Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.
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GROUND PROXIMITY WARNING SYSTEM FOR
~SE WITH AIRCRAFT HAVING DEGRADED PERFORMANCE
1 Technical Field
This invention relates to the field of aircraft
ground proximity warning systems and, in particular, to
systems that provide enhanced warnings in the event of
degraded aircraft performance near the ground.
Backqround of the Invention
_ _ _
Ground proximity warning systems -that provide
warnings of potential impact with the ground under
controlled flight conditions have been developed over the
past fifteen years. Examples of such systems are disclosed
in U.S. patents 3,946,751; 3,947,810; 4,060,793; 4,319,218
and 4,433,323. One of the objects of the ground proximity
warning systems illustrated in the above patents is to
utilize sensors that are normally present in commercial
aircraft, such as the radio altimeter, barometric altimeter
and glide slope receiver to provide the aircrew with timely
warnings of an impending but inadvertent contact with the
ground. These systems have generally proved to be highly
effective in preventing controlled flight into terrain type
accidents.
However, there are flight situations where the
performance of the aircraft itself becomes degraded and in
certain of these situations existing ground proximity
warning systems may not provide as timely a warning as might
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1 be desired. Reasons for degraded aircraft performance are
many and varied and as such include: wind shear, etc.;
improper configuration including gear down, partial
spoilers, flaps, etc.; degraded lift from rain, ice, excess
weight, improper flap settings, etc.; insufficient engine
thrust; and instrument errors leading to inappropriate
changes in thrust, attitude or airspeed. When reviewed with
respect to past aircraft accidents involving degraded
performance neither existing ground proximity warning Mode 1
which is the excessive descent rate warning mode described
in U.S. patent 4,060,793 nor mode 3 which is the negative
climb after takeoff warning mode described in U.S. patent
4,319,218 would always provide as much warning as might be
desired. For example, in certain wind shear situations the
warning generated by existing Modes 1 and 3 may not be
timely enough to be useful.
In addition to giving timely alerts it is also
highly desirable to give the aircrew an indication as to
what should be done to recover from a dangerous situation
especially under unusual circumstances such as wind shear or
misIeading instrument readings. For instance, there have
been situations where an aircraft has struck the ground
which could have been avoided if the aircrew had appreciated
that the aircraft had additional performance immediately
available in terms of airspeed that could have been
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1 converted to altitude or that additional ~hrust could have
been applied.
: With respect to degraded performance due to wind
she~r, there have been a number of proposed systems, as
described, for example, in U.S. patents 4,043,194;
4,079,905; 4,229,725; 4,281,383; 4,3~2,912 and 4,336,606,
for alerting an aircrew to a wind shear condition. However,
such systems are often difficult to implement or require
additional sensors or do not provide usable information in a
timely manner.
In one approach described in U.S. patent
4,189,777, airspeed rate is used to detect a wind shear
condition and in response thereto a ground proximity warning
system Mode 1 warning curve is modified to increase warning
time. Another approach relating to wind shear conditions is
described in U.S. patent 4,347,572 in which angle of attack,
stick shaker valùe, vertical speed, airspeed, flap position,
and thrust are used to provide climb out guidance on a pilot
flight director display in a wind shear situation.
None of the systems described above provide
enhanced ground proximity warning or guidance for a
comprehensive set of degraded aircraft performance
situations.
Summary of the Invention
It is ~here~ore an object of the invention to
provide an aircraft ground proximity warning system with
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l enhanced warning capability when aircraft performance is
degraded.
- It is a further object of the invention to provide
an aircraEt ground proximity warning system with enhanced
warning capability near the ground. Specifically the
warning envelope of Modes l and 3 are extended to within
five feet of the ground. Radio altitude rate and barometric
altitude rate signals are combined to provide a computed
altitude rate signal that is accurate near the ground for
use as an input to Modes l and 3.
It is an additional object of the invention to
provide an aircraft ground proximity warning system with
flight path deviation warning utilizing a measure of flight
path and aircraft altitude. The measure of flight path can
be based on aircraft vertical velocity. A flight path
warning is provided whenever the aircraft flight path angle
is less than a predetermined angle and when the aircraft is
below a predetermined altitude.
