Wind shear detector head-up display system

An aircraft performance instrument which provides a first indicator which represents the line of level flight of the aircraft and a second indicator representative of the maximum permissible angle of attack of the aircraft.

TECHNICAL FIELD 
This invention relates to the field of aircraft instruments and, in 
particular, to a head-up display and instrument which provides guidance to 
the aircraft's pilot so as to enable the pilot to fly the aircraft safely 
during degraded performance such as during a wind shear condition. 
BACKGROUND OF THE INVENTION 
Various wind shear detection systems are known. Among these systems are 
ground based systems, such as those that utilize a plurality of wind speed 
and direction measuring stations placed about an airport in conjunction 
with a system for analyzing the magnitude and direction of the wind at the 
various stations to provide an indication of a wind shear condition. U.S. 
Pat. Nos. 4,058,010 and 4,152,933 are examples. Other ground based systems 
utilize Doppler radars located near the airport (see U.S. Pat. No. 
3,427,581). 
Airborne systems are also known. Among these systems are systems that 
compare air mass derived parameters, such as air speed, with ground speed 
derived from a radar system. In the event of a rapid change in air speed 
relative to ground speed, a wind shear condition is indicated. Other 
systems compare air mass derived signals with inertially derived signals 
to generate a signal representative of wind shear when the rate of change 
of inertially derived parameters varies from the rate of change of air 
mass derived parameters by a predetermined amount. Two such systems are 
disclosed in U.S. Nos. 4,012,713 and 4,079,905. Both of these patents 
disclose systems that compare a longitudinal accelerometer signal that has 
been corrected for the effect of gravity with an air speed rate signal and 
provide a wind shear warning signal when the difference between the 
accelerometer derived signal and the air speed derived signal exceeds a 
predetermined amount. The '905 patent also takes into account a downdraft 
drift angle that is a function of vertical acceleration and air speed. 
Still other systems monitor the rate of change of deviation from a glide 
slope beam or an ILS beam to provide a signal representative of wind 
shear. 
While all of these systems do provide some indication of wind shear, the 
ground based systems are responsive only to conditions in the vicinity 
where the transducers are placed, and are not responsive to dangerous 
types of wind shear such as microbursts which form and dissipate rapidly. 
A microburst is an intense downdraft of cool air that, in some cases, 
drops 3,000 feet in two minutes. Such systems do not necessarily provide a 
timely warning to the pilot. 
While airborne wind shear protection systems are more responsive to 
conditions in the vicinity of the aircraft than are ground based systems, 
many of them require Doppler Radar, an Inertial Navigation System, glide 
slope signals, or other signals that are not available on older or small 
aircraft. In addition, it is desirable to develop not only a warning of 
degraded aircraft performance, but also timely guidance to the pilot on 
how to safely fly the aircraft during this degraded performance condition, 
such as a wind shear condition. 
There also have been teachings as to how aircraft performance is affected 
by wind shear (i.e., U.S. Pat. Nos. 4,043,194; 4,129,275; 4,281,383; 
4,342,912 and 4,336,606). These teachings involve systems which are often 
difficult to implement in an aircraft cabin environment, may require the 
use of special sensors, or do not provide usable information in a timely 
manner. 
Some systems have been devised which not only combine the detection of wind 
shear but also guidance information to the pilot. In U.S. Pat. No. 
4,189,777, air speed 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 is reflected in 
U.S. Pat. No. 4,347,572. There, angle of attack, stick-shaker value, 
vertical speed, air speed, flap position and thrust are used to provide 
climb-out guidance on a pilot flight director display by indicating an 
appropriate pitch angle. U.S. Pat. No. 4,040,005 teaches a servo driver 
composit situation analyzer. There, a complex mechanical instrument is 
disclosed which, in one embodiment, displays angle of attack flight path 
angle relative to an aircraft reference and a runway. The suggestion is 
also made that a CRT display may be used in lieu of the mechanical 
display. However, the displayed information amounts to a presentation of 
deviations from desired values of angle of attack and desired flight path. 
More importantly, since it depends on the pilot to insert a desired value 
for the angle of attack, it has the potential for giving misleading 
information during rapidly changing flight conditions. The complexity of 
the display also makes it likely that the pilot will not use it during 
emergency situations wnen the aircraft is in sight of the ground. 
In view of some recent aircraft accidents which have been attributed to 
wind shear and in view of maintaining the public's confidence in 
commercial aircraft, an aircraft instrument which provides timely guidance 
information to the pilot of tne aircraft so as to enable him to correctly 
guide the aircraft in a wind shear condition would be an instrument that 
would satisfy a clear long-felt need. 
