximum load acceptable in direct contact with aircraft floor per
surface unit.
This limitation is used to prevent the load in direct contact with the floor from exceeding the
capability of the horizontal floor panels ( metal sheet, honey comb and sandwich panels ).

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2 2
Unit : KG / M or Lb / Ft .

Weight of the Piece ( W )

Contact Load =
Contact Area ( S )
2
Let’s assume a Maximum Contact Load of 2000 Kg/M .

0.2 M 0.2 M

870 KG 0.5 M

1.2 M

Surface in direct
contact
2
with the floor ( 0.2 M ).

Figure 23. Contact Load Limitation.

2 2
Contact Load = 870 : 0.2 = 4350 Kg/M > 2000 Kg Kg/M .

Maximum Weight = 2000 x 0.2 = 400 Kg < 870 Kg.

1.2.2. Point Load Limitation.

The Point Load Limitation represent the resistance to puncture by a heavy load bearing onto a very
small surface of the floor panels ( Walking Load ).
2 2
Unit : Kg /M or Lb / Ft .

Never lay a heavy package on an edge or corner.

Figure 24. Point Load Limitation.

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When using a pinch bar, place a floor protector device beneath the pinch bar ( plank, piece of wood ).

Figure 25. Using wrong position pin bar.

Figure 26. Correct using pin bar.

1.2.3. Spreader Floor.

Spreader floor enables to transport load whose weight exceed on of the previous limit ( running load,
area load or contact load ).
It allows increasing either of the piece or the area in direct contact with floor.

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Example:
Maximum Running Load = 500 Kg/M.
2
Maximum Area Load = 1000 Kg/M .
2
Maximum Contact Load Area = 2000 Kg/M .

500 Kg 0.5 M
0.2 M

Spreader weight = 20 Kg. 0.2 M

1.2 M

Figure 27. Spreader.

.
Without Spreader Floor:

Running Load = 500 ; 0.4 = 1250 Kg/M;

2
Area Load = 500 : 0.35 = 1429 Kg/M ;
2
Contact Load = 500 : 0.16 = 3125 Kg/M .

With Spreader Floor:

Running Load = ( 500 +20 ) ; 1.2 = 434 Kg/M;

2
Area Load = ( 500 + 20 ) : 0.6 = 867 Kg/M ;
2
Contact Load = ( 500 + 20 ) : 0.48 = 1084 Kg/M .
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BIBLIOGRAPHY

1. FAA – Aircraft Weight and Balance Handbook - 2007.

2. CAA – Safety Sense Leaflet 9, Weight and Balance Leaf Let – 2005.

3. AIRBUS – Weight And Balance Manual – 2004.

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