close

Вход

Забыли?

вход по аккаунту

?

Патент USA US3052133

код для вставки
Sept. 4, 1962
l. s. WESTERBACK
3,052,122
FLIGHT PATH ANGLE COMPUTER
Filed Sept. 23, 1959
’l//2
,
BlASED
3
ACGELEROMETER
)
P
(ur'g)
SUMMING
AMPLIFIER
5
l0
/4
AIRSPEED
V
SENSOR
_-_
°°=5V
-
l6‘
8
24
-
RATE OF
CLIMB
5
INDICATOR
EJQALYQPEGR
6
26
MULTIPLIER
6
v sm 8
5
22
SINE
/8
3
FUNCTION
INTEGRATOR
§ GENERATOR
- smb’
‘
L)
20
——>ouTPuT
9
-
VERTICAL
GYRO
28
f
;
cos 6
cos ¢
MULTIPLIER
MULTIPLIER
30
32
SUMMING
347/ AMPLIFIER
5:.
————--ouTPuT
(
FIG .2.
INVENTOR
[VA A’ 5. W55 TERBA CK
BY
C
i Z
ATTORNEY
United grates
1
3,052,l22
Patented Sept. 4, 1962
2
only experienced if the ?ight path angle is changing. The
3,052,122
reason for this is that an aircraft in ?ight must change its
FLIGHT PATH ANGLE COMPUTER
pitch attitude in order to change its ?ight path angle.
Ivar S. Westerback, Glenhead, N.Y., assignor to Sperry
Rand Corporation, Great Neck, N.Y., a corporation of
When the pitch is changed, the forward movement of the
aircraft relative to the air mass in which the craft is lo
Delaware
cated produces a force on the craft which tends to move
Filed Sept. 23, 1959, Scr. No. 841,833
10 Claims. (Cl. 73-178)
7
me
the craft in the direction of its yaw axis. This force on
the aircraft is re?ected in an acceleration, i.e. centripetal
acceleration, the magnitude of which is determined by the
This invention relates to apparatus for use in the navi
gation and control of aircraft, and particularly, to appara 10 aircraft’s speed relative to the air mass multiplied by the
rate of change of ?ight path angle. Since the aircraft’s
tus for controlling the pitch attitude of the craft in such a
centripetal acceleration is determined by the aircraft’s
manner that the craft follows a ?ight path making a de
speed relative to the air mass, i.e. airspeed, multiplied by
sired or selected angle with respect to the horizontal. The
the rate at which the ?ight path angle is changing, a
apparatus is equally applicable for the automatic control
of the aircraft through an automatic pilot, or in the man 15 measure of ?ight path angle may be had by integrating a
signal which is representative of the quotient of the air
ual control thereof, through the use of an indicator indi
craft’s airspeed divided into the aircraft’s total yaw axis
cating the ?ight path angle directly, or through the use of
acceleration less the algebraic sum of all of the individual
an indicating instrument of the ?ight directed type, such
yaw axis accelerations other than the centripetal accelera
as that disclosed in US. Patent No. 2,613,352 in the name
of S. Kellogg II, and assigned to the same assignee as the 20 tion. Computation, as just described, is greatly simpli?ed
in the case where the aircraft pitches and rolls to only a
present invention.
small degree. In such cases, the total yaw axis accelera
Heretofore, it has been proposed to control the pitch
tion is merely the algebraic sum of the centripetal and
attitude of the craft so that the path along which the air
gravitational accelerations. If, however, large pitch and
craft travels makes a predetermined angle with respect to
the geo-horizontal regardless of the craft’s attitude. Such 25 roll angles are experienced the total yaw axis acceleration
is the algebraic sum of the centripetal acceleration plus the
an angle will hereinafter be referred to as the angle of the
gravitational ‘acceleration multiplied by the cosines of the
craft’s ?ight path or the ?ight path angle of the aircraft
pitch and roll angles.
>
and will be designated by the symbol 'y. As is known, the
Measurement of the total yaw axis acceleration may be
?ight path angle may be de?ned as the algebraic sum of
the craft pitch attitude angle 0 and the craft angle of 30 readily made by an accelerometer. However, accelerom
eters, like vertical gyroscopes and angle of attack ‘sensors,
attack a. Signals representing 0 and 0c may be respectively
detect and measure accurately short period changes but
derived from a vertical gyro and ‘an angle of attack sensor.
are inaccurate in detecting and measuring changes which
When algebraically added together, these signals produce
occur over an extended time period. Because of this,
,a signal representing the ?ight path angle 7.
