WO1992004698A1 - Improved method and apparatus for detecting weather related events and the like and devices for use therewith - Google Patents

Abstract

Method and apparatus for detecting weather related events and other temperature related phenomena including (a) a wobble method of detecting temperature differences (ΔT); (b) an infrared spectrometer (116) for implementing the wobble method; (c) various hazard index formulae for detecting the presence of low level wind shear (LLWS); (d) techniques for detecting LLWS and other temperature dependent phenomena where the detecting device (81) is relatively stationary with respect to the event being monitored including runway scans where the detecting means is mounted on an aircraft or landbased; (e) scanning systems for searching for phenomena such as the jet stream; (f) lapse rate correction (212); and (g) look distance calculations; and (h) utilization of unique wavelengths for clear air turbulence (CAT) detection (242).

Description

IMPROVED METHOD AND APPARATUS FOR DETECTING

WEATHER RELATED EVENTS AND THE LIKE AND

DEVICES FOR USE THEREWITH

FIELD OF THE INVENTION

The present invention relates to systems and devices for the detection of weather conditions such as low level wind shear in airborne or stationary aircraft. More particularly, the invention is directed to a system and apparatus by which a wide range of weather conditions can be accurately detected, in various installations including aircraft.

BACKGROUND AND SUMMARY OF THE INVENTION

As described in the above referenced co-pending application, low-level wind shear (LLWS) poses a major safety hazard for commercial and general aviation where aircraft are most vulnerable during landing and take¬ off. Furthermore, that application describes how an advance warning of LLWS is achievable. In particular, in a remote predictive mode, a temperature-based hazard index is calculated to achieve a warning factor on the basis of such data that is obtainable on board the aircraft as altitude, air speed and air temperature, both locally and remotely. For measuring outside air temperatures at a distance from the aircraft, infrared (IR) temperature measurements are utilized.

Various detection systems related to the detection of weather related events are known and are exemplified by U.S. Patents 3,427,581; 3,696,670; 3,498,132 3,537,136; 3,856,402; 3,926,522; 3,935,460; 3,938,386 4,012,713; 4,043,194; 4,079,905; 4,189,777; 4,250,746 4,282,527; 4,319,219; 4,342,912; 4,346,595; and 4,589,070.

It is a primary object of the present invention to provide a system and devices by which a warning of the existence of low-level wind shear can be provided with increased accuracy and reliability, especially in an aircraft-borne environment.

It is a further object of the present invention that the improved low-level wind shear warning system and devices be capable of distinguishing between various different types of weather situations,, including closely spaced, successive weather events.

Another object of the present invention is to provide temperature-based LLWS warning methods and systems which are able to utilize IR sensing and data processing technology with higher effectiveness than heretofore possible.

A still further object of the present invention is to provide a weather related event detection system which has applicability not only with respect to low level wind shear detection but also with respect to clear air turbulence detection, jet stream detection, volcanic ash detection, aircraft engine measurement, detection of leading or following aircraft, lapse rate correction, look distance calculation, etc.

In accordance with another feature of the present invention, event detection information is provided by systems which are relatively stationary with respect to the event being detected where the systems include those disposed at stations at an airport or the like and those disposed on aircraft on the runway preparatory to take off.

In accordance with another aspect of the present invention, the manner in which temperature signals are produced is improved by a wobbling between the infrared temperature detection frequencies.

Another object of the present invention is the • provision of scanning means to facilitate a search for the jet stream, volcanic ash, etc.

These and other features, objects and advantages of the present invention will become more apparent from the following detailed description when viewed in conjunction with the accompanying drawings .

BRIEF DESCRIPTION OF THE DRAWINGS

Figure 1 is a representation of the variation in gain as a function of wavelength for air of a. fixed temperature;

Figure 2 depicts a section of the spectrum at three different target temperature values, showing the relationship between the power signal as a function of wavelength for the portion of the spectrum illustrated;

Figure 3 represents the A and B wavelength signal differences from a bow tie as a function of target temperature;

Figure 4 is a graph of calibrated ΔT measurements versus ΔT between target and bow tie;

Figure 5 is a diagrammatic cross-sectional side view of an illustrative spectroscopic infrared sensing device of the parent application and for use as an IR sensor of the present invention;

Figure 5A is a perspective view of a further, preferred spectroscopic infrared sensing device of the present invention;

Figures 5B and 5C are a respective end view and a cross-section (along the line A-A of Fig. 5B) view of an illustrative damper body for use in the IR sensor of Figure 5A in accordance with a further aspect of the invention;

Figure 5D is an end view of an illustrative fiducial plate for use with the IR sensor of Figure 5A;

Figure 5E is a cross-sectional view of an illustrative optics/electronics package in accordance with the invention;

Figure 6 is an overall system diagram;

Figure 7 is an overall flow diagram illustrating the general operation of the system of Figure 6;

Figure 8 is a flow diagram of a main loop sequence of the system operation depicted in Figure 7;

Figure 9 is a flow chart of the steps of the LLWS mode of operation of the Figure 8 sequence;

Figure 10 is a flow chart of the steps of the CAT mode of operation of the Figure 8 sequence;

Figure 10A is a diagrammatic illustration of a . flight profile;

Figure 11A is an illustrative diagram of the derivative of the output signal from a detector as it enters and exits a microburst;

Figure 11B illustrates the detector signal of Figure 11A;

Figure 11C illustrates the microburst of Figure 11A;

Figure 12 is a block diagram/flow diagram of an illustrative overall system in accordance with the invention ;

Figure 13 (A) diagrammatically illustrates an aircraft entering a cold front, Figure 13(B) diagrammatically illustrates the derivative of the detector signal corresponding to Figure 13 (A) , Figure 13 (C) diagrammatically illustrates a FIFO signal associated with either a cold or warm front, Figure 13 (D) illustrates an OAT signal associated with a warm front, Figure 13(E) illustrates an OAT signal associated with a young microburst, Figure 13(F) illustrates the derivative of the detector output signal associated with Figure 13 (E) , and Figure 13 (G) illustrates the detector output signal itself associated with Figure 13 (E) ;

Figure 14 illustrates a thermal index signal associated with UAL flight 236 of July 11, 1988;

Figure 15 is an illustrative block diagram/flow chart of an illustrative system for providing FIFO compensation; Figure 16 diagrammatically illustrates the compensated response effected by the circuitry of Figure 15 in response to the sensing of a front;

Figure 17 illustrates the compensated response in response to the sensing of a microburst;

Figures 18-23 and 24 are waveforms illustrating various conditions compensated by an alternative FIFO compensation system in accordance with the invention;

Figure 25 diagrammatically illustrates an observation cone of an illustrative detector in accordance with the invention;

Figure 26 is a graph illustrating how look distance may be obtained if lapse rate is known and vice versa;

Figure 27 is an illustrative flow chart illustrating the correction of measured temperature to compensate for the pitch of an aircraft;

Figure 28 is a graph which illustrates signal change as a temperature change is approached;

Figure 29 illustrates the detector signals associated with SAWS I and II signals;

Figure 30 is an illustrative block diagram illustrating circuitry for processing an illustrative SAWS II signal such as that obtained from the Runway Scan system;

Figure 31 illustratively illustrates the jet stream;

Figure 32 is a diagrammatic illustration of an up/down, rotate 360^* search/scan unit in accordance with the invention;

Figure 33 is an illustrative flow diagram of Runway Scan logic in accordance with the invention.

DETAILED DESCRIPTION OF THE PRE. ERRED EMBODIMENTS

As discussed above, the present invention relates to apparatus and to methods for determining, the existence of weather conditions including those which pose hazards to airborne aircraft utilizing an IR spectrometer and temperature-based hazard index criteria, as is the case in the above-mentioned, co- pending application. Moreover, the present invention is directed to specific temperature-based hazard index calculating algorithms, and to various techniques for obtaining the requisite temperature-based data together with methods and apparatus utilizing such temperature- based algorithms, to obtain increased accuracy and reliability, as well as to expanding the applications to which such methods and apparatus can be applied.

Considering first the manner in which outside air temperature data is obtained, it is noted that most IR spectrometers, such as those disclosed in U.S. Patent Nos. 4,427,306 and 4,342,912 (both of which are incorporated herein by reference) , use detectors which respond to changes in IR intensity, and their output is a function of the temperature difference between the observed object and the detector temperature. To provide the required changes in IR intensity, a chopper is used, so that the signal alternately represents a function of the difference in temperature between the detector and the observed object and a function of the difference in temperature between the detector and the chopper blade. However, in accordance with the present invention, it is not necessary to employ a chopper, althoughas will be described below, a chopper may be employed either by itself or in combination with the wobble techniques of the present invention.

To explain the manner in which the present invention avoids the need for a chopper and the use of classical chopper techniques, reference is made to Figures 1-4.

Figure 1 shows the gain versus wavelength plot for one of the grating spectrometers used in the AWAS program. The gain is a product of Planck's equation

and the gain factors of the various elements of the spectrometer, such as the blaze enhancement of the grating (shown as a separate line on the plot) , the transmission of the filters, and the sensitivity of the sensors. The gain of the S/N002 unit is shown as a dashed line. In Figure 1 the M numbers represent the order of the refraction from the grating.

Figure 2 shows the spectral scan at three different target temperatures, using a bow tie at 23 degrees C. The detector system output is measured in millivolts. For an example of the wobble operation we will consider wavelength A to be 13.75 microns, and wavelength B to be 14.98 microns. These wavelengths will have been selected for properties such as their gain magnitude and look distance.

By stepping or wobbling from wavelength A to wavelength B to A to B etc. in accordance with the present invention, in a nearly square-wave manner, the detector output will be a function of the optical power at wavelength A and the optical power at wavelength B that have transferred through the optics. Thus, in this manner, a time varying signal is produced that is a function of the power difference between A and B, instead of, as in the classical IR system, where the signal is a function of the power difference at a single wavelength A between the target and the chopper blade.

Figure 3 shows that when measured against the bow tie both of these wavelengths are relatively linear functions of the Celsius temperature. Multi¬ dimensional calibration is provided to cause the two wavelengths to provide near identical readings as a function of common target temperatures. When the constants from this calibration are applied in the data processing, the reference does not need to be the bow tie, but can be the other calibrated wavelength. By wobbling from one wavelength to the other rather than

SU33T.TUTΞ SHΞE" looking in sequence at two different temperatures with the same wavelength, the reading obtained is a measure of the difference in temperatures between the regions seen by the two wavelengths.

Figure 4 shows a laboratory validation in which a calibrated AWAS is looking successively at a temperature controlled bow tie with one wavelength and at a temperature controlled target with the other wavelength. The accuracy of this validation is shown over a range of 100 degrees.

Thus, in the present invention, a chopper is not necessary, nor is it necessary to cool the sensor. These factors are unique for an infrared spectrometer, and make it readily applicable to aircraft operations and for use in remote or restricted access areas where full automation is needed, and where minimum supplies are available.

Although no chopper and no cooling are involved in many other IR instruments, these are not precision spectrometers, such as that of the present invention. Also they do not use adjacent sets of spectral wavelengths to establish accurate temperatures.

The distance that light of wavelength B travels through the atmosphere before it is sorbed and re- emitted in a random direction, or is scattered, is significantly different from that of wavelength A. Thus, the measurement of A-B can give significant information. The A-B measurement, gives the measurement of a difference between distant and local temperatures. Thus, one can measure a function of distant temperature minus near temperature, or a delta temperature. This delta T, or difference in temperature, can be used in the hazard index of the above-mentioned co-pending application or as the basis of a number of possible instruments, both aircraft borne and land/sea based for measuring temperatures and temperature changes, and interpreting them for various purposes although it is to be understood that A-B measurements may also be used for many purposes including the foregoing purposes.

There are several schemes by which these differences can be generated. One can sequentially or simultaneously detect the near and far temperatures and then subtract one from the other where the far temperature is detected with a conventional IR spectrometer of the chopper type, for example, and the near temperature is detected by conventional temperature measuring means provided on aircraft to measure outside air temperature (OAT) . However, these methods involve two complete three part noise systems to be squelched. The wobble method of the present invention has only one detector and two instead of three target noise sources per measurement. Thus, this method is less noisy than subtracting two conventional measurements.

To obtain the wobble readings of the present invention, an infrared spectrometer of the type disclosed in the above-mentioned co-pending application may be used. This IR spectrometer is shown in Figure 5 and the following description thereof, for simplicity and consistency, utilizes the same reference numerals found in that disclosure. In particular, the multi- spectral IR spectrometer is designated, as a whole, by reference numeral 11 and receives an infrared ray signal 70 from either a remote or local IR source through a window 71 of the spectrometer. After entering through window 71, the IR signal is reflected at a rotatable planar mirror 72 to a converging primary mirror 73. From primary mirror 73, the IR signal 70 is caused to converge to a harmonic filtered slit 74, after which it is reflected at a diverging secondary mirror 75. As a result, the IR signal is received in a parallel ray configuration at a rotatable planar surface assembly 76 which has two receiving surfaces 78, 79 disposed on opposite sides of a rotatable receiving block 77. The first of these surfaces 78 is a planar mirror, and the other surface 79 is a diffraction grating.