It is still a further object of the inven-tion to
provide a pitch warning system for generating a warning when
aircraft pitch is below a predetermined value after
rotation. The pitch warning system can utilize angle of
attack for pitch measurement.
It is another object of -the invention to provide
an aircraft ground proximity warning system with an output
indicating that additional aircraft performance is
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l available. Angle of attack is compared ~o stall angle of
attack to generate an indication that angle of attack should
be increased. A pilot indication to apply additional thrust
can also be provided.
Brief Description of the Drawi~s
Fig. l is a functional block diagram of a ground
proximity warning system with angle of attack and stall
warning margin inputs;
Fig. 2 is a graphical representation of a Mode l
warning envelope;
Fig. 3 is a graphical representation of a Mode 3
warning envelope;
Fig. 4 is a graphical representation of a flight
path warning envelope;
Fig. 5 is a graphical representation of a takeoff
angle of attack warning envelope;
Fig. 6 is a functional block diagram of the flight
path warning logic portion of the warning system of Fig. l,
used during takeoff;
Fig. 7 is a functional illustration of the
operation of the stall margin portion of the logic of Fig.
6; and
Fig. ~ is a functional block diagram of the flight
path warning logic portion of the warning system of Fig. l,
used during approach.
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1 Detailed _ escriptlon of the Invention
Fig. 1 illustrates in generalized block diagram
form the preferred embodiment of the invention. A source of
slgnals or data source for the warning system is indicated
by a block 10. The signals provided by the data source 10
include: radio altitude hR, barometric alti-tude hB, angle
of attack ~ , stall margin ~- ~s, vertical accelerometer an,
airspeed V, gear and flap position and glide`slope G/S.
Typically in modern digital commercial aircraft these
signals are available from the aircraEt digital data bus or
flight management system. On older aircraft, these signals
are normally available from individual instruments.
As shown in Fig. 1 the warning system has four
separate warning modes. these modes include a Mode 1
excessive descent rate warning mode, a Mode 3 negative climb
after takeoff warning mode, a flight path warning mode and a
takeoff angle of attack warning mode. Although only four
warning modes are described, it will be understood that the
system could include other warning modes such as those
disclosed in U.S. Patent 3,946,358.
A graphical representation of an improved Mode 1
warning envelope is provided in Fig. 2. This warning
envelope is similar to that shown in U.S. Patent 4,060,793
with the primary exception that the radio altitude cut off
has been moved down to five feet of radio altitude as
opposed to 50 feet in the prior art system. By lowering the
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1 warning boundary to five feet, warnings can be generated
much closer to the ground which can be useful in, for
example, wind shear situations on an approach to landing.
Lowering the floor of Mode 1 is made possible by producing a
computed altitude rate signal hC which overcomes error
sources in the barometric rate signal close to the ground.
As shown in Fig. 1 the Mode 1 warning envelope of
Fig. 2 is produced by applying the radio altitude signal hR
on line 12 and a barometric rate signal fiB on line 14 to a
computed altitude circuit 16. The barometric rate signal is
obtained from a differentiating circuit 18 which receives a
barometric altitude signal hB from signal source 10 over
line 20. The computed altitude circuit 16 which will be
described in detail in connection with Fig. 6 combines the
radio altitude rate signal XR with the barometric altitude
rate signal to produce the computed altitude rate signal
hc. This signal includes proportionally more radio altitude
rate the closer the aircraft is to the ground thereby
tending to eliminate error sources in the barometric rate
signals due to ground effects. Mode 1 warning initiated
signals are produced on a line 22 by a warning circuit 24
which receives the computed altitude rate signal over line
26 and the radio altitude signal on line 12. Suitable means
for implementing the operation of circuit 24 is disclosed in
U.S. Patent 4,060,793. A warning logic circuit 28 receives
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1 the Mode 1 initiated signal on line 22 and generates, where
appropriate, a voice warning on a cockpit speaker 30.
In a similar manner the effectiveness of Mode 3 is
enhanced by reducing the radio altitude cut off from 50 feet
to 5 feet as illustrated by the warning envelope of Fig.