35 computation of ?ight path angle from data derived from
Because the vertical gyroscope usually used in the mea
surement of pitch attitude is subject to drift which occurs
over a comparatively long time period before being cor
rected, measurement of the ?ight path angle by algebraical
an accelerometer will also be inaccurate over long time
periods. To compensate for this long term inaccuracy
in the measurement of the ?ight path angle, the instant
ly summing a pitch signal and an angle of attack signal is 40 invention employs the prior art technique of monitoring
the ?ight path angle measurement accurate only over a
short time period by a ?ight path angle measurement
fore, as proposed in the prior art, a second measurement
which is accurate only over a long time period. This
of ?ight path angle, which is stable over a long period of
monitoring technique as utilized in the present invention
time, is utilized to monitor the ?ight path angle measure
will be explained in detail later.
ment made from the summation of 6 and a.
45
Accordingly, it is the prime object of the present in
Still, the above-described ?ight path angle measurement
vention to provide apparatus for computing an aircraft’s
technique has additional undesirable aspects. The meas
?ight path angle without the need for angle of attack data,
urement of the angle of attack at is usually made by a
the computation being accurate over both long and short
probe extending from the wing of the aircraft and, be
cause of the probe’s location and the aircraft’s structural 50 time periods.
Another object of the invention is to provide apparatus
deformation in the airstrearn, the measure of angle of at
which does not require an accurate measurement of an
tack is susceptible to large errors. In addition, the probe
aircraft’s airspeed to accurately compute said aircraft’s
is a protuberance from the aircraft and causes an increased
insufficient to insure accurate measurement of '7. There
?ight path angle.
drag on the air frame by disturbing the air ?ow around the
Another object is to provide apparatus to accurately
craft.
55
compute the ?ight path angle of an aircraft, said apparatus
To overcome this disadvantageous need for an angle of
requiring the detection and accurate measurement of sig
attack sensor, the apparatus of the present invention
nals representing the short period changes of only a single
computes, in place of the algebraic summation of pitch
system variable.
angle and angle of attack, an expression of ?ight path
These and other objects of the invention will ‘become
angle which is equally as responsive over a short time 60
period, ‘but which requires no angle of attack data.
appreciated as the same becomes understood in the light
The instant apparatus computes the total of the accel
of the speci?cation and drawings of which——
erations which the aircraft experiences along the aircraft’s
FIG. 1 is a block diagram of a preferred form of
yaw axis. One of these yaw axis accelerations is centrip
the invention; and
etal acceleration. Centripetal acceleration, however, is 65 FIG. 2 is a block diagram of a circuit which, when
3,052,122
4
operable with the embodiment shown in FIG. ‘1, makes
erroneously low multiplier 22 output signal V sin 'y is
the operation of that embodiment accurate in the pres
ence of large roll and pitch angles.
Referring to FIG. 1, an accelerometer 11!}, which is
subtracted from the climb indicator output signal h
which is stable over long time periods. This summing
mounted in the aircraft in such a way that it measures the
?er 12 wherein it is algebraically added into the computa
tion of the yaw axis acceleration. Since the summing
ampli?er 12 has two erroneous signals applied thereto,
aircraft’s total accelerations car along the yaw axis, is biased
by an amount equal to the gravitational acceleration g to
provide an output signal avg.
This Signal aT——g is
coupled to a summing ampli?er 12 together with the
ampli?er 26 output signal is coupled to summing ampli
one being too high and one being too low, the errors
cancel each other and permit accurate computation of
output signal from the vsumming ampli?er 12. If the 10 the ?ight path angle '1. Obviously, for accelerometer 10
total acceleration aT should change due to a change in
output signals which are erroneously high the opposite
flight path angle the output acceleration signal from sum
will occur.
ming ampli?er 12 will correspondingly change. Since
The present invention also does not require an ex
tremely accurate measure of airspeed to accurately com
the output signal is fed back to the summing ampli?er 12,
it tends to null out the effects of the change in total 15 pute the ?ight path angle of the craft. That is, minor
acceleration aT. Nulling out the change in aT, i.e. mak
errors in the measurement of airspeed can be tolerated.
ing the algebraic sum of all yaw axis accelerations equal
For instance, when the airspeed signal V is slightly smaller
to Zero, can only be achieved if the feedback signal
in value than it should be, the output signal from di
equals the centripetal acceleration.
