Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.
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This application is a divisional of application
Serial No. 365,~04 filed November 25, 1980.
BACKGROUND OF THE INVENTION
This invention pertains to weight measuring systems
for an aircraft which measures the bend in a member which is
caused to bend by carried weight on the aircraft by means of
inclinometers and, more particularly, servoed accelerometers
mounted to the member and which have signal outputs representing
the angle of the member with respect to a reference plane and
which are utilized to determine the angle of bend in the member
as an indication of weight supported thereby.
A pilot in command of an aircraft must, before each
flight, determine that the aircraf-t's weight and-balance are
within the safe operating limits or boundaries of the
aircraft. This is typically accomplished by calculating the
total airplane weight and the distribution of that weight in
order to determine balance or center of gravity. The weight
and balance determination is usually made by auditing a
list of cargo, fuel, passengers, oil and crew members, there
frequently being estimation of passenger and crew member
weight and distribution thereof. Cargo may be actually
weighted or estimated. Fuel weight is determined by quantity
added to the aircraft and its temperature estimated to give
an approximate fuel weight. All of this is then added to the
basic empty weight and distribution of the empty weight of
the aircraft as determined at either the date of manufacture
or by subsequent weighing on scales.
Naturally, there are errors in such estimates and
human calculations but, for the majority of flights, the
built-in safety factors including extra runway length and
better than minimum airplane performance and control are
adequate to cover these errors. Even so, there have been many
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l examples of past incidents and accidents where it was later
determined that large errors in weight and balance were not
detected by the pilots. In air freight operations, there
is a large variation in loads, types of caryo, fuel and
airport conditions whereby there is a much greater possibility
for a serious error.
To further improve flight safety, an onboard weight
and balance system allows the pilots to cross-check data on
load and fuel normally provided by others. With sufficient
accuracy and reliability, the onboard weight and balance
system can be the primary ins-trument for the determination
of weight and balance. The system can assist in positioning
cargo at a more nominal balance point to reduce aerodynamic
drag which results in economic savings because of recluced
use of fuel with adequate flight control stability margins for
the pilots.
In recent years, several onboard weight and balance
systems have been developed for use on aircraft, including
the use of strain gages and pressure and magnetic variable
reluctance sensors. They have not measured up to expectations
because of problems in stability, accuracy, reliability and
ability to survive in a harsh environment. Because of these -
problems, many or most of these systems have been disconnected
or removed.
An added safety factor in aircraft operation is a
full complement of properly pressurized landing gear tires
which has required visual inspection and which can be difficult
to accomplish under conditions such as extreme winter weather.
SUMMARY OF THE INVENTION
A primary feature of the invention disclosed herein
is to provide a weight measuring system which utilizes the
angle of bend of a structural member as a direct indication
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of weight or load on the supported structure, such as an
aircraft.
The weight system utilizes inclinometers positioned
to measure the angle of bending in a structural member,
such as a bogie beam, axle or other structure, such as an
aircraft wing or fuselage. The angle or bending is
proportional to the weight or force on the structural
member. The inclinometers provide output signals which can
be summed to cancel out attitude and acceleration changes,
whereby only the beam bend angle is obtained and which is a
direct indication of the weight of the supported weight or
load.
Various embodiments of the weight system are disclosed
and, in each embodiment, servoed accelerometers are used
which are used at spaced locations to provide voltage outputs
which are summed to provide a signal indicative of the weight
causing a predetermined bend angle in the structural member.
In one embodiment, the inclinometers are associated
with the landing gear of the aircraft, including a pair of
inclinometers mounted on a bogie beam member for each of
the one or more main landing gears and a pair of inclinometers
similarly mounted on the gears and a pair of inclinometers
similarly mounted on the axle of the nose gear, whereby the
sum of the signals from the various pairs of inclinometers
provide an indication of aircraft weight.
In another embodiment and which may be in addition to
that previously described, the inclinometers are arranged at
various locations along the length of the aircraft wings and
positioned relative to variable weight sections, such as
fuel tanks, whereby the angle of bend of the wing can be
determined and thereby the weight of fuel carried in a fuel
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Additionally, inclinometers-can be located at various
positions along the length oE the fuselage of -the aircraft
to again determine bend angles at various locations and,
thus, the location and weight of loads carried by the
aircraft.