On the basis of program commands from a data processing means, such as a microcomputer, precisely controlled positioning of the receiving block is carried out whereby the filter slit 74 and the diffraction grating 79 are thus used to control the frequencies of the IR signals directed by the diffraction grating 79 to a focusing mirror 80. In particular, each position to which grating 79 is rotated corresponds to the sensing of a particular IR frequency. As the block is rotated back and forth between two positions, the frequencies B and A, for example, powers corresponding to the two positions of the block are sensed by an IR detector means 81.

In order to calibrate the spectrometer 11, calibration source 82 provides a reference black body, the temperature of which may be set to a series of known temperatures. The corresponding IR emissions radiated from the source 82 are directed to the focusing mirror 80 by the planar mirror 78 which has been rotated into the appropriate position. From focusing mirror 80, the emission from source 82 is reflected to detector 81 to thus facilitate calibration.

The foregoing wobble mode of operation (and associated IR detector) is a very important aspect of the present invention. As discussed above, this is a technique of IR spectrometry where the detector operation is switched between two wavelengths rather than using a single wavelength, and sensing in turn two different temperatures. In conventional IR technology, a "bow tie" element of a known temperature is swung across the incoming light beam such that first the sensor sees the light beam, then it sees the bow tie, in a repetitive sequence. This sequential operation generates an AC signal which is the only kind the IR sensor preferably used in the present invention can support on a continuing basis. That is, the sensor 81 preferably used in the present invention is AC responsive as opposed to DC responsive inasmuch as AC response detectors are more sensitive where a typical AC responsive detector may be of the lithium tantalate pyroelectric type.

The wobble technique of the present invention changes the wavelength between selected values in much this same sequential manner. In this way an AC signal is generated by the difference between the sensor response at the two wavelengths, without the need for a bow tie system although, as discussed hereinafter, a bow tie system may be employed if so desired in many of the applications of this invention.

With the wobble technique, the transition between wavelengths is very fast, and the dwell time at each of the wavelengths is sufficient that the response window is essentially a square wave. Moreover, the wavelength selector system may be provided with a large number of programmable steps where grating 79 may be stopped to provide input of a related wavelength. Two step numbers can then be selected, and the system will wobble between the related wavelengths to provide an AC signal.

Another and preferred embodiment of an IR spectrometer is illustrated in Figure 5A includes a lens tube 90, a slit 92, a focus mirror 94, a grating 96, a detector 98, a stepping motor 100, a motor back shaft 102, a thermal shield 104, an enclosure side 106, a flange 108, and an O-ring groove 110.

Light enters the lens tube 90 from an IR window (See 71 of Fig. 32, for example) in the skin of the aircraft. This light has, in most installation systems, been reflected from just off of the center line of the aircraft, such that it is coming from a great distance directly ahead of the aircraft. A nearly 45 degree gold coated, heated mirror reflects this light from ahead of the aircraft into the window, if the window is located approximately perpendicular to the line of motion of the aircraft. See Fig. 32 for a rotatable, tiltable mirror which may be used for this purpose. Alternatively, the mirror may be stationary and positioned at approximately 45^*, as stated above.

The lens tube 90 is sealed against the IR permeable window with a flexible tube. The lens tube contains an IR lens which focuses the light into the slit 92. The light reflects from the focus mirror 94 onto the grating 96 which is tipped at such an angle that only the desired wavelength is reflected onto two detectors 98. Other wavelengths are reflected at other angles from the grating such that they do not strike the active portion of the detectors.

Dual detecting techniques are known from U.S. Patent No. 4,266,130, incorporated herein by reference, and are used in the present invention with many logical advantages. Both detectors 98 accept light from the same horizontal cone direction where the first detector is typically directed straight forward along the aircraft line of flight and the second detector typically directed slightly (2 to 4 degrees) upward with respect to the other detector. This provides a capability to measure the look distance when the lapse rate is known, and vice versa, as discussed below. it also permits a slight up-look when in the take off mode. It also provides back up detection capability in case one should fail.

The two detectors are preferably built into the same package for accurate placement, common temperature, and common filter. The filter preferably serves as the front cover of the detector package, and the inside of the package is blackened to minimize the reflections. The detector temperature can be measured with an attached temperature sensor.

The grating is preferably directly mounted to the front shaft of the stepping motor 100. The stepping motor is thus able to establish the angle at which the grating is tipped, and thus the wavelength of IR which enters the detector. Because of the direct connection of the grating to the shaft, this removes all play from the positioning system that establishes the wavelength.

The motor back shaft 102 mounts a fiducial alignment plate and a mechanical damper, as discussed below. The fiducial alignment plate makes it possible for the grating to be automatically and accurately aligned to give the desired wavelengths to the detectors. The damper reduces the ringing of the grating/shaft system of the motor. This makes the change in wavelength most nearly a square-wave function. Since the detectors can respond to an alternating light signal, they report the

T differences between the signals at the two IR frequencies chosen in a given time period by the grating positions, as discussed above. The details of an illustrative damper body are shown in Figures 5B and

5C where the damper is mounted on back shaft 102 via tubular shaft 118 and includes a plurality of enclosed sections 120. Each section may be loaded with lead shot, brass clips or any other suitable damping medium.

The quantity of the lead shot in each section may be empirically determined so that the lead shot strikes the front wall of each section just as the stopping unit starts to spring back. There is an extended range of times over which the shot reaches the forward wall.

This damps the first return bounce over a wide range of frequencies, stopping most of the ringing of the system. Moreover, due to the number of sections, many walls are presented for contact thereby substantially enhancing the damping action. Further, the lead shot provides a large mass striking the forward walls. This mechanical type of damper is more desirable than the more conventional magnetic types, for the magnets radiate a sufficiently strong magnetic field among the aircraft instruments to cause concern about interference.

The fiducial plate is shown in Fig. 5D and includes a hole 124 for mounting on back shaft 102 and a slot 126. The leading edge 128 of the slot is sensed by a light source optical sensor system (not shown) mounted such that the plate operates between the light source and optical sensor. To generate the predetermined grating positions, the stepping motor rotates a known number of fixed size steps from the edge 128 of the slot. In the starting process, the motor returns to the edge 128 of the slot and again counts off the steps of its rotation to the correct number. For most continuing operations it keeps accurate track of its count.

The thermal shields 104, and one on the other side of the motor mount, help minimize the scattering of heat energy (IR) from the motor and the motor control drive circuits that are mounted on the enclosure side at 112. Heat from these sources warms the entire enclosure, but does not get beamed or reflected into the optics in such a way that it can interfere with the signals from the light outside. The heating of the enclosure may be corrected out by functions of the enclosure temperature, motor temperature, and detector mount temperature.

Local spots of heating, or cooling on the surfaces exposed to the detector or other parts of the optical system can lead to stray signals that undesirably affect the measurements. The entire enclosure, however, can operate at any sequence of temperatures from -55 to 70 degrees C. The thermal impedance of the container and optics exposed parts is kept low to minimize the local differences. The stepping motor may be a major source of heat, and its temperature m_.ay be separately measured and used in the thermal corrections.

In certain applications, the sensor is affected by the rate at which its temperature changes. It can operate accurately from -55 to +70 degrees C with no problems, but it cannot change temperature faster than about one degree per minute. Because the detector is deep inside its enclosure 114 (see Fig. 5E) , this is not a problem when the box is in a heated part of the aircraft. When the box is in a wheel-well, however, the take off and landing can expose the box to large, fast changes in temperature. Because these temperature changes occur at the most critical times in the operation of the system, electrical heaters and coolers may be employed although a highly insulating jacket on the enclosure may also be sufficient to cut the internal rate of temperature change to an acceptable value.

The enclosure 114 (Fig. 5E) including optics compartment 116 (Fig. 5A) and electronics compartment 117 provides a vacuum tight seal such that moisture cannot enter and condense on the optics or electronics. The enclosure or box also provides an accurate rim that can be used for physically mounting the enclosure to the aircraft structure while aiming its lens at the window in the aircraft skin. A mirror (as discussed above) outside the IR window directs light from ahead of the aircraft through the window into the optical bench in the box. The mirror may be heated by internal heaters.

A single multipin connector 119 joins the box with all of the required aircraft connections which may include the aircraft compass. Alternatively, a compass may be provided in the box. this will permit the system to be aware of changing inputs accompanied by changes in direction.

The box 114 preferably contains all of the system except the window and mirror system, and the cockpit warning devices. These are either separate or part of a cockpit aural warning system, and warning lights and reboot/test switch.

Power is provided by the aircraft power system, and conditioned by DC/DC converters and filters inside the box. Input from the aircraft instruments comes either directly from the instruments to dedicated receiver circuits, or by way of the aircraft computer (such as ARINC-429) , or a combination of both, as required.

The electronic compartment of the box may have a sealed auxiliary opening by which EPROM's containing software can be exchanged. This makes possible the calibration of the unit with a "standard" EPROM, and then the insertion of a specific EPROM via a removable plate 121 that contains the calibration constants after they have been established. The calibration EPROM may be electrically reprogrammed from outside the box.

The flange 108, as discussed above is a means for connecting the optical enclosure 116 with the electronics enclosure 117. The o-ring groove 110 provides a space for the o-ring that seals the two enclosures together. This permits the atmosphere inside to be established during the production calibration period and remain essentially unchanged during the use of the instrument. There are also o- ring seals around the lens, at the mount of the lens tube, at the top of the electronics-side cover, and about the operation constant memory element replacement cover in the electronics enclosure. The electrical feed through is also a sealing type.

The operating pressure is not as critical as is the dew point of the gas about the optical elements. A layer of condensed (or frozen) water on the surface of any one of the optical elements could seriously reduce the sensitivity of the device. The normal operating temperature of the enclosure could soon warm the elements to above most dew points, but it is important that the system operate correctly for the first take off after any cold night on the ground. Accordingly, a moisture sorption system may also be provided inside, in addition to sealing all the joints. An O-ring slot in the electronics mounting plate 136 of the electronics-side permits this cover to be removed without removing any of the electronics circuit boards or electrical connections. This is advantageous in trouble shooting the instrument, for it permits the operation of the instrument without this cover.

Several optical filter systems may be included in the system. Thus, a transmission filter (not shown) may be provided at the entrance to lens tube 90. The grating 96 provides the greatest filter action. To further eliminate the problems from looking directly into the sun, a focal plane isolation filter (not shown) may be employed which absorbs light from parallel rays striking the lens. The cap on the center of the lens, as is conventional, need not be employed.

Multi-pin electrical feedthrough 119 serves all the inputs and outputs. The 28 volt power from the aircraft may be used directly, and to convert to the other various DC voltage levels used inside. The converters and filters are all inside compartment 117.

As nearly as possible, the circuits are the same for all types of aircraft. The circuit boards of Fig. 5E are in part mechanically tied together by the multi- pin connectors which pass the electrical connections from board to board. Moreover, many of the circuits can be combined into several solid state devices. This permits the size and weight of the electronics part of the package to be very significantly reduced.

The EPROM (not shown) which holds the constants for the operation of each specific device may be plugged-in such that only after calibration can it be replaced with one which holds the correct values. Thus, a small sealable door or plots 121 may be provided over the EPROM to permit its replacement. Alternatively the EPROM may be reprogrammed in place thereby eliminating the need for the door.

The circuitry and software utilized in compartment 117 will be described hereinafter. The lowest layer between the electronics and the optics, is interface plate 135, it being a shield that prevents local warmth from the electronics from giving a non-correctable input to the sensor. Although the arrangement of Fig. 5E is preferred, it is also possible to provide the optical components in one enclosure and the electrical components in a second enclosure.

There are many different wavelength functions that can be applied with the wobble system. These have been given mode numbers and letters in the following explanations although there is nothing fundamental about these designations. Moreover, the following modes are intended to be illustrative of other modes that can be employed.

Mode 2 uses two wavelengths that have relatively short look distances in the atmosphere. The first of these, Wl, will typically see out of the AWAS lens system only a few meters. It thus relates to the temperature of the air just outside the aircraft. The second wavelength, W2, sees only a few hundred meters, if that far. Thus, these two temperatures can usually be considered to be the same. Even though the temperatures seen by these two wavelengths are usually identical, the detector sensitivity or gain with respect to them is significantly different, and thus there will be a difference in signal level between the two. This will provide an output, as described in detail hereinafter that represents the difference in light power from each of the two wavelengths times the sensitivities at each of them. Eq. 1 shows this. -. Signal = kl*Tl - k2*T2 (1)

Here kl is the sensitivity of the detector to the power of the light of Wl and k2 the sensitivity for W2, Tl is the air temperature in the region that Wl can see and T2 is the air temperature in the region that W2 can see.

From each wavelength, the immediate signal part is a signal that represents the difference in temperature between the detector and the total light input of the selected wavelength times the sensitivity for that wavelength. The total light input is a function of the temperature of the atmosphere at various distances. If all of the atmosphere that Wl can see is at temperature Tl, that part of the reading is K1*T1. When the wavelength changes to the second wavelength, W2, a new signal value will be established. This will be based upon the temperature in the region in which this wavelength can be sensed.

If the temperature in this local region is all the same, Tl will equal T2. Then Eq. 2 pertains.

Signal = Tl(kl - k2) (2)

Because the k values are constants, the signal is a linear function of this local temperature. Because the distance that the two wavelengths can see is very short, and very similar, the temperatures have little chance to be different. Thus mode 2 is quite accurate as an indicator of the local temperature.

Mode 3 is similar to mode 2, except that the wavelengths are chosen with long look distances. Typically the wavelengths are quite close together, but with significantly different sensitivities or gain values. This mode makes it possible to see temperature changes occurring at considerable distances.