vider 14 will be slightly erroneously high. Consequently,
The centripetal acceleration signal ac at the output of 20 integrator 18 and sine function generator 20 will also
summing ampli?er 112 is coupled to a divider 14- together
have slightly erroneously high output signals. However,
with a signal V representative of the craft’s airspeed, said
multiplier 22 produces the product of the erroneously
airspeed signal being ‘derived ‘from an airspeed sensor 16.
low signal V and the erroneously high output signal from
Since, as was earlier stated, centripetal acceleration is the
sine function generator 20. Since these signals have
product of the rate of change of ?ight path angle multi 25 errors which are in opposite directions, the errors tend
plied ‘by the craft airspeed, division of a signal representing
to cancel each other, resulting in a fairly ‘accurate multi
the centripetal acceleration ac by a signal representing the
plier 22 output signal. This multiplier ‘22 output signal is
craft airspeed V will yield a signal representative of the
then used, as described above, to aid in correcting the
rate at which the ?ight path angle is changing. The
summing ampli?er 12 output signal.
rate of change of flight path angle signal available at 30
Referring to FIG. 2, a block diagram is shown of a
the output of divider '14 is coupled to an integrator 18
circuit whose output signal, ‘when coupled to the input
wherein it is integrated to provide a signal representing
of summing ampli?er 12 of FIG. 1, enables the circuit of
the ?ight path angle ",1. Integrator 18 also serves to
FIG. 1 to accurately compute the ?ight path angle 7
?lter out spurious signals which arise in the computa
during periods when large aircraft roll and pitch angles
tion of the ?ight path angle and which are coupled to 35 are experienced. In FIG. 2 a vertical gyro 28, from
integrator 18.
which signals representative of the cosines of the roll
However, as earlier stated, the signal 7, because it is
angle (,6 and the pitch angle 6 may be derived, applies
derived from accelerometer output data, is subject to
these signals to a multiplier 30. Multiplier 30 operates
errors which tend to accumulate over extended time
periods. Therefore, long term compensation is provided
as follows:
It is known that the aircraft’s rate of climb divided by
its airspeed is equal to the sine of the ?ight path ‘angle 7.
Since the aircraft’s rate of climb is measured by baro
metric elements which are free of cumulative errors, a
to produce a signal representing the product, cos ¢ cos 0.
40 The signal cos ¢ cos 0 is coupled to a multiplier 32 to
gether with a signal representative of the gravitational
acceleration g to produce a signal g cos 0 cos ¢.
This
last-mentioned signal is representative of the true yaw
axis acceleration due to gravity which the aircraft ex
periences at times when aircraft roll and pitch angles
‘are large. Since the accelerometer 10 is biased by a
signal representative of the gravitational acceleration g to
simplify computation of the centripetal acceleration ac
signal representing rate of climb is utilized to correct the
computation of the yaw axis acceleration. The integrator
output signal v is coupled to a sine function generator 20
to produce a signal representing sin 'y. The signal sin 'y is
in FIG. 1, use of the biased accelerometer 10 in sys
coupled to a multiplier 22 together with a signal V derived 50 tems which accurately measure the yaw axis acceleration
from the airspeed sensor 16 to produce the output signal
due to gravity requires that the accelerometer bias signal
V sin 'y. This signal, because it is derived from apparatus
g be cancelled out. A summing ampli?er 34 is provided
having only short term accuracy is likewise only accurate
in the circuit of FIG. 2 to enable the cancellation of the
for short terms. A climb indicator 24 provides an output
accelerometer bias signal g. Summing ‘ampli?er 34 has
signal 71', representing the aircraft’s rate of climb. This
coupled thereto a signal g cos 0 cos ¢ representing the
signal it is coupled to a summing ampli?er 26 together
true yaw axis acceleration due to gravity and a signal g
with the signal V sin 'y. Summing ampli?er 26 operates
representing the gravitational acceleration. Summing am
to produce a signal representative of the difference be
pli?er 34 provides an output signal which is representative
tween h signal and said V sin 'y signal. The output signal
of the difference between its two input signals. When
from summing ampli?er 26 is then coupled to the input 60 the accelerometer 10 output signal cur-g and the output
of summing ampli?er 12 to be algebraically added into
signal 3 cos 0 cos ¢—g from summing ampli?er 34 are
the summation of aT—g and ac. Without the coupling
between summing ampli?ers 12 and 26, when the centrip
etal acceleration ac is in error due to a long term error
being generated by accelerometer 10, the output signal 7
from integrator 18 will be correspondingly in error. With
the coupling in, however, the error will be corrected.
