Maximum Landing Weight: The maximum weight approved for landing touchdown. Most multi-engine airplanes which operate over long stage lengths consume considerable weights of fuel. As a result, their weight is appreciably less on landing than at takeoff. Designers take advantage of this condition to stress the airplane for the lighter landing loads, thus saving structural weight. If the flight has been of short duration, fuel or payload may have to be jettisoned reduce the gross weight maximum or maximum landing weight.

Maximum Weight - Zero Fuel: Some transport planes carry fuel in their wings, the weight of which relieves; the bending moments imposed on the wings by the lift. The maximum weight - zero fuel limits the load which may be carried in the fuselage. Any increase in weight in the form of load carried fuselage must be counterbalanced by adding weight in the form of fuel in the wings.

Float Buoyancy: The maximum permissible gross weight of a seaplane is governed by the buoyancy of the floats. The buoyancy of a seaplane float is equal to the weight of water displaced by the immersed part of the float. This is equal to the weight the float will support without sinking beyond a predetermined level (draught line).

The buoyancy of a seaplane float is designated by its model number. A 4580 float has a buoyancy of 4580 lb. A seaplane fitted with a pair of 4580 floats has a buoyancy of 9160 lbs.

Regulations require an 80% reserve float buoyancy. The floats must, therefore, have a buoyancy equal to 180% of the weight of the airplane.

To find the maximum gross weight of a seaplane fitted with, say 7170 model floats, multiply the float buoyancy by 2 and divide by 1.8 (7170 x 2)/1.8 = 7966 lb.

C.  COMPUTING THE LOAD

A typical light airplane has a basic weight of 1008 lb. and an authorized maximum gross weight of 1600 lb. An acceptable loading of this airplane would be as follows:

Basic Empty Weight . . . . . . . . . . . . . .1008 lb.

Consisting of Weight Empty . . . . . . . . . . 973 lb.

Oil . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15 lb.

Extra Equipment . . . . . . . . . . . . . . . . . . . .20 lb

.

Useful Load . . . . . . . . . . . . . . .. . . . . . . . 592 lb.

Consisting of Pilot . . . . . . . . . . . . . . . . . . .150 lb.

Fuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 146 lb.

Payload: Passenger . . . . . . . . . . . . . . . . . .175 lb.

Baggage . . . .. . . . . . . . . . . . . . . . . . . . . . . 121 lb.

Problem

To find the maximum payload that can be transported a given distance and the amount of fuel required.

A seaplane on contract with a mining company is required to transport a maximum load of freight a distance of 300 nautical miles to a bush operation. The estimated groundspeed is 110 knots. The useful load for this airplane is 1836

pounds. Fuel capacity is 86 U.S. gallons. Fuel consumption is 20 gallons per hour or 120 lb of fuel per hour.

The time to fly 300 nautical miles is 164 minutes ((300/110) x 60). Add to that the 45 minutes required for reserve and the amount of fuel required must be sufficient for 209 minutes of flying time.

The amount of fuel required at 20 gallons per hour is 69.7 U.S. gallons ((20/60) x 209). That quantity of fuel weighs 418 lb (69.7 x 61b.).

The fuel calculations can also be computed by using the weight of fuel consumed per hour. The weight of fuel necessary for the flight is 418 lb. ((120/60) x 209).

The useful load is 1836 lb. The weight of the pilot (170 lb.) and fuel (418 lb.) is 588 lb. Therefore, the maximum payload permissible is 1248 lb.

What quantity of fuel in litres will be required? One U.S. gallon equals 3.785332 litres. The quantity of fuel required is, therefore, 263.8 litres (69.7 x 3.785332).

D.  BALANCE LIMITS

The position of the center of gravity along its longitudinal axis affects the stability of the airplane. There are forward and aft limits established by the aircraft design engineers beyond which the C.G. should not be located for flight. These limits are set to assure that sufficient elevator deflection is available for all phases of flight. If the C.G. is too far forward, the airplane will be nose heavy, if too far aft, tail heavy. An airplane whose center of gravity is too far aft may be dangerously unstable and will possess abnormal stall and spin characteristics. Recovery may be difficult if not impossible because the pilot is running out of elevator control. It is, therefore, the pilot's responsibility when loading an airplane to see that the C.G. lies within the recommended limits.