Mode 9 combines a short look distance wavelength and a long look distance wavelength. This mode has been typically used in the system of the above- mentioned co-pending application to provide the delta T measurement used in the determination of the hazard index F.

Mode 12 is a mode 9 type of operation used for CAT detection using water based wavelengths between 17 and 21 microns.

In mode 4, a number of wavelengths are stepwise scanned whereby a spectrum scan and calibration capability is provided.

As discussed above, a conventional bow tie modulator may be used in place of the wobble system of the present invention in many of the applications of the present inventions, even if it is less efficient than the wobble system where the bow tie system may be used by itself or in combination with a wobble system.

An infra red (IR) detector of the type preferably employed in the present invention is able to provide only a stable alternating signal. The alternating, or at least periodically changing aspect of its signal is caused by a change in light magnitude reaching it in periodic sequence. The wobble system periodically moves the frequency controlling grating to shine first one wavelength (WL) of IR light, then another on the sensor. By wobbling back and forth from one WL to the other, a periodically changing light input is established. This results in a periodically changing output which meets the operational need of the sensor system.

The more conventional response to the needs of the sensor system is the bow tie system that is most often rotated outside the optical system, in front of its entry. As a black arm of the tie swings in front of the optics, only the tie is seen by the optics, and thus by the sensor. When the tie rotates out of the way, the system sees its intended target. A moment later, the other arm of the bow tie blocks the optical entry. The result is an alternation between the bow tie and the intended target. This provides, the alternating aspect for the signal as required by the sensor.

The light power sensed by the detector is first that of the bow tie, as shown in Eq. 2A,

PWRbt = KITbt (2A) where Kl is the system sensitivity at the WL(s) involved, and Tbt is the temperature of the bow tie. When the bow tie arms are out of the way, the light power sensed is shown by Eq. (2B) ,

PWRobj = K1T1 (2B) where Tl is the effective temperature of the event or object seen, as measured at the WL(s) involved. The magnitude of the alternating signal is then the difference between these, as shown in Eq. 2C.

PWR1 = Kl(Tl-Tbt) (2C)

Because the temperature of the rotating bow tie is difficult to measure to the accuracy needed for the Tl measurement, Tbt must be removed from the equation by subtraction.

To do this, a second wavelength of light is used. This gives Eq. 2D

PWR2 = K2 (T2 - Tbt) (2D)

Equations 2C and 2D can be combined to remove Tbt, as shown in Eq. 2E.

PWR1/K1 = PWR2/K2 = Tl - T2 (2E) Thus, after two sets of measurements, with the assumption that Tbt remained constant between them, a function is obtained which is closest to that obtained with the first set of grating motions in a wobble unit. See Eq. 2F and Eq. 1.

PWRlw = K1T1 - K2T2 (2F)

In both cases, the K's must be known from calibration before the delta T values can be known. They must be applied in different ways, however, as shown by Eq. 2E and Eq. 2F.

The fact that twice the measurement operations are needed when using a bow tie is an important statement in favor of the wobble. It is true that the conventional system spends half of its time measuring Tbt which then must be removed from the data.

With respect to the foregoing wobble mode, the following describes certain features which are considered to be novel although, as discussed above, there are other features. Thus, the use of two wavelengths to eliminate the use of a chopper for accurate IR measurements is an important feature. Moreover, the use of two wavelengths that are quite close together to determine a single location temperature is novel. This takes the place of the use of a single wavelength, and a chopper whose blade temperature is known. Determining the blade temperature can present problems.

The use of two selected widely different wavelengths provides a delta temperature measurement between near and far locations, without actually establishing either the far or the near temperature, per se. This takes the place of two temperature measurements and a subtraction.

Use of a moving grating to select wavelengths is known but the use of a grating to take the place of the chopper is very unique. It has many advantages, one of which is the use of the detector all of the time. In the chopper system, the detector is looking at the chopper half of the time so that the system will work.

The dual wavelength measurements are a function of the detector temperature and calibration of the device at all possible values of detector temperature resolves this dependence.

Having now described various techniques for measuring the delta T's, an illustrative embodiment of an overall system will be now described where this system is directed not only to the detection of LLWS by a plane landing or taking off but also to the detection of (a) LLWS by a stationary plane on the runway preparatory to take off or (b) clear air turbulence (CAT) at cruising altitudes.

With reference to the overall system diagram of Figure 6, it can be seen that the system 1 is comprised essentially of two major components, i.e. , the electronics unit 117 and an optics unit 116, as described above with respect to Fig. 5E. Unit 117 has an aircraft instrumentation interface at 12, whereby any combination of instrument signals can be received either directly from the aircraft instruments, themselves, or from the aircraft computer or from combinations thereof. Such aircraft interface signals can provide data to the electronics unit 117 from for example, the aircraft's radio altimeter, pressure altimeter, air speed indicator, pitch indicator, air temperature gauge (outside air temperature) and heading indicator. The aircraft instrumentation signals received at 12 are delivered to a multiplexer 13 via a signal processor SP. Also fed to the multiplexer 13, are signals from sensors associated with the operational components of the IR spectrometer 116 and the signals relating to temperatures sensed by the optics of the spectrometer. Based upon instructions from the central processing unit (CPU) , the multiplexer 13 directs selected signals to the CPU, and in the case of an analog multiplexer, for time-share processing of analog signals from aircraft instrumentation, optics and the like, the selected signals are passed through an analog/digital converter, as shown.

The CPU provides master control, computation and sequencing. That is, it not only controls the flow of data to it from multiplexer 13, but performs the computations necessary to trigger issuance of audio and/or optical warnings to the pilot from the warning system, and at the same time controls the programming of a stepper motor computer that controls the frequency selection processes, described above, by issuing drive control signals for adjusting the diffraction grating via a grating drive 14 and receives a feedback signal from the grating drive for monitoring the functions performed at any given time.

The grating drive can comprise a stepper motor and a stepper motor driver in the form of a high current amplifier that is matched to the stepper motor characteristics. The stepper motor driver is controlled by the motor control computer to provide the repetitive sequences required for each function that the optics of the IR spectrometer 116 must perform.

Fig. 7 is a flow diagram depicting the flow format of the general system. When the power is turned on, usually before starting of the aircraft engines, the computer is booted and a series of tests applied to be certain that the computer and all of the testable aspects of the hardware are operating correctly. If the tests are failed, for example, three times, warning lights remain on to indicate to the pilot that the system is not available to help make weather related decisions.

During normal operation, the CPU will proceed on to the main loop. However, when the aircraft connection cable (12, Fig. 6) is disconnected, a configuration switch 15 automatically places the system in a diagnostic mode. In the diagnostic mode, switch 15 is open and allows other configuration switches to be attached and used for testing and calibration of the system. The details of such calibrations and tests, forming no part of this invention, are merely generally reflected on the diagram of Fig. 7,

During normal operational circumstances, as indicated, the system will directly pass from the starting loop into the main loop of the system, at which point the CPU must establish its situation since, otherwise, it would be unable to distinguish a re-start following a momentary power failure from initial s_.tart- up. Thus, the CPU examines the air speed and altitude signals from the aircraft instrumentation in order to determine whether the aircraft is on the ground, taking off, landing, between taking off and reaching 15,000 feet, between 15,000 feet and landing, or above 15,000 feet. As shown in greater detail in Fig. 8, the system first checks to determine whether the air speed is greater than 60 knots per hour, and if it is, the altitude is checked. If the altitude is greater than 15,000 feet, the unit transfers from a mode used for determining the existence of low level wind shear (LLWS) to one which checks for clear air turbulence (CAT) . If the pressure altimeter indicates that the altitude is less than 15,000 feet, a measurement is taken from the radio altimeter. An indication that the altitude is less than 2,500 feet is indication that the aircraft is either taking off or landing, while if the altitude is between 2,500 and 15,000 feet, a region where no action is required in most circumstances, the control will wait for a period of time, such as a second, and then repeat the test.

Now, when the altitude (RALT) from the radio altimeter indicates an altitude of less than 2,500 feet, the system proceeds to establish whether the aircraft is taking off or landing. This is achieved from the derivative of the air speed indication in the step designated ACCEL. If an acceleration over 2.5 knots per second per second is determined to exist, the computer enters the TAKE OFF mode, while reading less than that results in the computer entering the LANDING mode.

In the landing mode, the system re-checks the altitude to determine if it is greater than 2,300 feet, and if it is, it then performs self-tests and usual landing preliminaries before entering the LLWS mode since adequate time exists for that purpose. On the other hand, if the computer finds tnat it has "awakened" at too low an altitude during a landing or anytime that it starts or re-starts in a take-off, it immediately enters the LLWS modes since no time exists to test anything. From the above steps, it can be seen that no matter when the computer finds itself activated or re-activated, it finds its way to the correct mode of operation.

While the system is operating in the LLWS mode, it checks to determine whether it should still be operating. This is done by checking the air speed, and once the air speed is below 60 knots per hour, its function is completed since the aircraft has either completed its landing or has aborted its take-off. On the other hand, if the air speed is higher than 60 knots per hour, the program checks the RALT to see if the altitude is less than 1,600 feet and continues operation if it is. However, if the altitude is above that at which low level wind shear is a problem, the program goes to the WAIT or LAPSE mode, and checks to see if it is high enough for clear air turbulence (CAT) .

The flow chart of Fig. 9 indicates the steps that may be performed during the LLWS mode. In particular, first, the above mentioned mode 9 wobble may be performed by moving the diffraction grating to alternately and repeatedly sense the two different wavelengths under the direction of the motor control computer regulation of the grating drive. The data from these measurements is combined with the data from the aircraft instrumentation to calculate the wind shear index where this index can be calculated on the basis of the various hazard factor equations described hereinafter.

The hazard index calculation is transmitted for comparison against predetermined threshold values which, as noted in above referenced co-pending application, will be a function of the aircraft's performance capabilities and, for jet aircraft of the type used by scheduled carriers, normally is in the range of 0.12 to 0.15. If this threshold is exceeded, an alarm is sounded for a period of time, for example 1 minute, after which the alarm is held for a period of time, for example 30 seconds. The cycle is then repeated. In this regard, as reflected by the diagram of Fig. 8, even though not reflected in Fig. 9, the LLWS cycle includes re-checking of the air speed As and the altitude RALT from the radio altimeter, and a gain check is performed every 10 minutes to insure that the equipment is operating properly.

When the system switches to the CAT mode, a sequence of steps are executed, as shown in Figure 10, this sequence being similar to that of the LLWS mode except the CAT index calculation is different as described hereinafter.

The overall operation described above with respect to the programs of Figures 7-10 may be illustratively summarized as follows with respect to Fig. 10A, which depicts an illustrative flight profile where the following steps occur at the points indicated in Fig. 10A:

(A) POWER UP

Boot (automatic)

System check

ARINC check

Gain check All tests pass or else fail lamp is illuminated.

(B) WAIT

Until A/S is greater than 60 kts (automatic) or until Mechanic Call is pressed to instigate Runway Scan.

(C) RUNWAY SCAN (pilot initiated)

Upward looking detector scans for shear (D2) . OAT thermal feature engaged. 12 second duration test (Red and Amber lights flash) . If no event is detected, the lights extinguish. If event is detected. Red lamp and speaker so indicate.

(D) TAKEOFF ROLL (automatic)

Initiation when A/S is above 60 kts.

Upward looking detector used.

Initiate predictive and thermal features. Go to QUIET COCKPIT mode when A/S exceeds 90 kts.

(E) TAKEOFF CLIMB (automatic)

Radio altitude greater than 40 feet AGL, reconnect speakers. Continue predictive and thermal features. Use upward looking detector. Radio Altitude is less than 2500 feet AGL.

(F) LAPSE DATA ONLY MODE (no warnings, automatic)

Radio altitude greater than 2500 feet AGL. Data provided to recorder from modes 2,

3, 9, and 12. Both detectors in operation. Gain check performed.

(G) CAT MODE (automatic)

Pressure altitude exceeds 15,000 feet. Data updated every 15 seconds. Mode 12 used with water frequencies. Gain check performed each 10 minutes. Both detectors in operation. (H) LAPSE DATA ONLY MODE (no warnings, automatic) Pressure altitude is less than 14,000 feet, descending. Data provided to recorder from modes 2, 3, 9, and 12. Both detectors in operation. Gain check performed. (I) LANDING (automatic) Radio altitude is less than 2500 feet AGL. Unstable air warning provided by Amber lamp if lapse rate indicates unstable atmosphere. Radio altitude is less than 1500 feet AGL goes to mode 9. Returns to WAIT (B) when A/S becomes less than 60 kts.

With respect to the calculation of the LLWS hazard index factor of Fig. 9, various hazard factor equations which may be used will now be discussed. As described in the above-mentioned co-pending application, the hazard factor may be a function of at least (a) a first term which is a linear function of at least A T and (b) a second term which is a non-linear function of at least AT as exemplified, for example, by Eq. 3.

F = K*^T + (J/A_s)/( Δ T/Tm) ^l (3) where K and J are constants; ΔT is the temperature difference between the aircraft ambient temperature, Tm, and a remote temperature, T; and A_s is the air speed.