The fact that so many instrument systems and guidance systems have been 
proposed, and no one system has yet found acceptance by the industry, is 
an indication that these prior systems have been unsatisfactory and that 
the problem of providing guidance information to an aircraft pilot during 
a wind shear condition is one which has, so far, gone unsolved by the best 
minds in the industry. Clearly, a new approach to the problem is needed. 
SUMMARY OF THE INVENTION 
In accordance with one embodiment of the present invention, an aircraft 
head-up display is disclosed which provides, on a combiner, an aircraft 
symbol; a horizon line; a first symbol which is representative of the line 
of level flight; and a second symbol which is representative of the 
maximum permissible angle of attack. In one specific embodiment, the 
aircraft is provided witn a Stall Warning System which provides a signal 
representative of the stick-shaker angle of attack and a signal 
representative of the actual angle of attack of the aircraft, thereby 
providing a representation of the stall margin of tne aircraft. In that 
embodiment, the second symbol on the combiner is positioned by a signal 
formed by combining the stall margin signal and a pitch signal 
representative of the pitch of the aircraft. The pitch signal input 
improves flyability and damping. In that embodiment, the first symbol is 
positioned by a signal representative of the ratio of vertical speed to 
true air speed. Preferably, the vertical speed signal is the signal 
obtained by combining vertical speed and normal acceleration in a 
complementary filter. 
In another embodiment, an aircraft instrument is disclosed which comprises 
a scale index, a display of the aircraft flight path angle relative to 
said scale index, and a display of tne aircraft stall margin relative to 
said scale index. Preferably, the scale index is disposed between the 
stall margin display and the flight path angle display and the instrument 
is in the form of a circular dial gage with a pointer indicating flight 
path angle and a peripheral concentric display of stall margin. A digital 
vertical speed display may also be included on the dial face. 
As will be apparent from the discussion which follows, the two specific 
embodiments just described not only provide wind shear warning, but also a 
display which enables the pilot to recognize a dangerous wind shear 
condition, or other aircraft performance problem, and assist the pilot in 
safely flying the aircraft through such a condition. 
One advantage of using a head-up display (HUD) is that it is especially 
well suited as a display element for providing aircraft guidance 
information inasmuch as it allows the pilot to simultaneously view the 
outside world and the super-imposed HUD guidance symbology. The pilots 
know that wind shear is especially dangerous when the aircraft is at a low 
altitude, such as during a landing approach or during take-off. At these 
times, the pilot is generally viewing the outside world and would be most 
appreciative of any instrument that would give him guidance and wind shear 
warning information without having to move his head towards a conventional 
array of instruments. However, a HUD is generally more expensive than a 
conventional instrument; therefore, a less expensive conventional dial 
instrument that provides the same information would still prove to be a 
valuable contribution to flight safety. 
Numerous other advantages will become apparent from the following detailed 
description of the invention and its various embodiments, from the claims, 
and from the accompanying drawings.

DETAILED DESCRIPTION OF THE INVENTION 
While this invention is susceptible of embodiment in many different forms, 
there is shown in the drawings and will herein be described in detail, 
several preferred embodiments of the invention. It should be understood, 
nowever, that the present disclosure is to be considered an 
exemplification of the principles of the invention and that it is not 
intended to limit the invention to the specific embodiments illustrated. 
To better explain the unique aspects of the invention, the principle of 
operation will first be described, and then specific embodiments will be 
explained which employ the principles of the invention. 
Principle of Operation 
Referring to FIG. 1A, there is shown a pair of horizontal and vertical 
coordinates designated by the reference numberals 10 and 12, respectively. 
A representation of an aircraft 14 is also shown in FIG. 1A, and a pair of 
vectors 16 and 18 represent the longitudinal and normal axes of the body 
of the aircraft, with the vector 16 passing through the longitudinal 
center line of the aircraft and the vector 18 being perpendicular to the 
vector 16. A dashed line 20 represents the flight path of the aircraft 14, 
and in the illustration of FIG. 1A represents a descending flight path. 
The angle between the horizontal axis 10 and the longitudinal axis 16 of 
the aircraft is defined as the pitch angle and represented by the symbol 
".theta.". The angle between the flight path 20 and the horizontal 
reference 10 is defined as the flight path angle and is represented by the 
symbol ".gamma.". The angle between the longitudinal axis 16 of the 
aircraft 14 and the flight path 20 is known as the angle of attack of the 
aircraft and is represented by the symbol ".alpha.". The flight path angle 
".gamma." is negative for a descent and positive for ascent. Thus the 
pitch angle ".theta." is equal to the flight path angle ".gamma." plus the 
angle of attack ".alpha.". 