If the C.G. is too far forward, the airplane will be nose heavy, if too far aft, tail heavy. An airplane whose center of gravity is too far aft may be dangerously unstable and will possess abnormal stall and spin characteristics. Recovery may be difficult if not impossible because the pilot is running out of elevator control. It is, therefore, the pilot’s responsibility when loading an airplane to see that the C.G. lies within the recommended limits.

Usually the Airplane Owner's Manual lists a separate weight limitation for the baggage compartment in addition to the gross weight limitation of the whole airplane. This is a factor to which the pilot must pay close attention, for overloading the baggage compartment (even if the plane itself is not overloaded) may move the C.G. too far aft and affect longitudinal control.

The Airplane Owner's Manual may also specify such things as the seat to be occupied in solo flight (in a tandem seating arrangement) or which fuel tank is to be emptied first. Such instructions should be carefully complied with.

As the flight of the airplane progresses and fuel is consumed, the weight of the airplane decreases. Its distribution of weight also changes and hence the C.G. changes. The pilot must take into account this situation and calculate the weight and balance not only for the beginning of the flight but also for the end of it.

E.  DEFINITIONS

The center of gravity (C.G.) is the point through which the weights of all the various parts of an airplane pass. It is, in effect, the imaginary point from which the airplane could be suspended and remain balanced. The C.G. can move within certain limits without upsetting the balance of the airplane. The distance between the forward and aft C.G. limits is called the center of gravity range.

The balance datum line is a suitable line selected arbitrarily by the manufacturer from which horizontal distances are measured for balance purposes. It may be the nose of the airplane, the firewall or any other convenient point .

The moment arm is the horizontal distance in inches from the balance datum line to the C.G. The distance from the balance datum line to any item, such as a passenger, cargo, fuel tank, etc. is the arm of that item.

The balance moment of the airplane is determined by multiplying the weight of the airplane by the moment arm of the airplane. It is expressed in inch pounds. The balance moment of any item is the weight of that item multiplied by its distance from the balance datum line. It is, therefore, obvious that a heavy object loaded in a rearward position will have a much greater balance moment than the same object loaded in a position nearer to the balance datum line.

The moment index is the balance moment of any item or of the total airplane divided by a constant such as 100, 1000, or 10,000. It is used to simplify computations of weight and balance especially on large airplanes where heavy items and long arms result in large unmanageable numbers.

If loads are forward of the balance datum line their moment arms are usually considered negative (-). Loads behind the balance datum line are considered positive (+)*. The total balance moment is the algebraic sum of the balance moments of the airplane and each item composing the disposable load.

*In many cases the positive (+) sign is omitted, but the negative (-) sign is always shown. To simplify matters, both are included in our example

The C.G. is found by dividing the total balance moment (in inch-pounds) by the total weight (in lb.) and is expressed in inches forward (-) or aft (+) of the balance datum line.

The center of gravity range is usually expressed in inches from the balance datum line (i.e. +39.5" to +45.8"). In some airplanes, it may be expressed as a percentage of the mean aerodynamic chord (25% to 35%). The MAC is the mean aerodynamic chord of the wing.

To calculate the position of the C.G. in percent of MAC. Let us assume that the weight and balance calculations have found the C.G. to be 66 inches aft of the balance datum line and the leading edge of the MAC to be 55 inches aft of the same reference (Fig. 3). The C.G. will, therefore, lie 11 inches aft of the leading edge of the MAC. If the MAC is 40 inches in length, the position of the C.G. will be at a position (11 ~ 40) 27% of the MAC. If the calculated C.G. position is within the recommended range (for example, 25% to 35%), the airplane is properly loaded.