Aircraft weight and balance system with automatic loading error correction

A weight and balance system which automatically compensates for loading errors while the aircraft is stationary. A weight signal is generated from the outputs of a pair of inclinometers mounted on adjacent opposite ends of a landing gear member which bends in response to aircraft weight. The weight signal generated just after no movement of the aircraft is detected is held in a latch. The held weight signal and a subsequently generated weight signal are combined to provide a signal representative of the difference therebetween. The difference signal is scaled by an error correction coefficient to provide a loading error compensation signal which is combined with the subsequently generated weight signal to provide a corrected weight signal. The system further includes means for performing reasonableness tests in response to the restoration of power to the system after power has been down. The reasonableness tests include determining whether the aircraft had been moved or whether there had been a load change during the power down. If movement of the aircraft is determined to have occurred, the latch is reset to hold the weight signal generated just after power to the system is restored. If a load change is determined to have occurred, a power down load correction signal is generated and combined with the loading error compensation signal and the subsequently generated weight signal to provide the corrected weight signal.

BACKGROUND OF THE INVENTION 
The present invention relates to an aircraft weight and balance system and 
more particularly to a weight and balance system which automatically 
compensates for loading errors while the aircraft is stationary. 
Before each flight, the weight and balance of an aircraft must be 
determined to insure that they are within safe operating limits. This is 
typically accomplished by estimating the total aircraft weight when loaded 
and the distribution of that weight in order to determine the aircraft's 
center of gravity or balance. 
To insure proper distribution of the weight, onboard weight and balance 
systems for aircrafts have been developed which calculate the weight and 
center of gravity of the aircraft in response to various sensor signals. 
Known weight and balance systems using strain gauges and pressure or 
magnetic variable reluctance sensors have been found to be unreliable due 
to problems in the sensor's stability, accuracy, and ability to survive 
harsh environments. 
A known weight and balance system which overcomes the above problems is 
shown in Bateman, U.S. Pat. No. 4,312,042. This system uses inclinometers 
positioned to measure the angle of bending in a structural member, such as 
the landing gear of the aircraft, where the angle of bending is 
proportional to the weight or force on the member. The inclinometers 
provide output signals which are summed to cancel out attitude and 
acceleration changes, whereby only that portion of the angle of bending 
corresponding to weight is obtained. When the aircraft is stationary and 
being loaded and unloaded, it has been found that moments are created 
which constrain the bending of the structural member on which the 
inclinometers are mounted so that an error results in the calculated 
weight derived from the inclinometer outputs. Although the moments 
constraining the bending of the landing gear members are virtually 
eliminated once the aircraft is moved so that the members deflect to their 
true unconstrained value resulting in an accurate weight determination, it 
is impractial to require that the aircraft be moved after loading in order 
to determine whether its weight and balance are within the safety limits. 
SUMMARY OF THE INVENTION 
In accordance with the present invention, the disadvantages of prior weight 
and balance systems for aircraft have been overcome. 
The weight and balance system of the present invention accurately 
calculates weight while the aircraft is stationary, the system including 
means for automatically correcting loading errors in weight signals 
representing the sensed aircraft weight on a landing gear, weight signals 
being generated from the outputs of sensors which may be inclinometers, 
strain gage sensors, magnetic reluctance transducers or the like. 
The system for automatically correcting loading errors holds the weight 
signal, generated just after the aircraft stops moving in a latch. While 
the aircraft is stationary, the held weight signal is combined with a 
weight signal generated subsequently thereto to provide a signal 
representative of the difference therebetween. The difference signal is 
multiplied by an error correction coefficient associated with the 
particular landing gear the weight on which is being determined, to 
provide an error compensation signal. The error compensation signal is 
then combined with the subsequently generated weight signal to provide a 
corrected weight signal. 
The system further includes means for making reasonableness tests on 
individual landing gears in response to the restoration of power after 
power to the system has been down to determine whether the landing gear 
had moved or whether there had been a load change during the power down. 
In the event that a landing gear member is determined to have moved during 
the power down, the latch is reset to hold the uncorrected weight signal 
generated immediately after power has been restored to the system. In the 
event that the load on a landing gear member is determined to have changed 
during the power down, a power down correction factor is generated and 
combined with the uncorrected weight signal and the error compensation 
signal to provide the corrected weight signal, the power down correction 
factor compensating for the larger loading error which would otherwise 
result. 