Other forms of the hazard index equation may also be advantageously employed where, for example,

F = K*d( Δ T)/dt + (J/As)* /^"(ΔT) ' (4)

As can be seen in Eq. (4) neither of the terms contains Δ linearly. That is, both terms are non¬ linear where one is a time derivative and the other a square root. Moreover, another equation which may be advantageous in certain situations is that where the first term is at least a non-linear function of T such as the time derivative term of Eq. 4 while the second term is at least a linear function of ΔT such as J/A_s ( T) where the latter term corresponds to that of Eq. 3 except the square root and Tm are not utilized.

In general, the hazard factor is preferably comprised of two terms both of which are at least functions of _£T where the first and second terms may each be either linear or non-linear functions of T , as discussed above, for example, but where the first and second terms are not linearly combinable.

Also, the determination of T need not necessarily be in terms of a measurement of a near temperature and the separate measurement of a far temperature such that ΔT is determined from the difference of these measurements. Rather, this, temperature difference can be measured directly by the wobble technique discussed above with respect to the IR spectrometers of Figs. 5 and 5A.

An advantageous form of the hazard index equation is exemplified by Eq. (4) where the first term includes the time derivative of T where Δ is typically the far temperature minus the near temperature. The term ΔT is often thought of as a time derivative; however, in Eq. 4 it is a location derivative. A ΔΔT term, which will be employed hereinafter and which corresponds to the d(ΔT)/dt term of Eq. 4, is the time derivative of this location derivative. To reduce the effect of noise, three successive values of delta delta T are averaged to get the delta delta T value used in the first term of Eq. 4, as will be described in detail below with respect to Fig. 12.

Considerations relating to the desirability of a first term including ΔΔT and with respect to the averaging of three (or another suitable number of) successive values in order to obtain ΔΔT are as follows.

As an aircraft gets closer to a weather event such as a microburst, the ΔT signal difference as a function of distance to the thermal transition becomes larger and larger as the detector senses more and more of the new temperature air. Because of a logarithmic distance effect, the closest air has the largest effect on the signal. Thus, the largest change will come just before the aircraft encounters the event. However, when the aircraft is in the event, the change in signal with travel, or time, will immediately become zero, which corresponds to an unusual derivative function, as illustrated in Fig. 11A. Then, as the aircraft approaches the exit side of the microburst, the signal change goes through a negative version of this same curve.

In Fig. 11B, there is shown the general form of the sensor signal that provides the derivative shown in Fig. 11A. The absolute values of Fig. 11B can be quite small. It is the changes in these values with aircraft position, or time, that are of concern, thus the use of the derivative in the first term of Eq. 4.

The magnitude of the derivative is a function of:

1. The size of temperature change that is being approached.

2. How fast the aircraft is approaching it.

3. How close the aircraft is to the change. Calculations using the airspeed indication in the second term are used to remove the approach rate variable. The atmosphere is a thermally noisy medium. Thus, a single negathve derivative spike has little meaning. Thus, this signal should decrease over an extended number of seconds before a warning is given that a weather phenomenon of concern has been detected. This means that the derivative function should go negative and remain there, except for noise, for those seconds. Accordingly, three successive values of ΔΔ are averaged in order to obtain the Ac\ for the first term of Eq. 4.

As will be described in more detail hereinafter with respect to Fig. 12, delta T in the second term is preferably not taken directly from the delta T data, but is re-established over a period of 20 sequential measurements from delta delta T values where the 20 (arbitrary number) measurements are summed on a first in, first out (FIFO) basis to obtain the delta T-.used in the second term. This makes that value very stable relative to the individual measurement by measurement value. In addition, the delta delta T values are preferably clipped by setting limits to the values in order to avoid violent noise. Thus, the delta T in the second term is established by summing the sequential 20 (arbitrary number) of delta delta T values. This is a unique way to obtain a zero independent delta T value with considerable smoothing.

Empirically, it has been determined that a particular useful form of Eq. 4 is as follows:

F=K*d(delta T)/dt (+/-) (J/As)* J |<d(delta T)dt|' (5)

Note the square root term does not include the Tm divisor. Moreover, the sign of the sum is not included under the square root. It is used in front of the entire term. This permits the delta T sign to influence the result, rather than create mathematical problems in the calculation. Thus when the sequential summation that recreates delta T becomes negative, the second term is negative.

Referring to Fig. 12, the wobble technique for deriving Δ ^τ is illustrated in further detail. The output from detector 81 of Fig. 5 or 98 of Fig. 5A is applied to a wobble filter 200, the frequency of which corresponds to the rate that grating 79 of Fig. 5 or grating 96 of Fig. 5A alternates between the first and second wavelength measurement positions. Typically, these gratings alternate between these two positions at a rate of 3 Hz, for example, and thus the frequency of wobble filter 200, in this instance, would be 3 Hz. The 3 Hz output signal from the filter is applied to a pair of rectifiers 202 and 204 where rectifier 202 rectifies the positive going portions of the output signal from the filter and rectifier 204 rectifies the negative going portions. The rectified positive -.going portions are applied to an integrator 206 while the negative going portions are applied to an integrator 208. Typically the integrators 206 and 208 are reset every two seconds. Thus, since the output signal from filter 200 varies at a 3 Hz rate, about six positive going portions of this signal are summed in integrator 206 every two seconds while integrator 208 sums a corresponding six negative portions. Accordingly, the • integrators 206 build up the magnitude of the detected signals every two seconds. Moreover, these integrators assist in the removal of noise from the detected signals. The outputs of the integrators are applied to a subtractor which generates a T signal.

This A signal may be directly applied to a computer for further processing. Alternatively, the signal may be applied to the remainder of the circuitry shown in Fig. 12 which performs functions similar to those that would be performed by the computer. Thus, the lapse correction is performed at 212 and is described in further detail at Figs. 26 and 27. In general, the purpose of the lapse correction is to adjust the value of Δ T to compensate for the pitch of the aircraft. That is, the various index formulas of the present invention are typically based on a horizontal orientation of the aircraft and thus, if a ΔT measurement is made at an angle with respect to the horizontal, the measured ^T value should be compensated in accordance with either a default (predetermined) lapse rate or a measured lapse rate, as will be described below.

The lapse rate corrected value is indicated as AT ' in Fig. 12. However, it is to be understood that ΔT, as used in the various equations described above is not primed. However, in fact, the Δ^τ values used in these equations are preferably lapse rate corrected values.

In order to calculate the LLWS hazard factor of Equation (5), the lapse corrected value of Δ^τ is first applied to circuitry for obtaining the delta delta T value to be used in the first term of the equation. In particular, this circuitry includes a two stage FIFO shift register 214, the outputs of the two stages of which are applied to a subtractor 216 whereby the subtractor obtains the difference between successive values of with respect to time - that is, delta delta T. Each of the delta delta T values are applied to a clipping circuit 218, the purpose of which is described in more detail below. The clipped delta delta T values are applied to a three stage shift register 220, the outputs of the three stages being applied to a summing circuit 222 and then to a divider 224, which divides the sum of the three stages by three whereby an average is taken of three successive values of delta delta T to thus obtain the delta delta T value used in the first term of Equations (4) and (5) , as discussed above. In particular, this average value of delta delta T is inverted by invertor 226 and applied to LLWS index calculation circuit 228 where the hazard index Equation (5) is solved by employing an appropriate value where A_s is obtained from instrumentation conventionally employed with the aircraft and where J and K are empirically determined constants available to calculation circuit 228.

The purpose of clipping circuit 218 is to clip the delta delta T outputs from subtractor 216 so that no one particular pulse, which is greatly affected by noise, will distort the average value of delta delta T as calculated by elements 220, 222, and 224.

In order to obtain the ^τ value which is used in the second term of the hazard index Equation (5) , the delta delta T output of subtractor 216 is applied to a clipping circuit 230, the purpose of this clipping circuit being the same as that of clipping circuit 218. The output of clipping circuit 230 is applied -.to a twenty stage FIFO shift register 232. The outputs of the twenty stages are applied to a summing circuit 234 to obtain the requisite Δ ^τ value. Note that by summing the delta delta T value as applied to shift register 232 the derivative of the A T signal is effectively integrated to thus re-establish the ^τ signal. Moreover, this is done in such a manner as to substantially reduce the effect of noise on the £ T signal used in the second term of the equation.

Assuming hazard index Equation (5) is being employed, the absolute value of «ΔT is obtained by absolute value circuit 236 and the sign of the sum is obtained by sign detector 238. These are applied to LLWS index calculation circuitry 228 whereby the index calculation is made employing equation (5) .

The index calculation is made every two seconds - that is, each time the integrators 206 and 208 are reset and the shift registers 220 and 232 are shifted. As discussed in the above mentioned co-pending application, whenever the calculated index exceeds a predetermined threshold such as 0.15, a warning is provided.

As will be discussed hereinafter, other temperature related phenomena including weather related events may be detected utilizing the lapse rate corrected value of ΔT' and accordingly, this value is also applied to clear air turbulence (CAT) detection circuitry 242 and SAWS (stationary air warning system) II index calculation circuitry 244. Moreover, SAWS I index calculations, which utilize the delta delta T signal, may also be obtained from circuitry 246.

Note the foregoing hazard index equations can be used both in IR-predictive systems (where A ^τ is obtained from IR measurements) and thermal-reactive systems (where ΔT is obtained from successive outside air temperature (OAT) measurements made by temperature measurement means conventionally mouniced on-, the aircraft) . Moreover, a combination of IR and thermal, and a combination of either and both of these with the inertial system may be employed. Thus far, the combination of predictive IR and OAT thermal/reactive has proved most effective. The thermal reactive system senses the microburst at considerable distance if the aircraft altitude is sufficiently low, but may not see it until contact is made at higher altitudes where there may be no outflow. The thermal reactive system serves as a backup for the IR system if the IR window gets dirty, or an IR component fails. It provides prediction for most encounters, and is better than just an inertial system in most cases.

The thermal reactive system using the OAT temperature measurements from the aircraft's own outside air temperature instrument and using the algorithm of Eq. 5 will provide alarm between zero and about one minute before contact - if the IR predictive system has not already sounded a warning. The thermal prediction time is a function of where on the structure of the microburst the contact is made. If it is on the outflow, the warning is quite early. If the contact is on the stem of the downdraft, and there is almost no outflow there, the thermal system gives little if any thermal predictive warning and thus is almost completely reactive.

The IR system provides prediction in accordance with how far the IR can see through the weather. In an over 6 inch per hour downburst, a 36 second prediction time has been observed. Other trips through very heavy rain that contained no microbursts did not give nuisance warnings. Most of the IR predictive warnings were between 30 and 50 seconds before a 150-200 knot aircraft struck the wind determined events.

In most cases the three systems (IR predictive, thermal reactive, and wind-inertial reactive) provided a sequence of warnings on a recorder that is, the IR predictive at about one minute, followed by the thermal at 15 to 30 seconds, followed by the wind (inertial) system when the aircraft was being pushed down, or encountering a concerning loss of head wind at time zero.

In order to further illustrate the use of the FIFO 220 of Fig. 15, certain problems addressed by the LLWS detection system of the present invention will be discussed where the FIFO technique is implemented using the OAT signals of a thermal-reactive (TR) system where, as stated above, the OAT signals are not obtained by an IR spectrometer but from conventional temperature measuring devices conventionally used to measure the temperature of the air immediately outside an aircraft and where ^ is obtained as the difference between successive OAT measurements where the difference is calculated every two seconds for example.

The following situations will be considered: a. Entering a cold front b. Entering a warm front c. Passing through a young microburst (low level wind shear) d. Passing through a mature microburst e. Passing through unstable air

Figs. 13 (A)-(D) show the situation for a. and b. with respect to the OAT, the delta delta T and the computed delta T. The first 46 seconds (assuming the time for a sample to pass through the FIFO is 46 seconds) of the operation corresponding to the 20 unit period of the FIFO (after which the first delta delta measurement is discarded) are accurate, but there is a possible problem due to the increased or decreased sensitivity forced on the system by the accurate picture that it produces. It is only after the 46 second period of the FIFO has expired, and the delta delta T values of concern start to be replaced with normal, near zero values. During those 46 seconds, even a noise pulse might trigger a nuisance alarm.

Conversely, if a warm front is entered, or a cold parcel of air is exited, the reverse situation occurs. The 46 second time period holds the sensitivity at a below normal level so that a microburst or the like might not be detected.

Figs. 13(E)-(G) show a young (or forming) microburst, and the response generated for delta T and delta delta T. The delta T is without bias, and is a good copy of the shape of the actual temperature profile. However, 46 seconds after the start of the temperature fall, there is a delta T reverse picture that complicates the picture. This reverse function may decrease the sensitivity of the device for a period of time that is as long as it spent seeing the actual microburst.

With respect to a more complex mature microburst, Fig. 14 shows a case taken from a July 11, 1988 Denver event. Here the event was so wide that the reverse function started to occur before the peak of the event was reached. The subtraction almost caused the thermal-reactive system to miss the main part of the microburst. Fortunately the system fired on the outside ring of the system in this case. However, if the subtraction had been a small percentage greater, there may have been no earlier warning. Thus, this subtraction should be avoided in most situations.

From the foregoing, it can be seen two unique situations have been identified. These cases are the one for a front where the FIFO has too long a memory, and the one for an event where the FIFO has too short a memory.