It is also important to understand the various accelerations and 
accelerometer signals affecting the aircraft. One such signal is a 
longitudinal accelerometer signal A.sub.L which is the signal obtained 
from an accelerometer mounted parallel to the longitudinal axis 16 of the 
aircraft 14. The signal from the longitudinal accelerometer is a function 
of the longitudinal acceleration of the aircraft 14, and because of the 
influence of gravity, the pitch angle. Another such signal is the normal 
accelerometer signal A.sub.N which is the signal obtained from an 
accelerometer positioned parallel to the normal axis 18 of the aircraft 
14. The normal accelerometer signal is a function of acceleration along 
the normal axis of the aircraft 14 as well as gravity and pitch angle. 
Horizontal acceleration A.sub.H is a signal representative of acceleration 
along the horizontal axis 10. An accelerometer mounted parallel to the 
horizontal axis 10 would provide the horizontal acceleration signal 
A.sub.H. Finally, a vertical acceleration signal A.sub.V is a signal 
representative of acceleration along the vertical axis 12. An 
accelerometer mounted parallel to the vertical axis 12 would provide a 
signal representative of the sum of any vertical acceleration and the 
effects of gravity, "g" or 32.2 feet/second.sup.2. 
Recapitulating the various accelerations and angles to be considered are as 
follows: 
A.sub.N =N normal accelerometer signal 
A.sub.L =longitudinal accelerometer signal 
A.sub.H =horizontal acceleration 
A.sub.V =vertical acceleration 
g=gravity =32.2 feesecond 
.alpha.=angle of attack (AOA) 
.gamma. flight path angle 
.theta.=pitch angle=.alpha.+.gamma. 
In determining some of the parameters, other parameters will be required. 
These include: 
h=barometric altitude rate 
V=air speed 
and for small angles: 
.gamma.=h/v 
where ".gamma." is in radians. 
A particular severe performance problem exists when the aircraft flies 
through a microburst during take-off or when the aircraft encounters a 
wind shear condition during its approach that is severe enough to require 
a "go around". Referring to FIG. 1B, a microburst exhibits both downdrafts 
as well as strong horizontal wind shears. Downdrafts, coupled with 
decreasing headwind or increasing tailwind shears, will cause the aircraft 
14 to simultaneously lose air speed 22 and climb gradient. Even if the 
pilot selects maximum available thrust, the aircraft might not be able to 
out accelerate the horizontal shear and maintain an adequate climb 
gradient. Tne downdraft causes a decrease in aircraft angle of attack for 
a given pitch attitude; in other words, the pilot has to pitch-up to an 
abnormally high pitch attitude just to maintain an angle-of-attack 
equivalent to a non-shear condition. 
Thus, under such abnormal conditions, the pilot, after having selected 
maximum available thrust, has to control the airplane between the 
following two constraints: 
(1) He has to try to increase angle of attack to maintain a positive climb 
rate to avoid terrain contact 24. 
(2) He has to prevent the aircraft from stalling 26. 
Clearly, during such conditions, the pilot needs guidance that is not 
ambiguous, and that is presented in such a manner as to show the safety 
margins available. 
Turning now to FIG. 2A, a block diagram of an elementary head-up display 
(HUD) is illustrated. Bascially, such a display is provided by projecting 
tne output of a CRT through a set of optics on a semi-transparent screen 
or combiner. The combiner is disposed between the aircraft pilot's eye and 
an aim point exterior to the aircraft. The CRT displays images, including 
appropriate symbols, as directed by appropriate control circuitry. The 
apparatus of FIG. 2A is well known to those skilled in the art and is 
otherwise discussed in such patents as U.S. Pat. Nos. 3,686,626, 
3,851,303, 3,967,799, and 4,390,950. One commercial embodiment is the 
MD-80 made by the Sundstrand Corporation, Sundstrand Data Control, Inc. in 
Redmond, Washington. 
Turning now to FIG. 2B, the image formed on the combiner of FIG. 2A is 
illustrated. Specifically, the display includes an aircraft symbol 26, an 
aim symbol or "aimdot" 28, a horizon line 30 and a pitch ladder 32. The 
aircraft symbol 26 symbolizes an aft view of an aircraft controlled by the 
pilot to correspond to motions of his aircraft. During an approach, the 
pilot "flies" the aircraft symbol 26 to keep it centered on the aimdot 28. 