The system for correcting loading errors further includes means for 
detecting movement of the aircraft in response to the difference in the 
outputs from a pair of inclinometers mounted on adjacent opposite ends of 
the landing gear member, the difference signal being proportional to 
acceleration independent of the weight on the member. The acceleration 
signal is applied to a third order band pass filter which acts as a pseudo 
double integrator, the output of which represents movement of the landing 
gear in distance. The output of the filter is applied to a comparator 
which determines whether the landing gear is moving or stationary.

The weight and balance system of the present invention is used to determine 
the weight and center of gravity of an aircraft such as illustrated in 
FIG. 1, while the aircraft is stationary on a runway or loading ramp. The 
aircraft 10 has a fuselage 12, with a pair of wings, the right wing being 
shown at 14, and mounting a jet engine 12. In the illustrated embodiment, 
the aircraft 10 has main landing gear, one of which is shown at 18, and a 
nose gear 20. 
The weight and balance system, shown in FIG. 2, determines the weight of 
the aircraft 10 by measuring the weight on each of the main landing gears 
18 and 22 and the weight on the nose gear 20. The weight on each of the 
landing gears is determined by sensing the amount of deflection or bend in 
the bogie beams 24 and 26 of the respective main landing gears 18 and 22 
and in the axle 28 of the nose gear 20, the angle of bending being 
proportional to the weight or force on each of the members as disclosed in 
U.S. Pat. No. 4,312,042 and incorporated herein by reference. 
The angle of bending is measured by a pair of inclinometers mounted on 
adjacent opposite ends of each of the landing gear members such as the 
inclinometers 30 and 32 for the bogie beam 24 of the main landing gear 18, 
the inclinometers 34 and 36 for the bogie beam 26 of the main landing gear 
22 and the inclinometers 38 and 40 for the nose gear axle 28. The 
inclinometers 30-40 may be servoed accelerometers as disclosed in U.S. 
Pat. No. 3,702,073, the disclosure of which being incorporated herein by 
reference, with each accelerometer having its sensitive axis aligned with 
the axis of the bogie beam or axle with no load being applied thereto. 
Each of the inclinometers or servoed-accelerometers provides an indication 
of the actual angle of the beam or axle relative to a plane which is 
perpendicular to the force of gravity. 
The outputs of the inclinometers 30 and 32 represent the respective angles 
.theta..sub.1 and .theta..sub.2, measured with respect to an inertial 
reference, the angles being defined as follows: 
EQU .theta..sub.1 =.theta..sub.B +.theta..sub.L1 +.theta..sub.A1 (1) 
EQU .theta..sub.2 =-.theta..sub.B +.theta..sub.L2 +.theta..sub.A2 (2) 
where .theta..sub.B is the angle of the beam or axle caused by the airport 
ramp or runway tilt; .theta..sub.L1 and .theta..sub.L2 are the beam bend 
angles caused by a load; and .theta..sub.A1 and .theta..sub.A2 are the 
sensor axis misalignment and bias terms. The weight on a given landing 
gear member, bogie beams 24, 26 or nose gear axle 28 is given by the 
following equation: 
EQU W.sub.T =K(.theta..sub.L1 +.theta..sub.L2) (3) 
Where K is a scale factor which is dependent upon the beam or axle geometry 
and strength. 
From the equations 1, 2 and 3, it is seen that by summing the output 
signals from each of the inclinometers such as 30 and 32 representing the 
angles .theta..sub.1 and .theta..sub.2, a signal is provided which is 
proportional to the weight on the beam 24, independent of the airport ramp 
angle or runway tilt angle, .theta..sub.B. Similarly, by subtracting the 
output signals of the inclinometers, a signal is provided which is 
proportional to the acceleration of the member, independent of the weight 
applied thereto. 
Each of the main landing gear members 18 and 22 are provided with two pairs 
of inclinometers, the pairs 30, 32 and 42, 44 for the bogie beam 24 and 
the pairs 34, 36 and 46, 48 for the bogie beam 26. The output signals of 
the inclinometers 30 and 32, are summed by a voltage summing circuit 50, 
the output of which on line 52 is proportional to the weight on the bogie 
beam 24. Similarly, the output signals of the inclinometers 34 and 36 are 
summed by a voltage summing circuit 54, the output of which on line 56 is 
proportional to the weight applied to the bogie beam 26. The output 
signals of the inclinometers 42 and 44 are applied to a differencing 
circuit 58, the output of which on line 60 is proportional to acceleration 
independent of weight. Similarly, the outputs of the inclinometers 46 and 
48 are applied to a differencing circuit 62, the output of which on line 
64 is proportional to acceleration independent of weight. The 
inclinometers 38 and 40 associated with the nose gear axle 28 have their 
respective signal outputs on lines 66 and 68 applied to a means 70 which 
performs a summing operation, the signal output on a line 72 representing 
the weight on the nose gear axle. 