When fronts are encountered, FIFO tends to be too long, giving an undesirable sensitivity after the event could be identified, as discussed with respect to Fig. 13. As soon as the non-dangerous event can be recognized, there is no longer the need for the accurate delta T value. The delta T will remain accurate for 46 seconds when coming from calm weather. When the event is shown to be a front, the FIFO need not keep giving high accuracy for the remainder of the 46 seconds. It is better that it be dumped, and start over.

To make the decision that the event is a front, and not a microburst, the delta delta T function can be monitored after a group of detected significant values for delta T. If this averages within Z of zero for X seconds where Z and X can be empirically determined as will be further explained with respect to Fig. 15, the front has been crossed. It is then safe to fill the FIFO with zeros, and continue normal operation. This flushing clears out the delta T that is remembered in the FIFO, and returns the delta T to zero at this new temperature.

When microbursts (low level wind shear) are encountered, FIFO tend to be too short, dumping part of its memory to give incorrect delta T values as the computer tries to continue the search for dangerous events, as discussed with respect to Figs. 13 and 14. This is especially true when a series of events are encountered, and almost warning size events are hit before ones of serious proportion. Here, the system should be kept accurate and alert in order to provide safety. Here a decision must be made that this is a string of events or near events, or otherwise very rough weather. It must opt to increase the FIFO to be accurate for longer than the 46 second time constant of the FIFO circuit.

To make this decision, the computer can again monitor the delta delta T function. If this moves outside the noise band within X seconds after the detection of significant delta T activity, the values of delta T need to stay as accurate as possible. To accomplish this increvse in lifetime, the FIFO value that was first in is added to the incoming delta delta T before introducing it to the FIFO. This feedback system keeps the FIFO from loosing its reference to zero. After the delta delta T values return to noise level, the FIFO needs to be flushed, for it is getting very full. Any FIFO sum that is between plus and minus Q will offer a good time to dump where Q can also be empirically determined.

In this way, the life of the FIFO can be lengthened to eliminate the errors that would normally occur at the time that the FIFO exits a significant event, or combination of them.

The following is needed to implement the above features:

(a) Detection of an event at above noise level.

(b) Knowledge of noise level for delta delta T.

(c) Monitor of delta delta T for relation with noise for timed periods after a.

(d) Switch on results of c.

(e) FIFO switched drain to zero.

(f) FIFO switched feedback.

Figure 15 shows the functional diagram for this system that prevents the above noise level events from detracting from the FIFO function of providing a stable delta T value.

Figure 15 diagrammatically depicts a system for increasing or decreasing the life of the retained temperature values. In block 31 the outside air temperature is obtained every two seconds, for example and utilized to calculate the delta delta T value in block 32 where this value is the average of three (for example) successive values of delta delta T and is utilized in the first term of the above discussed hazard index equations, as discussed above with respect to elements 220, 222, and 224 of Fig. 12. This value may be applied through block 33 clipped if necessary to a first-in-first-out (FIFO) memory 34. Alternatively, delta delta T values may be directed applied to memory

34, as in Fig. 12, where FIFO 34 corresponds to FIFO

232. Each A A is retained for a set period of time, such as the noted 46-second interval as it is stepped through the FIFO. At 35 the sum of the twenty

(for example) current delta delta T values stored in the FIFO memory 34 is taken every two seconds to produce a bias-free delta T value that is utilized in the second term of the warning index algorithm at block

36 in accordance with the previously described techniques. At block 37, the hazard index value F that has been generated by the algorithm calculation is compared with an empirically determined value A which is less than the value required to trigger an alarm

(0.15, for example) , but large enough that it will cause future significant errors in the value of F when the FIFO memory unloads the related delta delta T values. If the value of F exceeds this value A, a process indicated at 38 is started for avoiding an undesired increase or decrease in system sensitivity.

At block 39, a maturity delay M occurs permitting the time required for a typical front to pass before starting an evaluation of delta delta T via the block

40 switch. Thereafter, block 41 tests for X seconds to see if any of the values of delta delta T are outside of the noise level Z established by block 42. If the delta delta T signals do not exceed the noise level Z for this defined period of time, it is reasonably safe to assume that a front has been passed and that the aircraft is not in the middle of a microburst as discussed above with respect to Fig. 13. As a result, block 43 is triggered which sends a zero set signal to the FIFO memory replacing all contents with zero, and at the same time it resets the timers of blocks 39 and

46. Moreover, at block 44, it prevents the change in the hazard index value F created by zeroing the FIFO memory 34 from retriggering block 37 or the warning systems .

If the delta delta signals do not remain within the noise level Z of bloc 41, the event is considered to be a microburst or other unstable weather which warrants extension of the time that the FIFO memory retains the delta delta T signals that caused the original significant F values. To prevent these delta delta T values from escaping from the sum obtained by summing circuit 35, a feedback control switch in block 40 is closed which will result in the exiting values from the FIFO to be reinserted, such that it retains its reference basis for an extended period of time. After delta T falls within plus or minus Q about zero, it can be assumed the aircraft has passed through the event that triggered the system. That is, at block 45 a value of Q is used that represents a value of delta T that is sufficiently close to zero that when delta T is below this value replacing the contents of the FIFO memory with zeros will not significantly change warning detection. This thus prevents the exiting FIFO contents from causing a reflection of the delta T pattern, as discussed with respect to Figs. 13 and 14, for the FIFO is flushed between a single set of data points. At the same time, the timers are reset so that the system is again ready to continue its search. However, should the decision made by block 45 be incorrect, delta T may not come back to zero to dump the FIFO, and delta T could remain high as in the front case. In this situation, block 46 provides a backup function, causing the FIFO memory to be reset to zero after a predetermined time limit.

In summary, it can be seen that the front system shortens the period of time that the F value is offset by the front of Fig. 13, if it is a reasonably steep one. Fig. 16 shows this effect. The initial contact with the front does not differentiate the temperature change cause. Once the most probable cause is established, the related correction is made. This operation may be seen as resetting the zero for the delta T to the temperature at the following side of the front.

The microburst system permits the FIFO to give the correct delta T until the correct delta T value is near zero, or the new maximum time has expired. Then it clears and resets FIFO. This will reduce the possibility for errors occurring during an event. Fig. 17 shows the results of this correction. This operation can be seen as extending the lifetime of the temperature value used for the zero in the delta T, i.e. , the ambient temperature. The delta T is effectively the difference between the 46 second ago ambient temperature and the presently measured temperature.

The effect cf unstable air would in most cases be present with the microburst pattern. Even when not present with a microburst pattern, the delta delta T would be unlikely to remain within the noise level during the test period. As a result, it would involve the FIFO with extended lifetime. This would not create any error for the system, for the FIFO would provide an accurate delta T over the extended lifetime of the FIFO and not slip into a possibly erroneous condition after 46 seconds. The crossing of zero for delta T would in many cases be often enough that no effect would ever be seen.

Should the system get into the front mode due to a quiet period at the test time, the FIFO would get reset, and no error would result. Thus, there is no negative situation resulting from unstable air.

In the above discussed thermal reactive (TR) system, delta Ta can be seen to be a difference at time "a" from ambient temperature, Ta b, as established by OAT history, for use in this system. See Eq. (6) Delta Ta = Ta - Ta b Eq. (6)

By this definition, delta Ta is actually a temperature based on a scale of Centigrade degrees with zero set at the value of the ambient temperature.

As the aircraft crosses fronts, or changes altitude, the Tamb changes. Thus, the system measures Tamb as well as Ta in order to provide the delta Ta measurement. The thermal-reactive system does this by computing delta Ta over a 46 second interval. As discussed above this can sometimes be too short, and sometimes too long. When it is assumed ambient temperature has changed, it is best to forget the old one, and measure a new one immediately. However, when an event is entered, it is best to remember the old ambient temperature until the event is exited, at which time a new meaningful ambient temperature can be established.

In this consideration, the effect of lapse rate (discussed hereinafter) is seen on the T-R system during take off and landing, for here the ambient temperature is changing, and the system sees ambient as the temperature from 46 seconds ago. Meanwhile, the delta delta T is the change per time step of the delta T. If the ambient temperature were constant for a few moments, delta T could be seen as a special temperature measure from ambient, and delta delta T would be a more conventional delta T/delta time, or stepwise time derivative. Thus, it can be seen that the first delta is a difference from ambient, and can be considered a time function only because this difference is measured on a time schedule. Because the aircraft is moving, this can also be seen as difference in temperature with distance.

These temperature functions correspond correctly with the inertial equation discussed in the above- mentioned parent application using horizontal -.wind acceleration and vertical wind velocity to thus relate the thermal reactive system to the inertial one.

The IR predictive system also generates a delta T which represents a similar ambient temperature reference. The difference is that delta T is a measure of the temperature at some distance ahead of the aircraft minus the local temperature, taken without measuring either temperature directly in the case of a wobble measurement.

The FIFO system of Fig. 15, which may also be used with IR predictive systems, generates the sum of the differences between adjacent delta T measurements. This provides the delta T free of any output biases. The delta T that it indicates is the most recent one. With each step, time wise, the FIFO looses a 46 second old delta delta T for each new one entered. If 46 second old delta delta T's are near zero, their loss does not alter the delta T value significantly.

If the FIFO were infinite, the delta T would be the same as the delta T measured. Because the FIFO is finite, it can cause an unwanted function. When an event enters, a problem arises after 46 seconds when the FIFO starts to give up the event data. This is much like the situation discussed above with respect to Figs. 13-14. Thus, the same problem is faced with the IR predictive FIFO's that were encountered with the T-R FIFO and, as stated above, the system of Fig. 15 can also be used with IR predictive systems.

Referring to Figures 18-24, there is illustrated an alternate method of generating and processing the delta T average of Fig. 15. In this alternate method, FIFO input data for demonstration may be clipped to a maximum height of plus and minus 1.0, for example. Thus, the greatest sum is 20.0 built over 40 seconds (2 seconds per value) . Consider a thermal peak that has delta delta T over 1 unit high for 40 seconds of travel. At 185 knots this is 2.4 miles radius, a rather large microburst. Fig. 18 shows the sum over this isolated situation. For the 40 second wide event the solid line shows the sum.

At the end of 40 seconds the FIFO register 34 is filled with + ones's. These values will start to be dumped, one every 2 seconds. If there are only near zeros entering at this time, the curve that has reached 20 will start down with the same slope that it came up..

If the positive side of the event lasts longer than the 40 seconds, there will be a one entering every 2 seconds which will make up for the one that is lost at the other end of the register. Thus, the total remains at 20 as shown by the dashed line. At the end of the event at 60 seconds the register starts to empty. It will take it 40 more seconds to empty completely.

During the 40 seconds required for emptying, the contact with a new positive input will stop the emptying by entering a one while one is being lost at the other end. An increase in total cannot occur until the full 40 seconds of emptying are used. This is shown in Fig. 19. At 80 seconds, the first positive record is gone, and the sum starts to increase. It stops increasing when the second positive input is over at 90 seconds. The sum remains constant until the one's move to the end of the register at 100 seconds, and start to exit. Note that in these figures the top shows the entry values.

Figure 20 shows the addition of two smaller positive inputs if they occur within the same 40 second interval. This keeps the noise on top of a longer event from stopping the summation. However, several small ones can not be distinguished from a longer one.

When both negative and positive entries are employed, there is a more critical situation. If there is a full load of minus 20, and a positive event is encountered, a -1 is lost and a +1 is added at each 2 second interval. This changes the value of the sum by 2 at each 2 second interval. This is twice the slope attained before. See Fig. 21. Thus, the large positive slope of the curve is an indication of a positive input and a negative unloading. If a negative sum is hit that is not ready to unload with a positive input, the sum changes with the unit slope, and thus the system has a long way to go to get to a warning. See Fig. 22.

Because the input to the FIFO is a delta delta T, it is the derivative of the zero based thermal signal. The delta T is not really a derivative, it is more a zero referencing operation. The result is that the input from a microburst reverses sign of input as the center of the symmetrical storm is encountered. Thus, there tends to be a symmetry to the input. A large positive signal is followed by a large negative signal. As seen in Fig. 23, the result when starting with an empty FIFO is a simple peak. Now the FIFO is full. The FIFO starts dumping +l's because their 40 seconds are up. The result is a 40 second period of dumping in which the sum is reversed to become a negative peak. This keeps the next storm from having a chance to register positively relative to zero for about 40 seconds after the end of the previous storm. During the first 20 seconds after the end of the previous storm, the new storm can cause the sum to remain zero. If it continues to enter positive one's for the next 20 seconds, the sum will increase at 2 units per 2 seconds. If the second storm is as long as the first one, negative one's will be entered into the FIFO when it is dumping negative one's from the previous storm. The result is no response at all.

This means that two nearly identical storms adjacent to each other can result in no net signals in the sum for the second one. This is fine if the first one sounds a warning, but is a miss if the second one is the stronger, and the first one did not quite exceed the threshold.

The reverse pattern caused by the dumping can be prevented from confounding the registration of the second storm if the entire FIFO is dumped each time the sum reaches zero. This keeps the pattern in the FIFO from having a negative effect on the sum. The noise cancellation is caused by the ability of the positive inputs to be cancelled by the negative inputs. The pattern inside the FIFO when the sum is zero plays no part in this, but may come back later to be a problem. Thus, dumping at each zero crossing reduces the continuing effect of the stepwise dumping. The stepwise dumping can continue when the FIFO is full and zero has not been crossed. This keeps the system from building to cause a nuisance alert.