3. A warning mode logic circuit 32 receives the radio
altitude signal over line 12 and the computed altitude rate
signal hC over line 26 from the computed altitude rate
circuit 16. It is the accuracy of the computed altitude
rate signal that permits the Mode 3 warning of Fig. 3 to be
reduced to five feet of radio altitude and hence resulting
in a more responsive warning system. The logic circuit 32
operates in a conventional manner such as the systems
disclosed in U.S. Patents 3,947,810 or 4,319,218 to produce
warning initiate signals on line 40 when the aircraft
descends a predetermined amount of altitude after takeoff.
Accident analysis has shown that flight safety can
also be improved by giving a warning for inadequate flight
path angle ~ when the aircraft is close to the ground
either during takeoff or a landing approach. An
illustration of the preferred embodiment of a flight path
warning envelope for the takeoff phase of flight is provided
in Fig. 4. Here the cross-hatched portion to the right of
line 42 ir.dicates that a flight path warning will be
initiated for flight path angles less than 0.5 for radio
altitudes of 35 feet or greater.
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1 Wind shear can cause a sustained loss of
airspeed. With a loss of airspeed a loss of altitude may
follow and as such it is desired that the aircraft be in a
climb attitude in order to prevent or minimize any dangerous
loss of altitude near the ground. Therefore, under
conditions of a negative airspeed rate, the warning curve of
Fig. 4 is shifted to the left as indicated by the dashed
line ~4 so that a warning is given earlier at a greater
flight path angle.
The flight path warning logic is represented by a
logic block 46 or Fig. 1 the details of which are shown in
Fig. 6. Inputs to the logic block 46 include radio altitude
on line 12, computed altitude rate 26 and airspeed rate on
line 48. Airspeed V is obtained from data source 10 and
applied over line 50 to a differentiator circuit 52.
Referring to Fig. 6 the computed altitude circuit
16 produces the computed altitude rate signal hC on line 26
by blending the barometric rate signal hB with a radio rate
signal hR below a predetermined radio altitude hRMAX. The
radio altitude signal is differentiated by a differentiator
circuit 54 and applied to a first multiplier circuit 56. A
multiplier K having values from 0 to 1.0 as a function of
radio altitude is produced by a function generator circuit
58. The value K-l produced by a summing junction 60 is also
applied to the first multiplier 56 resulting in the value
(l-K) ~R on a plus terminal of a summing junction 62. A
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1 second input to the summing junction 62 is the quantity K hB
produced by a second multiplier circuit 64. The second
multiplier circuit 64 receives the barometric rate signal
over line 14 and the multiplier K from function generator
circuit 58. In operation the circuit 16 will produce a
computed altitude rate signal that at hRMIN and below is
equal to radio altitude rate and at hRMAx is equal to
barometric altitude rate.
In addition the computed altitude circuit 16
includes a detector circuit 66 responsive to radio altitude
on line 14 to start a timer circuit 68 at lift off. The
timer 68 inputs to a limiter circuit 70 that outputs a
signal over a line 72 to the function generator circuit 58
that has the effect of making the value of K equal to 1.0 a
predetermined time after the aircraft lifts off the runway.
As discussed above the warning curve of Fig. 4 is
shifted to the left as a function of a decreasing rate of
airspeed. A function circuit 78 in Fig. 6 responds to the
airspeed rate signal on line 48 and serves by means of line
~ 80 to bias the output of logic circuit 46 to provide a
warning at greater flight path angles as a function of
increasing negative airspeed rate.
With respect to the flight path warning, once a
warning has been generated by the circuit 46 indicating that
the aircraft may have an unsafe flight path, it is
considered desirable to provide the aircrew with guidance as
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1 to what action will tend to maximize the safety of the
aircraft. Logic which can form a portion oE the warning
logic 28 of fig. l is shown in Fig. 6. A stall margin
signal ~ - ~s from the signal source lO is applied over a
- 5 line 82 to a comparator circuit 84. If the stall margin
signal indicates that the aircraft's angle of attack ~ is
within a predetermined amount of the stick shaker angle of
attack ~s, the comparator 84 will apply a logic signal
over a line 86 to an OR gate 88. A positive logic output
from gate 88 will cause an aural warning such as "add
thrust" to be generated by the warning logic 28. The flight
path logic 46 will put out a signal suggesting that the
pitch attitude or flight path angle of the aircraft is too
low. Normally the preferred aural warning will be "nose up"
or "pitch up" to indicate that the aircraft pitch attitude
should be increased due the proximity to the ground.