The display may also be provided with an indicator 34 which moves relative 
to a horizontal scale 36 to depict the aircraft's heading, airspeed (i.e., 
138 kts) and digital radio altitude (i.e., 200 ft.) Thus, the display 
shown in FIG. 2B is conventional. 
In the case of the MD-80 HUD, the aircraft symbol 26 is displayed fixed in 
a location representing the roll axis of the aircraft (i.e., an axis 
perpendicular to the plane of FIG. 2B). The horizon line 30 and the pitch 
ladder 32 are displayed "conformal" to the external or real world. This 
means that the horizon line 30 overlays the true horizon independent of 
the aircraft pitch or roll attitude. In other words, the pitch ladder 32 
and the horizon line 30 are pitch and roll stablized and indicate pitch 
and roll when compared to the fixed airplane symbol. Finally, the aimdot 
28 moves relative to the airplane symbol and represents a "command" from 
the flight director portion of the HUD. 
In FIG. 2C, a HUD combiner is depicted which has two additional symbols 
imposed thereon in accordance with the present invention. Specifically, 
there is illustrated a single dashed line 38 representing the level flight 
line and a double dashed line 40 representing the maximum angle of attack. 
Thus, these two symbols form a "take-off window" inside of which the 
aircraft symbol 26 is located under normal flight conditions. The take-off 
window is roll stabilized about the aircraft roll axis like the horizon 
line 30 and the pitch ladder 32. The displacement or location 42 of the 
upper dashed line 40 or upper window limit, relative to the aircraft 
symbol 26 is representative of or indicates the stall margin of the 
aircraft (i.e., stick-shaker angle of attack minus actual angle of 
attack). If the aircraft stall margin is large, the upper dashed line 40 
might be out of view, as specifically shown in FIG. 2C. The relative 
location of the single dashed line 38, compared to the aircraft symbol 26, 
is representative of the flight path angle ".gamma." of the aircraft 44. 
As long as the aircraft is climbing, the aircraft symbol 26 is above the 
lower limit 38. Therefore, the "size" of the take-off window is a measure 
of the take-off performance capability of the aircraft. More importantly, 
a take-off performance problem is immediately evident by a shrinking 
window size and/or the aircraft symbol being outside of the display 
window. 
Another embodiment of the present invention implementing the take-off 
window concept is shown in FIG. 2D. In this implementation, the head-up 
display of the model MD-80 HUD has been modified by letting the aircraft 
symbol 26 represent flight path angle relative to the real world and the 
horizon line 30. Thus, the horizon line 30 indicates the level flight line 
in relation to the aircraft symbol 26. As long as the aircraft is 
climbing, the aircraft symbol 26 is above the horizon line 30. The 
take-off window is still displayed relative to the aircraft symbol and is 
essentially the same, except that its lower limit can now be replaced by a 
horizon line and no separate symbol is required. 
The window concept can be accentuated by connecting the two horizontal 
limit lines 38 and 40 with vertical lines 46 and 48 and by laterally 
moving the window relative to the aircraft symbol 26 as a function of 
Heading Error "E" (i.e. desired heading minus actual heading). Similarly, 
an aircraft symbol 26 may be manipulated to show rotation about the 
aircraft's roll axis (i.e., see FIG. 3B). In that case, the window would 
be "roll stabilized". 
Scenarios showing various flight performance situations are shown in FIG's. 
3A 3C, 3D and 3E. FIG. 3A corresponds to FIG. 2D with a separate line 38 
used in lieu of the horizon line 30. Specifically, FIG. 3C illustrates the 
display presented to a pilot in the case of marginal climb performance. 
FIG. 3D illustrates shrinking stall margin, and FIG. 3E illlustrates an 
emergency situation, whereby the pilot has to use his judgment to trade 
flight path angle for stall margin. It should be evident that the take-off 
window concept not only indicates performance problems due to wind shear, 
but also problems due to inadequate thrust, excessive drag or improper 
flight path control (i.e., sinking after take-off). 