If random values of both signs are entering FIFO as noise, the dumping of the 40 second earlier random numbers would seem to cancel some of the entries, and double others. Thus, it adds a different frequency and magnitude of noise to the system. Doing a complete dump at crossing zero should eliminate the new noise caused by this coordinated one cell dump with each input.

Increasing the size of the FIFO to hold more than associated with a anticipated single event could prevent the build-up from being cut short. The zero dump in addition would make this a safer thing to do.

When contacting a front rather than a microburst, it is important to dump the FIFO as soon as the situation is recognized. Otherwise, the system is either held at increased or decreased sensitivity until 80 seconds are over. The function diagram of Fig. 15 describes how to recognize the front and thus, the foregoing modification of the Figure 15 diagram would constitute elements for sensing the output of summer 35 in response to block 41 sensing the presence of a microburst, assuming the output of the summer crosses zero, the entire contents of FIFO memory 34 are then dumped to reset all locations thereof to zero. Thus, the feedback step 33 of Fig. 15 is not employed in this modification of the invention.

If an 80 second wide microburst is encountered, a sum of 20 is built in 40 seconds, and then the sum falls to zero in 20 more seconds. See Fig. 24. If dump to zero occurs at crossing zero, the rate of falling due to the negative half of the derivative is changed to 10 units of sum in the following 20 seconds. However, the aircraft is at the end of the storm, and the system is at a negative value. This would take 40 more seconds to return to zero if it was not interpreted to be a front, and dump to zero after about 20 seconds. Although this is a concern with the zero dump, it is solved by the front solution described above.

From the foregoing, it can be seen that provisions are taken for not only sensing the start of the storm for the aircraft, but also providing a continuous monitor of the continuing danger. Thus, provision is made to follow the danger through the storm. This could keep the approach to a second microburst from being confounded by the response to the first one, etc. If the first one is not quite large enough to trigger the warning, that response must not be of a sort that will prevent the system from giving a suitable warning relative to the second or third down draft encountered.

In addition to the provision made above with respect to Figs. 13-24, other provisions can be made. For example, the IR predictive system should foresee each event. If the thermo-reactive system picks that one up, the predictive one can forget the present one and reset for the next one. Moreover, an ambient zero from before any storm contact can be retained, and this used as a reference for a period of a few minutes:, if the temperature is changing. When the temperature out front is not changing, then the present ambient temperature is acceptable. When the temperature starts to change, freeze a value for the reference temperature.

Furthermore, when a cold storm is viewed from another cold one, the delta T may be too small to provide a warning. This may be accommodated by providing a local temperature measurement to compare with the known ambient, and then a correction of the delta T being based on the ambient.

The OAT temperature compared with the recorded ambient could also provide a better indication of when an aircraft is out of a storm. That is, the use of the OAT system provides a more stable zero for the system.

The CAT index calculation of Figs. 10 and 12 and, in particular, block 242 of Fig. 12, will now be discussed in further detail.

By using certain wavelengths provided by water vapor, an IR system can see ahead by as much as 50 miles. That is, these wavelengths have a look distance of as much as 50 miles. Even at 600 Knots, the water wavelengths provide over 4 minutes to prepare for CAT. In particular, the water wavelengths used in accordance with the present invention are between 17 and 21 microns, specifically at 17.5 and 19.5 microns where the T signal obtained from these wavelengths may be obtained by the wobbling method or the other methods described above.

Moreover, the nature of CAT is such that the interface between different temperature weather streams is not smooth in some cases. Rather, there are regions of one weather system scattered into the other. Striking these small regions with different wind patterns and densities provides the momentary violent changes in air speed and lift relative to aircraft motion. Thus, with respect to the CAT index of the present invention, a search is made for aspects of high noise in the thermal signals that reach the IR from great distances, rather than a search for major transitions. Instead of looking directly at the IR signal, the present invention preferably looks at the standard deviation of the signal from distant water molecules, this technique being described in U.S. Patent No. 4,266,130 to P. Kuhn, this patent being incorporated herein by reference.

This standard deviation signal is then compared with an experimentally established reference. When the reference is exceeded for a sufficient time period, such that a single extreme measurement cannot cause the problem, the warning is provided. Because the warning is typically 4 to 10 minutes before the CAT is encountered, there is time to adequately notify the passengers, check that they are safely strapped into their seats, and clean up the aisles - that is, remove carts from the aisles.

It is possible for clouds to exist in mass at the 15,000 feet and higher altitudes where the CAT mode, of Fig. 8 operates. This means there will occasionally be regions of very high humidity that can have an effect upon the IR signal. If the cloud is about the look distance away from the IR detector, there is no sharp transition, only a growing loss of signal meaning. The problem encountered here is that the water density is sufficient that the water based signal now virtually all comes from the front surface of this nearby cloud. The look distance thus grows shorter and shorter. Because the detector is looking at a water signal coming from a great mass of water, it has temporarily lost its ability to look sufficiently into the distance to provide adequate warning time should a serious CAT be approached.

Thus, even though the water related frequencies chosen for their long distance abilities work well most of the time, they can be frustrated by clouds at these high altitudes. When this type of blockage occurs, there will typically be a rather constant signal at the water wavelengths . It is then time to change to wavelengths which can penetrate the water better than the water related wavelengths. Carbon dioxide related wavelengths can be used to provide extended look distances in this case. In particular, the carbon dioxide wavelengths of 13.5 and 15.0 microns are advantageously used in accordance with the invention although in this regard any water based wavelengths and any carbon dioxide based wavelengths in the 1-30 micron range may be used when switching from water based to carbon dioxide based wavelengths where the selected wavelengths can be empirically determined.

As soon as the water wavelengths return to normal relationships, the system can switch back to them, but there will be extended capability when in high water situations. The best analysis will involve the use of both wavelengths.

Heretofore, the systems described have been airborne and may be termed Advance Warning Airborne Systems (AWAS) . Hereinafter, in addition to AWAS, there will be described stationary systems which may be termed Stationary Advance Warning Systems (SAWS) . There will also be described various systems where SAWS techniques may be used in AWAS systems. In each of these systems it is important to understand exactly what the infrared sensor "sees" in these devices. Assume an IR detector is sealed in a light tight box with its only contact with the outside being the brightness of a chalky window in the top of the box. It can only detect the brightness of different colors in the glow of the window. The system also has electrical signals from the following aircraft instruments to use in its calculations: airspeed; outside air temperature; pitch; and radio and pressure altimeters.

From this the system is required to measure the presence of dangerous cold fronts ahead of the aircraft, and to warn the pilot if they are severe. The system can see only one color at a time. It is the brightness, intensity, or power of the light of these several colors from which all of the optical- information must come.

Further, all of the light reaching the detector comes from a narrow detection cone typically only 2 degrees wide and 2 degrees high. This cone of acceptance is established by the optics of the system.

See Fig. 25 for the concepts regarding this cone, segments of the cone, the volumes of those segments, and the areas of the segment separators. These areas are spherical surfaces, but flat ones are shown, in that the differences are iniscule. Moreover, the air itself is providing the light that the system sees. It is not reflected sun light as one would see with their eyes. The system looks into a faintly glowing mass of air. The intensity of this infrared glow is a positive function of the air temperature, in accordance with Planck's law.

The total light power measured by the detector is the sum of the light power received from all distances. The distances that quanta of infrared glow can travel are limited by their tendency to be absorbed by the molecules that they contact in their travel. Once a quantum of light headed toward the detector is absorbed, the chances of its again being emitted within the acceptance angle that would strike the detector are very, very small. Thus in many cases the absorption of light can be seen to follow Beer's law:

P = PO x e^~(qxD) (7) where P is the light power reaching the detector from the light power PO originating from distance D- and beyond. q is a constant, called the absorbence. This is an exponential relationship that indicates that as the distance increases, the signal that is received from that distance is less than that from nearby. If the absorbence, q, is large, the system can see only a short distance. Conversely, if q is small for the wavelength used, the system can see for a great distance. The reading that the system obtains is the sum of the P values for all of the D values.

The PO value originating at each distance is a function of the temperature of the emitting gas at that distance. The total power reading is then a function of the temperatures of the gas at all distances that can be seen by the system using that particular IR wavelength. Significant changes of temperature at any of these distances will change the light power reading by changing the P values from these distances in the summation that reaches the detector. It is these changes that correspond to the delta T's of the equations discussed above and also utilized below.

As AWAS moves with the aircraft toward new temperature regions in its detection cone, they will provide larger and larger fractions of the light power readings, for less and less light from these regions will be sorbed on its way to AWAS. This provides a change in light power with time that in effect describes the contents of the cone, and thus may be an important part of the criteria for detecting dangerous conditions ahead of the aircraft.

It should be noted that the classical form of the Beer/Lambert law given in Eq. 7 is correct for most calculations. However, for measurements at frequencies which involve emission from C0₂, it should be noted C0₂ is a sufficiently strong sorber of these same frequencies, at the central frequency involved, that the detector receives only the light of the outboard frequencies, and this is controlled by the central frequency. Because this light is not free in the sense that a less self sorbing light would be, its form of the transmission equation is different. The form that is found to fit the data is shown as Eq. 1.

p = Po X e -J^{ q. d⁾' (7a)

This has taken the square root of the q*d in the exponent of e. The effect of this is a weakening of the slope of the transmission curve. Accordingly, the data taken for clear air turbulence modes used at high altitude and typically used for LLWS calculations employing Eqs. (3) through (5) may be so compensated.

With the foregoing concepts in mind with respect to the detection cone, various applications of the invention will now be described. The first of these will be lapse rate correction. That is, the temperature of the atmosphere is routinely a linear function of the altitude. The slope of this temperature versus altitude curve is the lapse rate. This value is typically between 5 and 8 degrees Celsius per km.

In the foregoing descriptions, it is assumed the system is aimed parallel with the ground, such that its signals represented a most nearly constant temperature layer of atmosphere. Not always will this situation be available. For example, during the take off of a commercial jet, the rotation may be as much as 30 degrees. AWAS sees the effect of a different temperature at each distance in its cone. This looks like a cold front with ever colder temperatures as the aircraft gains altitude. Thus, since the various index formulae of the present invention are typically based on an horizontal orientation of the aircraft, the measured value of Δ T should be compensated with a default (predetermined) lapse rate or a measured lapse rate whenever the measurement is made at an angle with respect to the horizontal as described hereinbefore with respect to block 212 of Fig. 12. The manner in which this compensation is effected will now be described with respect to Figs. 26 and 27.

It should also be noted that, when the upward looking detector is employed for an index determination such as the stationary LLWS determination discussed hereinafter, the upward angle of 2 to 4 degrees of this detector must be used. This 2 to 4 degree angle should be added to the pitch of the aircraft.

Referring to Figure 26, there is illustrated the technique used to determine the look distance of the detector where the look distance is the distance the detector 81 or 98 sees into the atmosphere to make the Δ reading, the look distance being a function of the sensed wavelengths and the sensed wavelengths being a function of the alternate grating positions.

The angle o is the angle between the detectors and is typically 2 to 4 degrees. The temperature measured by the detector directed along the line of flight is T₂ and the temperature measured by the upwardly directed detector is T^. The distance between these temperatures is B and the lapse rate is (T₂ - T_Q)/lkm where 1 km is the distance between T₂ - T_Q. A typical default value of the lapse rate is 6.5" Celsius/km.

The foregoing default value of the lapse rate is typically used during take off and, in particular, below 1500 feet, for example, where LLWS determinations are made. During descent, a calculated or measured (as opposed to default) value of the lapse rate is typically employed inasmuch as there is sufficient time to effect such calculation during descent. In general, this calculation is effected between 15,000 feet and 1,500 feet by obtaining the difference between successive readings of the OAT temperature and dividing them by successive differences in altitude readings where the OAT and altitude measurements are made by conventional measuring devices on the aircraft.

Referring to Figure 27 , it can thus be seen that the lapse rate obtained at step 300 of Fig. 27 is typically a default lapse rate if the aircraft is ascending and a measured lapse rate if descending.

The look distance (LD) of step 302 can be obtained as

where B and tan oζ are defined in Fig. 26 and where B = (T₂-T₁)/(T₂ - T_Q) (9) where T₂ and T-_j_, are measured and T₂-T_Q = 6.5" Celsius as described above, by default, or is as measured on descent.

The pitch angle of step 304 is obtained from the aircraft instrumentation where the angle -C should be added thereto if the index determination is made using the upwardly directed detector (which is preferably the case in the runway scan mode) , as discussed above.

The error (or correction) temperature, T_c , is calculated at step 306 as the product of the lapse rate obtained at step 300, the look distance of step 302, and the tangent of the pitch angle of step 304.

The error temperature is added at step 308 to the measured / T of Fig. 12 to obtain the lapse rate corrected ^ T' of Fig. 12 if the aircraft is ascending and subtracted if descending.

From the foregoing, it can be seen the twin set of detectors preferably see light from a difference in height direction of approximately 2 to 4 degrees. One is set to look directly along the line of flight, and the other looks up from this by 2 to 4 degrees. This is a sufficient difference in height direction that a clearly different temperature is seen. From Fig. 26, it can be seen if the difference is very large, the look distance is also very large, and the two sensors collect light from over a large distance. If-, the temperature difference is very small, both detectors are seeing essentially the same air. Thus, the look distance is very short. This measure of how far the system can see can be important in interpreting the meaning of the time-temperature differences. If the look distance is short, the temperature difference is local, and the magnitude can be easily interpreted. If the look distance is long, it is more difficult to- interpret the temperature difference. Its change in magnitude with time gives the information as to size.