The weight signals generated on lines 52, 56 and 72 and the acceleration 
signals on lines 60 and 64 are applied to a computer 74 which may be a 
digital computer or an analog computer. The computer 74 calculates the 
weight on each of the landing gears according to equation 3 above and from 
those weights the computer calculates the total weight of the aircraft, 
the calculated aircraft weight being displayed on a pilot display unit 76. 
The center of gravity may also be determined directly from the geometry of 
the aircraft and the weight on each of the landing gears using well-known 
formulas, the calculated center of gravity also being displayed on the 
pilot display unit 76. The computer 74 interacts with a Calibration Data 
Module (CDM) 80 which is a non-volatile memory for retaining key weight 
data in the event that power to the aircraft is removed. 
The forces acting on the nose gear 20 and the main landing gears, such as 
the gear 18, are illustrated in FIGS. 3A and 3B respectively for a 
stationary aircraft. When the bogie beam 24 or the axle 28 is loaded and 
unloaded with weight, the resultant bending of the beam or axle creates a 
force tending to move the tires 80, 82 on the runway or ramp surface 84. 
This force, however, is opposed by frictional static forces between the 
tires and the ramp, creating counter-moments in the bogie beam and axle 
which constrain the bending of these members, the loading effects for the 
nose gear 20, where the axle is at right angles to the track of the tires 
being larger than the effects for the bogie beam 18 where the track of the 
tires and the beam are in the same axis. Because the weight and balance 
system calculates the aircraft weight from the angle of bending of these 
structural members, a loading error, if uncompensated for, is introduced 
into the calculated weight, due to the constrained bending of the beams 
and axle when the aircraft is stationary. 
Once the aircraft is moved a short distance, such as one meter, the 
moments, constraining the bend of the main gears bogie beam and the nose 
gear axle, are eliminated so that the bogie beams and the axle deflect or 
bend to their true unconstrained value resulting in an accurate weight 
determination. Because it is impractical to move the aircraft after it is 
loaded in order to determine whether its weight and center of gravity are 
within safe operating limits, the weight and balance system of FIG. 2 is 
provided with a loading error correction system, illustrated in FIG. 5, 
which automatically compensates for the loading effects. Although the 
correction system is shown for the weight and balance system of FIG. 2 
using inclinometers, the correction system of FIG. 5 may be used with any 
system providing a signal representative of the aircraft weight on a 
landing gear such as provided by strain gage sensors, pressure and 
magnetic variable reluctance sensors or the like. 
FIG. 4 illustrates the worst case loading errors in the calculated weight 
and center of gravity, the worst case occurring when the aircraft is 
loaded from its empty weight to its maximum departure weight. From the 
graph it is seen that the error band of the weight and balance system, 
with the automatic loading error correction system shown in FIG. 5, is 
much less than the percent error in the calculated weight and center of 
gravity where the loading errors are left uncorrected. 
The automatic loading error correction system is shown in FIG. 5 for either 
of the main landing gears 18 or 22 or the nose gear 20 where the strut 
weight 86 is continuously sensed from the outputs of the summing circuits 
50, 54 or 70 on lines 52, 56 or 72. The sensed uncorrected strut weight 86 
for a landing gear member is applied on a line 88 to a differencing 
circuit 90 performing a subtraction operation, and on a line 92 to a latch 
94. When the aircraft is moving as determined by a strut movement detector 
96 described in detail below, the detector applies a high signal on a line 
98 to the latch 94 so that the uncorrected weight signal on line 92 is 
passed through the latch to the circuit 90 on a line 99. When the aircraft 
is moving, the uncorrected weight on line 99 is thus equal to the 
uncorrected strut weight on line 88 so that the output of the circuit 90 
is zero. However, when the output of the movement detector 96 goes low 
indicating that the aircraft has stopped moving, the latch 94 holds the 
uncorrected strut weight applied thereto just after no movement is 
detected, the held weight also being stored in the CDM non-volatile memory 
80. During the remainder of the time the aircraft is stationary the signal 
applied on line 98 to the circuit 90 is the uncorrected weight signal 
generated just after aircraft movement has ceased. This held weight signal 
is combined by circuit 90 with the weight signal applied on line 88, which 
signal varies as the aircraft is unloaded and reloaded. 