BRIEF DESCRIPTION OF T~E DRAWINGS
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Fig. 1 is a side elevational view of a typical
aircraft with which the invention disclosed herein is
associated;
Fig. 2 is a diagrammatic plan view of the aircraft
landing gear and having a block diagram showing the control
with sensor locations and computer interface therewith;
Fig. 3 is an enlarged elevational view, partly in
section, of one of the main landing gear and taken along the
line 3-3 in Fig. 4;
Fig. 4 is an end elevational view, looking toward
the right-hand end, as shown in Fig. 3;
Fig. 5 is a diagrammatic view of the main landing
gear shown in Fig. 3;
Fig. 6 is a front elevational view of the aircraft
shown in Fig. l;
Fig. 7 is a block diagram of the part of the weight
and balance system utilizing inclinometers associated with
the wings and fuselage of the aircraft; and
Fig. 8 is an electrical schematic of one of the
servoed accelerometers, which is used as an inclinometer.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
A first embodiment of the invention is shown in
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Figs. 1 to 5 wherein the weight and balance system will
determine aircraft weight and center of gravity by measuring
the weight on each of the main landing gear and the nose
gear.
The aircraft, indicated generally at A, is shown
in a particular configuration only for purpose of
illustration. The invention disclosed herein can be used
with all types of aircraft now utilized, as well as many of
those that will come into use in -the future. The aircraft
has a fuselage 10, with a pair of wings, the right wing
being shown at 11, and mounting a jet engine 12. In the
illustrated embodiment, the aircraft has main landing gear
and nose gear, with one of the retractable main landing
gear supported from the aircraft wing being identified at
15 and the nose gear at 16.
Referring more particularly to Fig. 2, the aircraft
additionally has main landing gear 17.
The aircraft weight is determined by measuring the
weight on each of the main landing gear and on the nose
gear. This measurement is accomplished by sensing the
amount of deflection or bend in the bogie beams of the main
landing gear and by suitable means associated with the nose
gear. This latter means can be accomplished by sensing
the amount of deflection or bend in the axle of the nose
gear.
The angle of bending of the structural member, such
as the bogie beam is proportional to the weight or force Oll
the member. Referring to the schematic view of the main lan-
ding gear, shown in Fiy. 5, a bogie beam 20 is shown with an
exaggerated bend to illustrate the principle of the
invention. An inertial reference plane is indicated by a
horizontal, broken line 21. The angle of bending of the
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beam is represented by the reference angles ~1 and ~2
The angle ~ for a uniorm beam is a function of the
displacement caused by the weight or force on the beam and
the length of the beam. The angle of bending for a beam
with uniform bending is four times the beam displacement
divided by the length of the beam, while the angle for a
beam with uniform cross section is three times the
displacement divided by the lenyth of the beam. Knowing the
length of the beam and measuring the angle of bending, then
provides an indication of displacement.
If the angle of bending is properly measured by
sensing means providing an output representative of the
angle, it is possible to use this angle along with a
constant, which is the relationship of the load or weight
deflection constant on the beam. This will depend on the
beam material, its length to cross section, its modulus of
elasticity, as well as other factors. This constant will
remain a fixed value as long as the beam load is kept within
the linear yield of the beam and the beam is not deformed by
other factors.
Measurement of the angle of bending is more
particularly shown in ~igs. 2 to 5 wherein the main landing
gear 15 and 17 each have a pair of inclinometers associated
therewith. Referring particularly to Figs. 3 and 4, the
well-known type of main landing gear 15 has a retractable
shock strut 25 which is retractably positionable within the
envelope of the aircraft and which has a pair of braces 26
and 27 pivotally connected to a collar 28. A torque
linkage, indicated generally at 29, has members pivotally
interconnected. The lower end of the shock strut mounts a
bogie beam 20 which, at opposite ends thereof, mounts
transverse axles 31 and 32 which are shown in section and
which mount the respective wheel assemblies 33 and 34 having
pressurized tires 33a-d and 34a-d. The weight of the
aircrafk supported by the main landing gear 15 acts
downwardly centrally of the bogie beam 20 to cause bending
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1 thereof in the manner sho~n in Fig. 5. This angle of bend
is measured by a pair of inclinometers ~0 and 41 whlch have
their casings mounted to opposite ends of the bogie beam.
Each of the inclinometers ~0 and 41 has a signal output-
representing the angle of the end of the beam relative to a
horizontal reference plane perpendicular to the force of
gravity. For reasons set forth hereinafter, the use of the
inclinometers in pairs results in the summed output of the
two inclinometers being an indication o~ angle of bend of the
bogie beam regardless of the angle of the aircraft resulting
from other factors, such as the tilt of an airport ramp or
runway upon which the aircraft rests. The main landing gear
17 has a pair of inclinometers 42 and ~3 mounted adjacent
opposite ends of a bogie beam 44. The nose gear 16 is shown
as having an axle 45 and with a pair of spaced-apart inclino-
meters 46 and 47 associated therewith. The inclinometers
associated with the main landing gears and the nose gear sense
the angle of the bogie beams or the axle with respect to
an inertial reference.