In the foregoing applications of AWAS, the aircraft airspeed has not been significantly interrelated with the observations. It can play an important role, for it establishes the rate at which the detection cone moves into the weather ahead of the aircraft, and thus, the rate that its picture can change.

See Fig. 28 which illustrates the signal change per km of travel toward a front. The signal change per time values become distance derivative values when divided by km/hour, the air speed. It is the pattern and magnitude of the signal distance derivative that may play an important part in the analysis as to whether a dangerous front is being approached. The greater the aircraft velocity, the greater the time related signal. Division by airspeed provides a velocity independent parameter for analysis.

When a stationary unit is considered, or SAWS, the motion of the device is zero, and cannot be used to generate the derivative signal from the detector output. Moreover, division by the airspeed can not be used for division by zero explodes the results.

Attempting to use wind velocity as the speed that the front moves is also inappropriate, for there is often no relationship between local windspeed and the velocity of the front seen in the distance. Windspeed can change sign, or direction, and also be zero. Thus, local air motion can not be included in the interpretive equation for SAWS, for it provides no consistent information regarding what the SAWS sees in the distance.

At least two classes of SAWS can be established: Process the derivative of the cone total- SAWS-I Process cone total - SAWS-II

SAWS-I sees only the time derivative of the event that enters its cone. This corresponds to the A A T signal of Fig. 12. See Fig. 29 for a typical signal form. This device sees a change enter its field of view, leave its field of view, or change within its field of view. It gives no information as to where it is, or what direction it is moving since the detector is relatively stationary.

This first type may encounter difficulty with noise, especially if the noise contains items of the same magnitude and frequency as the entry and exit events. Integration to reduce the noise does not work well here, because the entry and exit events have no aspect to differentiate them from the noise.

SAWS-II differs from SAWS-I by its ability to repeatedly or continuously obtain information about the sum of the temperature profile in its cone of vision without anything in the cone needing to change. This corresponds to the signal of Fig. 12. Fig. 29 contains a typical signal/time form.

It should be noted that AWAS can be used as a SAWS-II when the aircraft is on the runway ready to take off. Because the aircraft velocity in the direction AWAS is seeing is zero, the signal level changes with time only in response to a changing temperature somewhere in its cone of view. The sum of the temperature-distance function (temperature profile) is measured repeatedly (every 2 seconds, for example), so that the measurement is not just a time derivative function. This use of SAWS-II technology makes it possible for AWAS to check the safety of the flight path before take off. To operate in this runway scan mode, AWAS must substitute a finite default air speed, or change equations to avoid division by AS=0. Keeping in mind that the foregoing alternative embodiments may be used, an illustrative embodiment will now be described of a runway scan system wherein a different type of approach, is used which does not involve division by the air speed.

Referring to Fig. 30, the T ' signal of Fig. 12, which is applied to SAWS II index calculation 244, is applied, in particular, to a clipping circuit 320, the purpose of which is the same as those of clipping circuits 218 and 230 of Fig. 12. Six stage FIFO, summing circuit 324, and divider 326 obtain an average of six successive ΔT readings to thereby insure a single reading will not trigger an alarm. The number of successive readings is arbitrary and depends on the degree of confidence described before triggering an alarm and on the amount of time needed by a pilot or the like before the event is encountered. Since in the runway scan mode the amount of time before the event is encountered depends on when take off occurs, whatever time is necessary to provide the desired degree of confidence may be taken before it is decided to take off.

The output of divider 326 is compared with a threshold value 330 at comparator 328 to provide a warning 332 if the average value of A T ' exceeds the threshold value where the threshold value is empirically determined to correspond to presence of LLWS in the intended flight path, for example, where the threshold value is based on aircraft type.

A more sophisticated runway scan system which is capable of detecting a variety of conditions is illustrated in Fig. 33. In particular, this runway scan mode can :

1. Sense already formed microbursts ahead;

2. Sense forming microbursts ahead;

3. Sense forming microbursts engulfing A/C from ahead;

4. Sense formed microbursts engulfing A/C;

5. Sense forming microbursts engulfing A/C from behind;

6. Signal OK or Danger.

Figure 33 is a general logic diagram for the AWAS system to meet all of these situations. The system involves the computation of a distant temperature, TDC that is expected, and the comparison of this with the measured one, TD. In addition, the change of the local temperature, TL , and the change of the distant temperature, TD, with time over the test time period are monitored and judged against constants. Finally, the dual detector lapse rate measurement is shown to be reasonable.

For each concerning weather condition, the logic scheme is now described.

1. Sense already formed microburst ahead. On the runway scan signal the computed and measured distant T's will be obtained and compared. The comparison should fail with the Rl limit. The distant TD will be smaller than the calculated TDC due to the microburst.

2. Sense forming microbursts ahead. The Rl test may be failed if the microburst is sufficiently formed, but may pass. If it passes, the routine passes through to the R3 test which looks for up to a 30 second change in the distant temperature. If there is no change in the distant temperature over the R3 limit, the lapse rate is measured using the two detectors. If this gives a reasonable value, the runway scan tests continue for up to the full 30 seconds, and are then passed. R3 is the major test for sensing this condition.

3. Sense forming microbursts engulfing A/C from ahead. The Rl test should fail, but it cannot be guaranteed that it will if the local temperature decline has occurred more than 30 seconds before the runway scan signal was given. The R2 test looks for changes in the local temperature. R2 will detect a microburst just reaching the A/C as the tests are being conducted, as should Rl, and it will also detect the usual changes in temperature within the microburst. R3 should see the heart of the microburst, unless it has now moved over the A/C. If the problem has not been detected by this time, the lapse rate is measured, and should be quite out of bounds. All tests may be continued for up to 30 seconds for example. A failure of any of them will provide warning during that period.

4. Sense formed microbursts engulfing A/C from side or above. This situation should decrease the local temperature, but would not tend to influence the distance temperature. This should fail the R4 test if there is very little effect on the distance temperature. If both TL and TD change, the R2 test will fail, as should R3. The lapse rate should also be unusual. Note another type of lapse rate test is disclosed in U.S. Patent No. 4,947,165.

5. Sense forming microbursts engulfing A/C from behind. In this case, the local temperature, TL, would drop with time. This would fail R2. Because of the TL finite time delay .-for Rl, this test might not fail during the test period.

Thus, to perform these tests, there has been used a computed value for the distance temperature compared with the measured value. This can fail Rl if we are looking into a microburst, and R4 if we are looking out of one. The calculated value is based on a local temperature measurement made 30 seconds before the calculation. This helps the system see the edge of a weather system forming ahead of it.

The system also looks at the stability of the local temperature, failing R2 if this is becoming cooler. This will be the case if the A/C is becoming engulfed from any direction.

Also, the stability of the distant temperature is measured and judged, failing R3 if the temperature is increasing, and R5 if it is decreasing. R3 failure could imply an engulfing microburst moving in opposition to the intended A/C motion. R5 failure could imply an engulfing microburst moving in the intended A/C direction, or one forming at a distance near the look distance.

Passing all of these tests could be possible under only a few cases in a million where a down draft storm compensated quite accurately for its temperature differences-altitude-lapse rate combinations to escape the warning situation. Thus, there is added the lapse rate measurement to see the picture at two different angles. The chances of the sensed lapse rate also being within bounds in such a situation is very small. Thus, the chance for safety is increased beyond that of the more direct tests.

It should be emphasized that one of the tests described above is the situation where the downdraft part of a microburst occurs behind the aircraft when the aircraft is ready to take off. This results in a sudden tailwind that can surround the aircraft as it attempts to gain speed. The result is the inability to accelerate the air speed adequately before the end of the runway is reached. A number of cases of this type have been reported. The wind sock is not readily available once the roll is started. The first clue is the inability to gain air speed.

The momentum (or inertial) type windshear devices do not detect this type of phenomenon. However, AWAS can detect the unexpected temperature change. Especially if the runway check uses the bi¬ directional system, it can spot either a fore or aft microburst just prior to take off from considerable distances. Even a system using only the OAT and the appropriate SAWS electronics or computer can detect a relatively close fore, aft or side microburst by analyzing the temperature-time function as the aircraft approaches the runway.

The clue to these dangerous rear-side microbursts is a very rapid loss of more than 3 to 4 degrees Celsius in local temperature, with the temperature remaining cold. Detecting such a situation is not difficult, but it is not an expected thing for a pilot to look for in addition to his other duties. Thus the local temperature may be monitored by AWAS as the aircraft taxies to its take off point to analyze this data to see if sudden negative changes in temperature take place. If all of the conditions are correct, a warning is sounded.

The SAWS-II measurement would also tend to indicate that the temperature profile ahead of the aircraft was not as expected. This would occur because the local temperature would cause AWAS to predict a lower cone total than would be sensed.

The runway warning system would thus be sensitive to aft microbursts by reporting too high a cone total, and fore microbursts by reporting too low a cone total.

As also discussed above, one test that the above runway scan embodiment utilizes is a determination as to whether the lapse rate is within a predetermined range. If not, a low level windshear indication may be provided. In the thermal-reactive embodiments of the invention, lapse rate correction cannot be effected in the foregoing manner since in the thermal-reactive embodiments, only the OAT measurement is utilized. Accordingly, in the thermal-reactive embodiments for take off and landing, lapse rate compensation is effected by adjusting the critical value of the hazard index. As described above this value is typically 0.15 for LLWS detection. Accordingly, in thermal-reactive systems the critical value of the hazard index is increased to 0.17 when taking off and adjusted to 0.13 when landing to thereby provide the desired lapse rate correction for such systems.

Two pitch angles may be employed as a result of dual detectors. As discussed above with respect to lapse rate correction, these make possible a lapse rate measurement using a default value for the look distance or vice versa. Not indicated on the logic chart is the possibility of performing the full logic of Fig. 12 or Fig. 33 with data from each of the detectors. This multiplies the probability of the results being correct.

It is important that the landing aircraft not be warned away from the runway if a weather event is present significantly beyond the other end of the runway. It is also important that the aircraft not be aborted in take off because of an event that is so far away that the aircraft would have climbed to over 1500 feet before contacting it. Both of these situations could be considered nuisance warnings even if the system indeed reported a real event of significant magnitude. Thus, in accordance with a further aspect of the invention, the look distance of the longer seeing wavelength is chosen to be about 4 to 5. km, which insures that the aircraft can see only to where the aircraft would reach the 1500 foot altitude, below which LLWS is a problem, when taking off without seeing substantially beyond where the aircraft would reach 1500 feet and when landing the aircraft, as it reaches 1500 feet, sees to the runway but not substantially beyond the runway.

In a runway system which utilizes an LLWS detection formula of the type of Eqs. (3) through (5), a default airspeed of 135 knots, for example, may be employed in lieu of the actual zero airspeed. Typically, the hazard index is calculated every two seconds and judged against an empirical Ft value, or an average thereof can be calculated based on six, for example, samples using averaging circuitry such as elements 322-328 of Fig. 30. Assuming the average over this twelve second period (Each sample taken every two seconds. ) exceeds the empirically determined threshold 330, a warning 332 is provided.

Various applications of AWAS will now be discussed. Of course, LLWS detection, as extensively discussed hereinbefore, is important. Moreover, with respect to the stationary case discussed above, where the aircraft looked down the runway just prior to take off, a SAWS-II measurement is described which gives safety information prior to starting to roll. Typically, AWAS, when utilizing Eqs. (4) or (5) , requires a 60 mph roll speed before AWAS can accurately look down the runway, because of the inability of the airspeed detector at lower speeds to derive the requisite derivative data. Thus, the system may switch to AWAS operation at 60 mph, but use SAWS-II from 0 to 60 mph as an added safety feature.

Another application of AWAS relates to jet stream detection. The jet stream is a fast complex layer of wind blowing in a west-east general direction across at least the USA and the Pacific ocean. The velocity of this moving stream of air is often between 100 and 300 mph. A typical local cross-section is shown in Fig. 31.

An aircraft that can find its way into the pocket on either side of the jet stream can be carried smoothly with the jet stream at very high velocity. The velocity relative to the ground is the sum of the aircraft airspeed plus the velocity of the jet stream. The savings of fuel on a single crossing of the Pacific ocean for a 747 may be about $8000, depending on the cost of fuel.

Thus, it is very desirable to have a sensor which can help the pilot locate the jet stream and the smooth region in it. On the return trip it is critical for the aircraft to stay out of the jet stream. AWAS can provide both of these functions, as described below.

The jet stream is cooler than the surrounding air by many degrees. The transition to the warmer air is very turbulent. Thus, the CAT mode of AWAS (using the standard deviation) can sense the edge of the jet stream when the turbulence intersects the cone. Where both the aircraft and the jet stream are traveling in about the same direction, there is relatively little chance of seeing the jet stream with the conventional forward looking AWAS.

Accordingly, as illustrated in Fig. 32, AWAS can be set to scan right and left and up and down relative to the aircraft and thus use CAT and/or SAWS-II modes to see the jet stream. Alternately, if AWAS is mounted low on the aircraft, some downward search can be added.