However, if the stall margin logic signal on line 86
indicates that the aircraft attitude is already close to
stall, a "pitch up" type advisory may be inappropriate.
Therefore, and AND gate 90 serves to inhibit the "pitch up"
warning when the aircraft is approaching stall. In the
preferred embodiment of the invention, the "add thrust"
advisory will always be generated since added thrust should
always be considered by the aircrew when in difficulty close
to the ground. Note that the circuit of Fig. 6 includés a
circuit 92, a limiter 94 and a summing junction 96 to
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1 provide a stall margin rate lead term to the comparator
84. This will speed the response of the circuit 84 if the
rate of increase of angle of attack should indicate a rapid
pitch up of the aircraft. Operation of this circuit is
illustrated by Fig. 7.
Flight path logic 46 for use when the aircraft is
on approach is illustrated in Fig. 8. When on approach the
function generator 46 of Fig. 1 will operate somewhat
differently from the function generator of Fig. 6
illustrated by the warning envelope of Fig. 4. Therefore,
the function generator of Fig. 8 will be indicated by 46'.
Flight path angle ~ which is defined as the angle that the
direction of travel of the aircraft makes with the horizon,
can be approximated by vertical speed such as hB or hc.
lS Computed altitude rate was used in the circuit of Fig. 6. A
more accurate approximation of flight path is vertical speed
divided by airspeed V. This approach is illustrated in Fig.
8 where a divider circuit 98 divides the computed 20
altitude rate on line 26 by the airspeed on line 50. Thi5
provides a flight path angle input over line 100 to the
warning envelope function generator 46'.
Since the logic of Fig. 8 is used when the
aircraft is on approach the normal flight path angle will be
negative. The warning envelope shown in 46' of Fig. 8 will
provide a first warning initiate signal on line 102 and a
second on line 104 when flight path exceeds a second
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l amount. The first signal on line 102 applied to an AND gate
106 will cause a "nose up" or "pitch up" aural warning. As
described in connection with Fig. 6 the approaching stall
margin signal on line 86 can inhibit the "pitch up" aural
warning via AND gate 106. A pull up warning on an AND gate
108 can also be inhibited by a logic signal on line 86.
A glide slope signal G/S input from the signal
source 10 of Fig. 1 on a line 110 can provide additional
warning logic. This signal, input through a function
generator circuit 112, can be used to inhibit the output of
gate 106 when the aircraft is not below the glide slope
criteria of function generator 112. The glide slope signal
on line 110 can also be used to modify the bias applied by
the function generator 78 to the warning envelope 46' over
line 80.
An additional "add thrust" warning can be
generated by OR gate 88 by coming through an AND gate 113
the airspeed rate signal on line 80 and the below glideslope
signal from function generator 112.
The use of the logic of Fig~ 6 or Fig. 8 for
flight path warning depends on the phase of flight. If the
aircraft is in a takeoff or go around phase of operation,
the circuits of Fig. 6 is used. If the aircraft is in an
approach phase, the circuit of Fig. 8 is used. In the
preferred embodiment a takeoff logic circuit ll~ is used to
select the appropriate flight path warning circuit. Logic
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1 for such a circuit is disclosed in U.S. Patents 3,947,810
and 4,319,218. A phase of flight signal is transmitted from
the takeoff logic 114 over a line 116 to circuit 46.
Under certain circumstances it may be desirable to
give a warning of potentially insufficient angle of
attack. The criteria for such a warning is illustrated in
Fig. 5. Durin~ takeoff, once the aircraft has rotated to a
predetermined angle of attack for, example 2, any decrease
in angle of attack will result in a warning. Logic for
generating such a warning is indicated by a block 118 in
Fig. 1. Duration of this warning mode can be a function of
time from lift off or radio altitude or barometric altitude.
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This application is a divisional o-f Canadian Patent
Application, Serial No. 481,522, filed May 14, 1985.
While the invention has been described with reference to
preferred embodiments, the invention is not so limited.
Many modifications and variations will now occur to a person
skilled in the art. For a definition of the invention,
reference is made to the following claims.