For instance, when the accelerometer output signal is
erroneously low, the summing ampli?er 12 output signal
applied to summing ampli?er 12 for accurate measure
ment of the ?ight path angle 7 during periods when the
aircraft experiences high pitch and roll angles, the bias
accelerations cancel each other and the remaining ac
celeration signals are the total yaw axis acceleration my,
the centripetal acceleration ac and the acceleration due
to gravity g cos 6' cos ¢. These remaining accelerations
is correspondingly lower than it should be. This being 70 when algebraically added together equal zero.
While the invention has been described in its preferred
the case, the divider 14 output signal, the integrator 18
output signal, the sine function generator 20 output signal
embodiments, it is to be understood that the words which
and the multiplier 22 output signal will all be erroneously
have been used are words of description rather than of
low. However, the output signal from summing ampli?er
limitation and that changes within the purview of the
26 will be higher than it should be. This is because the 75 appended claims may be made without departing from
3,052,122
r
J
the true scope and spirit of the invention in its broader
aspects.
What is claimed is:
1. Apparatus for computing the ?ight path angle of an
aircraft comprising means computing ?ight path angle
from an expression of the aircraft acceleration along the
yaw axis, said expression including terms which are ac
fi
senting the ‘algebraic sum of said ?fth and sixth means
outputs, said seventh means being also coupled to said
?rst summing means to algebraically add said seventh
means output signal to the signal representing said
centripetal acceleration.
5. The structure of claim 2 wherein said computing
means comprises means producing a signal representing
the total yaw axis ‘acceleration of said aircraft, means
producing
a signal proportional to the gravitational ac
second expression of aircraft ?ight path angle, said sec
ond expression including terms which are accurate over 10 celeration and third means algebraically adding to the yaw
axis acceleration signal and the gravitational acceleration
a long period of time, means comparing the results of
signal to compute the centripetal acceleration of said air
both said computations to produce the difference there
craft.
between, and means algebraically adding said difference
6. The structure of claim 2 wherein said computing
to the result of the computation of said ?rst expression,
means comprises means producing a signal representing
thereby providing a measure of ?ight path angle which
the total yaw axis acceleration of said aircraft, means
has both long and short term ‘accuracy to instill long
producing a signal proportional to the gravitational ac
term ‘accuracy in the result of the computation of said
celeration, means producing a signal proportional to the
aircraft acceleration.
cosine of the angle said aircraft pitches, means coupled to
2. Apparatus for computing the ?ight path angle of an
receive
the gravitational acceleration signal and the signal
aircraft comprising computing means producing a signal 20
representing the cosine of the pitch angle to produce a
representing the centripetal acceleration of said aircraft,
signal proportional to their product, and means coupled
?rst summing means coupled to receive said acceleration
to
receive the yaw axis acceleration signal and the product
signal, second means producing a signal representing the
signal to produce a signal proportional to their algebraic
airspeed of said aircraft, third means coupled to said
‘sum, said last mentioned means output signal being pro
?rst summing means and said second means to produce a
portional to the centripetal acceleration of said aircraft.
signal representing the quotient of said ?rst summing
7. The structure of claim 2 wherein said computing
means and said second means outputs, fourth means cou
means comprises means producing a signal representing
pled to said third means to produce a signal representing
the total yaw axis acceleration of said ‘aircraft, means
the integral of said third means output, said integral sig
producing
a signal proportional to the gravitational accel
nal being a signal-representing ?ight path angle, ?fth
curate over a short period of time, means computing a
means coupled to said fourth means to produce a signal
representing the sine of said fourth means output, sixth
means coupled to said second means and said ?fth means
eration, means producing a signal proportional to the
cosine of the angle said aircraft rolls, means coupled to
receive the gravitational acceleration signal and the signal
representing the cosine of the roll an?le to produce a signal
proportional to their product, and ?fth means coupled to
receive the yaw axis acceleration signal and the product
signal representing the rate of climb of said aircraft, ‘and
signal to produce a signal proportional to their algebraic
eighth means coupled to said sixth and seventh means
sum, said last mentioned means output signal being pro
to ‘produce a signal representing the algebraic sum of
portional to the centripetal acceleration of said aircraft.
said sixth and seventh means outputs, said eighth means
8. The structure of claim 2 wherein said computing
being also coupled to said ?rst summing means to alge 40
means comprises means producing a signal proportional
braically add said eighth means output signal to the signal
to produce a signal representing the product of said sec
ond and ?fth means outputs, seventh means producing a
representing said centripetal acceleration.