Servoed accelerometers are utilized as the
inclinometers to provide the angle measurements, with the
servoed accelerometer being a commercially-available product.
The assignee of this application markets a Q-FLEX accelerometer
which contains a seismic element and servo electronics all
in one miniature package. This basic structure is disclosed
in the applicant's U.S. Patent No. 3,702,073 which issued
November 7, 1972. The seismic assembly is electronically
held in the center position to provide a high level voltage or
current output. This gives an accurate and continuous
measurement of the angle taken by the supporting structural
member, such as the bogie beam. An electrical schematic
for the servoed accelerometer is shown in Fig. 8. A
pendulously-supported proof mass 50 has a high compliance
along a sensitive axis, indicated generally at 51, as
provided by a flexure mounting 52. Pick-off
means, in the form of capacitors 53 and 5~,
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are disposed at opposite sides of the proof mass 50 which
causes the capacitance of one capacitor to increase and the
other to decrease when the proof mass in subjected to
acceleration. ~hrough the servo system, indic~ted as 56 and
as shown in Fig. 5 of the above-mentioned patent, a current
is caused to flow through a pair of coils 57 and 58 which
interacts with magnetic fields provided by magnets 59 and 60
to produce a force on the proof mass which counteracts the
tendency of the values of the capacitances S3 and 54 to
changeO The signal voltage developed across the resistor R
represents acceleration measured by the accelerometer an~
the voltage developed across the resistor R varies in
proportion to the measured acceleration.
The accelerometer is rigidly attached to the
associated structure, with the accelerometers 40 and 41,
previously identified, being mounted at opposite ends of the
bogie beam 20 and each having their sensitive axis 51
aligned with the axis of the beam with no load applied to
the beam. The casing of the accelerometer fixedly mounted 20 to the beam is caused to follow the angle of bend of the
beam and the proof mass 50 of the accelerometer will give an
indication of the inclination or actual angle relative to a
plane which is perpendicular to the force of gravity.
Similarly, accelerometers are used as inclinometers 42 and
43 associated with the bogie beam 44 and as the
inclinometers 46 and 47 associated with the nose gear axle
The angles ~1 and ~2 sense the angle of the
supporting structure with respect to an inertial reference
and with the sensed angles being defined as follows
(Formulas 1 and 2):
~1 ~B ~ ~Ll ~ ~Al (1)
~2 0B ~1,2 ~A2 (2)
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In the foregoing equations, 0B is the angle of the beam or
axle caused by airport ramp or runway tilt. ~Ll and ~L2
are the beam bend angles caused by a load. ~Al and ~A2
are sensor axis misalignment and bias terms.
Weight is given by the follo~ing formula:
WT = K (9Ll ~ ~L2)
From the foregoing, it will be seen that the weight on
the beam is proportional to the ~L components of the total
measured angles. K is the scale factor which depends on
beam or axle geometry and strength, as previously referred
to. Referring to bogie beam 20, the outputs of the two
servoed accelerometers 40 and 41 are summed to cancel out
the beam angle ~B caused by airport ramp or runway tilt
and the angle factors ~A are measured during automatic
zeroing of the system. In summary, weight on the bogie beam
or a nose gear axle is proportional to the sum of the output
signals of the two accelerometers associated therewith.
Referring particularly to Fig. 2, the outputs of the
inclinometers (servoed accelerometers) 40 and 41 are
directed through lines 60 and 61 to a known type of means 62
for summing thereof such as a voltage summing circuit, with
the output thereof directed to a computer 63. Similarly,
the inclinometers 42 and 44 have their respective signal
outputs delivered through lines 64 and 65 to a means 66
which performs a summing function and outputs to the
computer 63. The inclinometers 46 and 47 associated with
the nose gear axle have their signal outputs delivered
through lines 67 and 68 r respectively, to a means 69
performing a summing function which outputs to the computer
63. The computer 63 which can be simply digital or analog
computer determines the weight on each landing gear and the
total weight of the aircraft and causes weight readings at a
pilot display unit 70. ~he center of gravity can also be
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determined directly from the geometry of the aircraft and
the weight on each of the landing gear using well known
formulas for the computate of the center of gravity of a
body. Using the weights on the nose gear and main landing
gear in a center of gravity formula that includes constants
based on the airplane geometry, it is possible to have an
output which indicates the center of gravity.