The signal output from the circuit 90 represents the difference between the 
uncorrected weight signal generated just after no movement has been 
detected and the subsequently generated uncorrected weight signal. A 
multiplier 100 multiplies the difference signal applied thereto on line 
102 by a loading error correction coefficient, K, which is dependent on 
the structure of the bogie beam or nose gear axle being monitored. For the 
nose gear, K is set equal to 0.05 and for each of the main gears, K is set 
equal to 0.025. The loading error compensation signal output from the 
multiplier 100 on a line 104 is applied to a summing circuit 106, which 
combines the uncorrected weight signal 86 applied on a line 112 with the 
loading error compensation signal from line 104 to generate a corrected 
weight signal on a line 114, the corrected weight signal being stored in 
the CDM non-volatile memory 80. 
In order to compensate for fluctuations in the sensed uncorrected weight 
signal 86 caused by the passengers rushing towards the nose of the 
aircraft in order to exit after the aircraft has stopped, the generation 
of the corrected weight signal on line 114 is delayed by a 4-second delay 
116 connected to the output of the summing circuit 106. The 4-second delay 
compensates for changes in the weight distribution before the corrected 
weight signal is stored in the non-volatile memory 80. 
In the event that power has been removed from the system, once the power is 
restored, reasonableness tests are performed to determine whether the 
aircraft has been moved or whether there has been a load change during the 
power down. If aircraft or strut movement is detected, a power down strut 
movement detector 118 outputs a reset signal on a line 120 to the latch 
94, the contents of which are reset to the uncorrected weight 86 applied 
on line 92 after power is restored. If a load change occurring during the 
power down is detected after power is restored, a switch 108 is moved from 
an open circuit terminal 109 to contact a terminal 110 to which is applied 
a power down correction factor. The power down correction factor is 
applied to the summing circuit 106 where it is combined with the error 
compensation signal on line 104 and the uncorrected weight signal on line 
112 to compensate for the larger loading error which would otherwise 
result. The power down reasonableness tests are discussed in detail below 
with reference to FIG. 8. 
The strut movement detector 96, as illustrated in FIGS. 5 and 6, provides a 
high output on line 98 to the latch 94 when the aircraft is moving and 
provides a low output to the latch when no movement is detected. The beam 
or axle angle signal 124 provided by the outputs of the differencing 
circuits 58 and 62 on respective lines 60 and 64, is proportional to the 
acceleration of the aircraft. A typical beam angle signal output from the 
differing circuits 58, 62 is illustrated in FIG. 7 bearing the unique 
signature of its associated landing gear member when moved. The beam angle 
signal 124 representing acceleration is applied to a third order band pass 
filter 126 which acts as a pseudo double integrator, the output of which 
represents strut movement in distance. The particular frequency components 
of interest associated with aircraft movement are typically different for 
the nose gear axle 28 and the bogie beams 24 and 26. As such, the limit 
frequencies F.sub.1 and F.sub.2 for the band pass filter 126, which 
isolates the particular frequency components of interest, are stored in 
the CDM non-volatile memory 80 for each of the landing gear members. For 
the nose gear axle, F.sub.1 and F.sub.2 are set equal to 0.1 hz and 5.0 hz 
respectively and for the bogie beams F.sub.1 and F.sub.2 are set equal to 
0.05 hz and 20 hz respectively. 
If the beam or axle angle signal 124 falls within the shaded area 128 of 
the filter indicating movement, a high signal is output from the filter 
126 and applied to a comparator 130, a zero output being produced by the 
filter if the angle signal is outside of the shaded area. The comparator 
130 compares the signal output from the band pass filter 126 to a .+-.0.3 
reference signal supplied by the CDM non-volatile memory 80. If the 
amplitude of the filter output signal is greater than +0.3 or less than 
-0.3, the comparator produces a high output on the line 98 indicating 
aircraft movement. When the filter output signal is within 0.3 of zero, 
indicating the aircraft has come to a stop, the output of the comparator 
130 goes low causing the latch 94 to hold the uncorrected weight signal 
generated at that time. 
The output of the comparator 130 is applied to a 0.4 second hold delay 134 
to eliminate short term vibrations in the output signal which are not 
indicative of movement. The delay 134 prevents the generation of a high 
output signal on line 98 until a period of 0.4 seconds has elapsed from 
the time that the comparator 130 has generated a high output signal, the 
delay circuit 134 providing an immediate indication, i.e., with no delay, 
when the output of the comparator drops to a low state indicating that 
movement of the aircraft has ceased. 
It is noted that the output of the strut detection circuit is also applied 
to the pilot's display unit 76. A high output, indicating movement, causes 
the aircraft weight and center of gravity indication on the display 76 to 
freeze since movement eliminates the loading errors in the calculated 
weights and center of gravity. However, when zero movement is detected, 
the output of the comparator 130 going low, the output signal unfreezes 
the display 76 to show the corrected and updated weight and center of 
gravity as calculated by the system of FIG. 5. 