3. Apparatus for computing ?ight path angle of an
to the total yaw axis acceleration of said aircraft, means
producing a signal proportional to the gravitational accel
eration, means producing a signal proportional to the
aircraft comprising means producing a signal represent
ing the aircraft acceleration along the yaw axis of the 45 cosine of the angle said aircraft pitches, means producing
a signal proportional to the cosine of the angle said
craft, second means producing a signal representing the
aircraft rolls, means coupled to receive the signals repre
acceleration due to gravity, third means coupled to said
senting the cosines of the pitch and roll angles to produce
?rst and second means to produce a signal reprmenting
a signal proportional to their product, means coupled
the algebraic sum of said ?rst and second means output
signals, fourth means producing a signal representing the 50 to receive the gravitational acceleration signal and the
product signal to produce a signal proportional to their
airspeed of said aircraft, ?fth means coupled to said
product, and means coupled to receive the output signal
third and fourth means to produce a signal representing
from said last mentioned means and the yaw axis accel
the quotient of said third and fourth means output signals,
eration signal to produce a signal proportional to their
and sixth means coupled to said ?fth means to produce a
signal representing the integral of said ?fth means output 55 algebraic sum, said last mentioned means output signal
being proportional to the centripetal acceleration of said
signal, said sixth means output signal being a signal repre
aircraft.
senting the ?ight path angle of said aircraft.
:1
4. Apparatus for computing the ?ight path angle of an
aircraft comprising computing means producing a signal
representing the centripetal acceleration of said aircraft,
?rst summing means coupled to receive said acceleration
signal, second means producing a signal representing the
airspeed of said aircraft, third means coupled to said
9. Apparatus for computing the ?ight path angle of
an aircraft comprising means producing a signal repre
senting the centripetal acceleration of said craft, means
producing a signal representing the speed of said craft,
?rst computing means receiving said acceleration signal
and said speed signal producing therefrom a ?rst signal
representing the craft ?ight path angle, means producing
a signal representing the quotient of said ?rst summing 65 a signal representing the craft rate of climb, second com
puting means receiving said climb rate signal and said
means and said second means outputs, fourth means cou
speed signal producing a second signal representing ?ight
pled to said third means to produce a signal representing
path angle, comparison means receiving said ?rst and
the integral of said third means output, said integral sig
second ?ight path angle signals producing a signal repre
nal being a signal representing ?ight path angle, ?fth
?rst summing means and said second means to produce
means coupled to said second and fourth means to pro
senting the error therebetween, and means receiving said
duce a signal representing the product of said second
?rst ?ight path angle signal and said error signal alge
braically summing said signals to produce a signal accu
means output and a function of the output of the fourth
rately representing the ?ight path angle of said craft.
means, sixth means producing a signal representing the
10. Apparatus for computing the ?ight path angle of
rate of climb of said aircraft, and seventh means coupled
to said ?fth and sixth means to produce a signal repre 75 an aircraft comprising means producing a signal repre
3,052,122
U
senting the centripetal acceleration of said aircraft, means
producing, a signal representing the speed of said aircraft,
means receiving said acceleration signal and said speed
signal producing a signal representing the quotient thereof,
means receiving said quotient signal producing a signal 5
representing the time integral of said quotient signal,
said integral signal being a ?rst signal representing ‘the
flight path angle of the craft, means producing a signal
representing the craft climb rate, and computing means
receiving said climb rate signal and said speed signal 10
producing a second signal representing the craft ?ight
path angle, comparison means receiving said ?rst and
second ?ight path angle signals producing a signal repre
senting the error therebetween, and means receiving said
?rst ?ight path angle signal and said error signal alge
braically summing said signals to produce a signal accu
rately representing the ?ight path angle of said craft.
References Cited in the ?le of this patent
UNITED STATES PATENTS
' 2,896,145
Snodgrass ___________ __ July 21, 1959
2,932,467
Scorgie _-_ ____________ __ Apr. 12, 1960
2,934,267
Wirkler et al. ________ __ Apr. 26, 1960
Документ
Категория
Без категории
Просмотров
0
Размер файла
642 Кб
Теги
1/--страниц
Пожаловаться на содержимое документа