The embodiment shown in Figs. 6 and 7 illustrates a
weight and balance system where there is not suitable room
or provision to install inclinometers on the landing gear as
well as showing a technique which can be used to determine
wing and tail loading as well as distribution of fuel and
fuselage loads while airborne in addition to functioning as
a weight and balance system. In this embodiment, servoed
15 ` accelerometers, such as shown in Fig. 8, are used as
inclinometers and are installed in the wings and fuselage.
Enough accelerometer must be utilized to accurately
determine various load distribution, such as can happen in
wings having multiple fuel tanks.
Referring to Fig. 6, the aircraft has the right wing
11 previously identified and also a left wing 75, with t~e
wings each having a pair of fuel tanks. The right wing 11
has fuel tanks in the areas indicated generally at 76 and 77
and the left wing 75 has fuel tanks indicated generally in
the areas 78 and 79. It should be noted that the bend o
the wing has been exaggerated for illustrative purposes.
For measuring wing loading the example given, there
are 7 accelerometers shown in association with the fuselage
and wings and which are identified at 80-86.
Various deflection constants can be obtained by
measuring the wing with various fuel loads and fuselage
weight and accurately calculating the differential structure
beam bend angles. The structural bend or beam angle across
accelerometers 81 and 82 will be directly related by the
fuselage weight times a constant, the fuel weight in tank
located at 77 times a constant and the fuel weight in the
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tank located at 76 times a constant. As shown in Fig. 7,
the outputs of the accelerometers 80-86 are fed to a summing
and computer unit, indicated generally at 90, with the
output thereof being delivered to a display unit 91.
Similarly, the fuselage bend can be measured and the
load and its distribution determined by locating a plurality
of accelerometers along the length of the fuselage and as
indica~ed generally by accelerometers 92, 93 and 94 which
all have their signal outputs delivered to the summing and
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It will be noted that the system shown in Figs. 6 and
7 can be utilized while airborne to determine wing and tail
loading and distribution of fuel and fuselage loadsO
With the systems disclosed herein utilizing
accelerometers offering long-term stability, infinite
resolution, and eliminating many of the limitations of
strain gages and various types of transducers, it is
possible to have an onboard weight and balance system which
can become a primary instrument for the determination of
weight and balance. With such a system, there is improved
flight safety as well as econornic savings in operation
because of controlling the location of cargo within the
aircraft to reduce aerodynamic drag.
The weight and balance system has substantially all of
the components required for low tire pressure or blown tire
detection. The bogie beam tilt angle with respect to an
inertial reference will be small if all tires are properly
pressurized. A low tire will cause detectable bogie beam
tilt and torsional twist. A 75% depressurized tire causes a
bogie beam tilt of about 1.4 and a torsional twist angle
of 0.5 on a level runway.
For weight and balance operation, the ~B terms in
previously given Formulas 1 and 2 were cancelled by summing
~1 and ~2. For low tire detection, however, the ~B
terms are of interest. By subtracting 0~ from ~1~ the
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~B terms are of interest. By subtracting ~2 from ~1'
the ~B term can be measured:
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~ Ll and 3L2 (bend angles caused by load) are equal
and cancel out. ~L2 (bend angles caused by load) are
egual and cancel out. ~Al ~ aA2 is due to alignment
error which the system will measure automatically~
Each of the main landing gear 15 and 17 ~Fig. 2) has
four tires 33a-d and 34a-d. The beam tilt angle (~B) will
identify which pair of tires (front or back) has a low
tire. In order to determine which of the front tires or
which of the rear tires (left or right~ is low, two cross
axis inclinometers 100,101 and 102,103 are installed on the
respective bogie beams 15 and 17. One cross axis
inclinometers is on the front of the beam and one is on the
rear of the beam. The inclinometers 100-103 ea¢h have their
sensitive axis at right angles to that of the adjacent
inclinometers 40~43 and output through respective lines
105-108 to the computer 63. Means 62A and 66A include known
circuitry for performing the subtraction function. A low
tire will cause a torsional twist of the beam which can be
measured by the cross axis inclinometers.
In practice, runway slopes and variations will impact
the sensed bogie beam anglesO The effects of runway slopes
and variations are minimized in three ways. First, a
fuselage mounted inclinometer senses runway slope so that
its effect can be eliminated from the measurements. Second,
the two bogie beam angles are compared so that ~ariations
between the two angles will be detected. Finally, the twist
angles measured by the front and back lateral inclinometers
on each bogie are compared so that the effects of lateral
runway slopes wlll be minimized.
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In summary, outputs of the system inclinometers can
detect and identify blown or underpressured tires on the
main landing gear.