Control System 
A block diagram of a control system for implementing the head-up display 
just described is shown in FIG. 4. Specifically, the position of the upper 
window limit 40 relative to the aircraft symbol 26 is essentially 
controlled by "stall margin", a stall margin signal 50, which is 
considered to be the difference between the "stick-shaker angle of attack" 
and actual aircraft angle of attack. Modern aircraft are provided with 
stall warning systems which can provide such an output signal. The 
stick-shaker is an impending stall warning device and, in the case of jet 
transport aircraft, is automatically activated when the angle of attack is 
at a "never exceed" value corresponding to an airspeed about 10% above the 
stall value. To reduce the effects of turbulence, the stall margin signal 
50 is filtered in a standard first order lag type filter 52. A rate signal 
54, generated by a differentiator circuit 56, is fed to a limiter 58 to 
limit the rate signals generated by the differentiator 56 to rates typical 
for the motion of the aircraft in which the system is installed. The 
output signal 60 from the limiter 58 is combined with the unfiltered stall 
margin signal 50 and both are fed to the turbulence filter 52 to reduce 
its overall lag effect without significantly reducing its turbulence 
filtering function. To improve flyability and damping, a pitch signal 62 
from a vertical gyro is put through a wash-out filter 64 and then 
subtracted from the output of the turbulence filter 52. 
As long as the aircraft angle of attack is below the stick-shaker 
activation point, the stall margin signal 50 is positive and the upper 
window limit 40 is above the aircraft symbol 26 (see FIG. 3A). As the 
aircraft angle of attack reaches the stick-shaker activation point, the 
upper window limit 40 will coincide with the aircraft symbol 26 (see FIG. 
3E). Under normal conditions, the take-off window acts as a situation 
display. In emergency situations, however, the pilot can use the upper 
window limit 40 as a guidance symbol to fly the aircraft at its maximum 
allowable angle of attack for a limited amount of time. 
The control law for the lower window limit 38 is flight path angle 
".gamma.". As long as the flight path of the aircraft is positive (i.e., 
climbing) the aircraft symbol 26 is above the lower window limit 38 (see 
FIG. 3A). The preferred signal source for the flight path angle is the 
aircraft's Inertial Reference System (IRS). For aircraft which do not have 
an IRS, the flight path angle may be computed from a vertical speed signal 
66 and true air speed 68. A standard complementary filter 70 is employed 
to filter a vertical speed signal 66 obtained from the aircraft's air data 
computer or from the rate of change of barometric altitude and to 
complement it with a signal representative of normal acceleration A.sub.N 
to restore responsiveness. 
Relatively recent wind shear related accidents have indicated that pilots 
were reluctant to counteract the wind shear by increasing the angle of 
attack to an optimum level. A downdraft, coupled with a decreasing head 
wind shear, decreases the angle of attack and air speed simultaneously 
(see FIG. 1B). This abnormally high pitch attitude is uncomfortable to the 
pilot especially when flying close to the ground and, therefore, there is 
reasonable anxiety that the aircraft will reach a stall condition. The HUD 
just described should reduce this anxiety factor inasmuch as it keeps the 
pilot simultaneously aware of his stall margin and flight path angle while 
allowing the pilot to view the outside world and manuever free of ground 
obstructions. To counteract the decreased angle of attack, a pitch 
attitude higher than normal is required. Under some critical wind shear 
situations, the aircraft cannot accelerate to make up the air speed loss 
and to simultaneously climb to counteract the downdraft. 
FIG. 5 illustrates another embodiment of the invention. Here, instead of 
using a head-up display, a modified vertical speed indicator is presented. 
Specifically, the indicator 80 displays flight path angle ".gamma." 
instead of vertical speed 88 alone. A moving angle of attack index 82 on 
the periphery of the indicator 80 provides a display of stall margin 42 
and optimum angle of attack 90 guidance when read relative to the flight 
path angle indicating needle 86. Such an instrument would show both 
constraints (see FIG. 1B) that the pilot faces when flying through a 
severe wind shear. The flight path angle index 86 indicates whether the 
aircraft maintains a positive climb gradient and the angle of attack index 
82 shows how close the aircraft is to being stalled or the optimum angle 
of attack 90. Preferably, for further guidance, the instrument of FIG. 5 
may be provided with a readout 88 in digital form of the aircraft's 
vertical speed. 
From the foregoing, it will be observed that numerous variations and 
modifications may be effected without departing from the spirit and scope 
of the novel concept of the invention. For example, a digital voice 
synthesizer may be included or one or more warning lights may be provided 
to alert the pilot that any one of the window limits 38 and 40 have 
decreased to some preselected value. Similarly, further enhancements or 
display information may be added to the combiner screen to provide 
additional guidance to the pilot of the aircraft. It should be understood, 
however, that no limitation with respect to the specific apparatus 
illustrated herein is included or should be inferred. It is, of course, 
intended to cover by dependent claims all such modifications as fall 
within the scope of the claims.