The system for performing reasonableness tests in the event that power to 
the system has been lost is shown in FIG. 8. In response to a power up 
signal 136 generated when power to the system is restored after having 
been lost, the uncorrected strut weight on hold since the last detected 
movement 138 and the corrected strut weight 140 generated prior to the 
power down, both of which are stored in the CDM non-volatile memory 80, 
are read and applied to a differencing circuit 142. The output of the 
differencing circuit 142 represents the corrected strut weight 140 minus 
the uncorrected held strut weight 138. The stored corrected strut weight 
140 is also applied to a differencing circuit 144, to the negative 
terminal of which is applied the power up uncorrected strut weight 146 
generated upon the restoration of power. The output of the differencing 
circuit 144 is applied to a multiplier 146 which multiplies the signal 
representing the difference between the stored corrected strut weight 140 
and the power up uncorrected strut weight 146, by the loading error 
correction coefficient K, where K is equal to 0.025 for the main landing 
gears and 0.05 for the nose gear. The output of the multiplier 146 is 
applied to a second multiplier 148 which multiplies the input thereof by 
one-half to provide the power down correction factor on a line 159 to 
terminal 110, the correction factor being equal to K/2 multiplied by the 
difference between the stored corrected strut weight 140 and the power up 
uncorrected strut weight 146. 
In order to determine whether a load change had occurred during the power 
down, the output of the differencing circuit 142, representing the 
difference between the stored corrected strut weight 140 and the held 
uncorrected strut weight 138, is applied to a multiplier 150 which 
multiplies the output of the circuit by the loading error correction 
coefficient K for the particular landing gear being monitored. A 
differencing circuit 152 combines the output from the multiplier 150 with 
the output of the differencing circuit 144 to provide a signal 
representative of the difference therebetween. The output of the 
differencing circuit 152 is applied on a line 156 to a comparator 154 for 
comparison to a reference signal of .+-.0.3. If the signal on line 156 is 
greater than 0.3 or less than -0.3, the output of the comparator 154 goes 
high, indicating that a load change occurring during power down has been 
detected. 
The output of the comparator on line 158, when high, is used to actuate a 
relay or the like to cause the switch 108 of FIG. 5 to contact the 
terminal 110. The power down load correction factor from line 159 is then 
applied throu9h the switch 108 to the summing circuit 106 where it is 
combined with the error compensation signal applied on line 104 and the 
uncorrected strut weight signal applied on line 112 to provide the 
corrected strut weight signal. 
In order to detect whether aircraft movement or strut movement had occurred 
during the power down, the output from the differencing circuit 152 is 
applied to a zero comparator 160, the other input of which is .+-.0.1. If 
the output from the differencing circuit 152 is within 0.1 of zero, the 
output of the comparator 160 generates a high inhibit signal indicating no 
movement during the power down, the inhibit signal being applied on a line 
162 to a comparator 164. A high inhibit signal indicating no movement 
prevents the comparator 164 from going high regardless of the inputs 
applied thereto. 
The output from the differencing circuit 144 representing the difference 
between the corrected strut weight 140 and the power up uncorrected strut 
weight 146 is applied to the comparator 164 through a one-second delay 
166, a .+-.0.1 reference signal being applied to the other input of the 
comparator 164. If the difference between the corrected strut weight 140 
and the power up uncorrected strut weight 146 is within 0.1 of zero and 
there is no inhibit signal on line 162, the output of the comparator 164 
goes high, producing the reset signal on line 120 of FIG. 5 to reset the 
latch 94 to hold the uncorrected strut weight 86 which is the power up 
uncorrected strut weight 146. If the output of the differencing circuit 
144 is not within 0.1 of zero, the output of the comparator 164 is low, 
indicating no detected strut movement. If the output from the differencing 
circuit 144 is within 0.1 of zero but a high inhibit signal is applied on 
line 162 to the comparator 164, the output of the comparator remains low. 
The one-second delay of the output from the differencing circuit 144 
prevents the output of the comparator 164 from going high when a 
determination of no movement is made by the comparator 160 generating a 
high inhibit signal on line 162. 
It is noted that the system of FIG. 8 performs reasonableness tests; that 
is, load changes and gear movement occurring during power down may not 
always result in high outputs on lines 120 and 158. If the system of FIG. 
8 determines that the load changes or gear movements occurring during 
power down are within limits so that weight signal held prior to power 
down and the error compensation signal on line 104 will adequately 
compensate for the errors, the outputs on lines 120 and 158 will be low.