Block 242 of Fig. 12 may be employed to detect the jet stream if the standard deviation technique of CAT detection is employed. Note in this regard that the jet stream typically occurs at altitudes above 35,000 feet while unrelated CAT does not. Hence, the altimeter reading may be used in conjunction with the CAT detection output to differentiate the detection of CAT from detection of the jet stream.

Moreover, block 244 of Fig. 12 and, in particular, the circuitry of Fig. 30 may be employed to detect the jet stream if the SAWS II technique is employed. Here the threshold 330 of Fig. 30 would- be empirically determined to correspond to the presence of jet stream where again the fact that the jet stream typically occurs above 35,000 feet may be employed to differentiate the detection of it from other phenomena.

In utilizing the circuitry of Fig. 30 (or the CAT detection 242 of Fig. 12) , means are provided, as is known to those skilled in this art, whereby the measured

T signals are correlated with the direction of orientation of the scanned detecting means so that the direction of the detected jet stream with respect to the aircraft may be determined.

With reference to Fig. 32, it can be seen that outside mirror 20 is tiltable through an angle eC by a tilt motor 22 and is provided with an electric heater H to prevent frost and condensation and, most importantly, ice from collecting on its reflective surface. Mirror 20 and motor 22 are mounted upon a rotatable ring 24 which closes an opening in the aircraft skin surrounding window 71 of the IR spectrometer. Ring 24 is carried upon slide contacts 26 which enable power to be transmitted to the tilt motor 22 and heater H. A ring gear 28 engages the rotatable ring 24 for producing rotation of the ring 24 (and with it the mirror 20) about window 71, when driven by rotation motor 30. As can be appreciated, such a search unit enables the infrared radiation to be directed through the mirror 71 into the IR spectrometer from virtually any direction due to the combined effects of the 360 degrees of rotation of the mirror about the axis of rotation of ring 24 and tilting of the mirror in a plane of tilting movement that is perpendicular to the axis of rotation. The outside mirror system of Fig. 32 with rotational capability and tilt control provides standard AWAS functions plus rotational and vertical jet stream search capability while using a standard optical bench in the system. A limit system may be employed so that the AWAS does not look at parts of the aircraft, or its exhaust stream. The exhaust stream typically falls and spreads away from the aircraft, such that searching horizontally or overhead would not contact it.

This multidirection AWAS is able to search an approximate 100 mile diameter for signs of jet stream activity. Once such activity is detected, the aircraft may move toward it. If up-down capability is available on the AWAS, the center of the wall can be estimated for penetration. If this up-down capability is not present, the aircraft roll can be used to scan the wall. Once inside the smooth column, the left-right CAT mode can help keep the aircraft centered in the jet stream flow.

Another application of AWAS relates to aircraft lead/follow protection. The lead/follow protection unit also utilizes the outside mirror system of Fig. 32 and searches for the heat given off by the engines and exhaust of other aircraft, and makes certain the pilot is notified of any aircraft detected. The search is for a positive rather than a negative entity. The major search is for the exhaust of an aircraft ahead. The after search is complicated by the aircraft's (A/C's) own exhaust. However, exhaust and turbulence from a large aircraft are typically rolled downward by the action of the wings at a rate of approximately 900 feet per minute. This does then provide the opportunity to look backwards for following aircraft whose engines or exhaust, or turbulence can give clue to their presence. A SAWS-II mechanism gives a measure of their distance and rate of approach. Accordingly, the circuitry of Fig. 30 may be employed where the empirically determined threshold value corresponds to the presence of another aircraft.

On the ground, the use of SAWS-II permits the detection of the turbulence of warm exhaust remaining from a recent aircraft motion. Under some weather conditions, severe disturbance can remain for several minutes or so after the passing of a large aircraft. These winds can be enough to flip a smaller aircraft, or otherwise damage it. For example, a small aircraft looking down a runway too soon after the take off of a large jet can see that ahead there was a dangerous situation that it should avoid. Present small aircraft procedures involve great care and prolonged waits before moving where their big brothers have been. There is always a danger that these "standard times" are not sufficient in a given situation.

Crossing the path of a larger aircraft in the air can also be a difficult experience if the turbulence has not subsided. The lead/follow function can also detect the temperature and turbulence ahead, and provide warning.

^• Thus, the lead/follow protection can serve an important warning function both on the ground and in the air.

Another application of AWAS relates to dust/smoke detection. Aircraft engines can suffer extremely costly damage from prolonged exposure to smoke, dust, and other particulates. Smoke and dust such as that from a volcano or forest fire may destroy the engines on aircraft flying over head at great altitudes where the pilot cannot see the pollution in the air. For example, the Mt. St. Helen's volcano eruptions destroyed many millions of dollars worth of jet engines on the aircraft that tried to fly through the western areas of the country during the several months of that disaster. There were wind streams containing these particulates thousands of miles from the mountain.

Volcanic ash from the recent Redoubt Volcano in Alaska has been blamed for the jet engine failure on a flight 115 miles from the volcano. All four engines stopped when the 747 flew through the stream of ash that was not detected by radar. Fortunately the engines were restarted in time to avoid a crash.

Air polluted with particulates scatters the infrared light that AWAS would normally collect. Thus, AWAS will see a much smaller IR light level than is appropriate for that air temperature. This will look like windshear when approached from below 2000 feet. It would not typically trigger the CAT warning above 20,000 feet. However, when entering the pollution, correlation with the measured outside air temperature (OAT) would fail. When this correlation failure is one of measured air temperature (OAT) being larger than IR detected air temperature, there is danger of particulates in the air.

A particulate detecting version of AWAS may include (a) a mechanism to interpret the OAT-IR Temp correlation, (b) a "Particulates" warning system, and (c) means for powering the warning system in response to the dangerous values of the correlation. An LLWS type of measurement would serve at all altitudes.

In addition to the OAT-IR system, the wobble system of Fig. 12 may be employed where the SAWS II ^ T' signal is applied to index calculation circuitry 244 of the type illustrated in Fig. 30 where the empirically determined threshold value corresponds to a predetermined level of particulate matter such as volcanic ash.

A further application of AWAS is local temperature detection. AWAS units are capable of reading temperatures of local objects when the object is present in the cone. By combining this with the wide range of aiming that was described under the Jet Stream application with respect to Fig. 32, it is possible for AWAS to look at specific objects on the aircraft, or on the ground, and determine their temperatures. Calibration for a wide range of temperatures involved can provide accuracy to plus or minus several degree Celsius.

The derivative function provided by aircraft motion does not exist when looking at parts of the aircraft, thus, a SAWS-II mechanism should be employed. This function of the system can be used to test the temperature of the engines on a large aircraft in flight in quite a different way than the usual thermocouples, etc., in that no new wires are needed.

Thus, using a system of the type illustrated in Fig. 12 to obtain T' measurements and with the calibration of these T' values to local object temperatures, the ' readings will correspond to the requisite temperature readings.

SAWS II applications will now be described where the system is truly stationary with respect to the earth. As can be appreciated from the foregoing, many of the AWAS applications utilize SAWS II. Any measurement taken where there is no certain motion between the IR sensor and the object or atmosphere being sensed can be interpreted by a SAWS II system.

The above applications were included under the AWAS listing because they were aircraft borne applications, rather than being truly stationary.

A land-based LLWS detection SAWS-II technology may be used to detect dangerous cold fronts crossing the runways in either vertical or horizontal modes using the various runway scan modes described above. A system with a mirror of the general type illustrated in Fig. 32 capable of scanning the runway approach, exit, and overhead provides a very meaningful system for giving warnings of danger to either a towerman, or for radio broadcast to approaching or take off prepared aircraft.

Multi-sensor systems may also be used for simultaneous measurement in the various directions. Thus a single unit can be used for each runway direction to provide the fastest response to a microburst. Many different combinations can be established to protect especially the smaller airports which can not afford huge weather radar systems. These smaller airports may also host many of the smaller aircraft which may not use the AWAS systems due to their cost. Thus, the need here for an airport type of warning system is great. Where there is no tower or contact radio, the SAWS system may provide a go/no go light system for take off and landing advisory. The SAWS could also contain its own warning radio system.

The immobile SAWS systems may use somewhat different algorithms for establishing conditions of concern. Because of their greater exposure to conventional ground/weather systems, their sensitivity to change may be based upon more thorough and exacting critera, for this is where most of the reported measurements have been made.

The SAWS systems can also detect smoke, dust and other air pollution that can be dangerous to aircraft engines, especially the jet engines. Sufficient warning to avoid an engine stopping at a critical part of a take off or landing could be most important.

The positive potential for these ground-based systems is very great. Only aircraft, weather, and air transport applications have been presented herein but other applications are contemplated to be within the scope of this invention.

Claims

WHAT IS CLAIMED IS:

1. A method of detecting weather related events or other temperature related phenomena with respect to an aircraft comprising the steps of: determining whether the aircraft speed exceeds a predetermined value of speed; determining, in response to the aircraft, speed exceeding said predetermined value, whether the altitude of the aircraft exceeds a first predetermined value; determining, in response to the altitude exceeding the first predetermined value, whether a clear air turbulence (CAT) index is exceeded and issuing a warning if it is exceeded; determining, in response to the aircraft not exceeding the first predetermined value, whether the altitude is less than a second predetermined value; determining, in response to the altitude being less than the second predetermined value, whether a low level windshear (LLWS) index is exceeded and issuing a warning if it is exceeded.

2. A method as in claim 1 where said step of determining the low level windshear index includes determining, in response to the altitude being less than said second predetermined value, whether the aircraft acceleration exceeds a predetermined value of acceleration so that said LLWS index determination can be made based on (a) the aircraft taking off if the acceleration exceeds the predetermined value of acceleration or (b) the aircraft landing if the acceleration is less than the predetermined value of acceleration.

3. A method of detecting a weather related event or other temperature related phenomena comprising the steps of : obtaining a wobble signal by alternately detecting at a predetermined wobble frequency at least two temperature dependent signals emitted from the air where said signals have different wavelengths and where the magnitude of the wobble signal is a function of the difference in magnitude between the two signals and the frequency of the wobble signal is a function of the predetermined wobble frequency at which the two temperature dependent signals are alternately detected; and processing said wobble signal to detect said weather related event.

4. A method as in claim 3 where the first of the signals corresponds to a first distance from a location where said wobble signal is obtained and the second signal corresponds to a second distance from said location.

5. A method as in claim 4 where said wobble signal corresponds to the temperature difference between the temperature of the air at said first distance from the location and the temperature of the air at said second distance from the location.

6. A method as in claim 5 where the difference in the wavelengths is such as to provide said temperature difference.

7. A method as in claim 6 where the wavelength of the second signal is so chosen as to correspond to a predetermined altitude.

8. A method as in claim 7 where said predetermined altitude corresponds to an altitude below which low level wind shear is potentially dangerous.

9. A method as in claim 8 where said predetermined altitude is about 1500 feet.

10. A method as in claim 7 where the maximum look distance associated with said wavelength is about 4 to 5 km.

11. A method as in claim 4 where said first and second distances are near said location.

12. A method as in claim 4 where said first and second distances are far from said location.

13. A method as in claim 4 where said first distance is far from the location and the second distance is near the location.

14. A method as in claim 4 where said location is on an aircraft.

15. A method as in claim 4 where said location is substantially fixedly positioned with respect to the weather related event.

16. A method as in claim 3 where said alternately detecting step is effected by alternately positioning a wavelength sensitive grating between first and second positions so that the wavelength of the first of the two signals is directed to a detecting means in response to the grating being positioned at the first position and the wavelength of the second of the two signals is directed to the detecting means in response to the grating being positioned at the second position.

17. An infrared ray spectrometer comprising: a wavelength sensitive grating; means for directing infrared rays to -.the grating ; detecting means responsive to the infrared rays f or produc ing s i gnal s c orre sponding to the infrared rays ; and pos i t i oning me an s f o r al te rnate ly positioning the grating at at least first and second predetermined positions so that the wavelength of a first one of said inf rared rays is directed to said detec ting means in re sponse to the grating being positioned at the first position and the wavelength of a second one of said infrared rays is directed to said detec ting means in response to the grating being positioned at the second position; whereby a wobble signal is produced at the output of the detecting means , the magnitude of which is a function of the difference in the magnitudes of the first and second infrared rays and the frequency of which is a function of the f requency at which the grating is alternately positioned at said first and second positions .

18 . A spectrometer as in claim 17 where said positioning means includes a stepping motor with a shaft with said grating mounted thereon and means for alternately positioning the shaft and the grating mounted thereon between said f irst and second positions .

19 . A spectrometer as in claim 18 including damping means mounted on the shaft for damping vibrations of the grating as it is alternately positioned at the first and second positions , said damping means comprising a multi-chamber element mounted on the shaft where each chamber contains a predetermined amount of damping elements which impact against the wall of their associated chamber as the grating reaches one of said first or second locations to thus effect the damping.

20. A spectrometer as in claim 19 where said damping elements comprise lead shot.

21. A method of detecting low level wind shear (LLWS) with respect to an aircraft comprising the steps of: obtaining the difference in temperature between first and second temperatures of the air outside the aircraft where the first temperature is that of the air located a first distance from the aircraft and the second temperature is that of the air located a second distance from the aircraft; determining the value of a hazard factor F which is a function of at least (a) a first term which is a function of at least A T and (b) a second term which is also a function of at least Δ ^τ where the functions of Δ T of the first and second terms are different and can not be linearly combined; comparing the determined hazard factor F with a predetermined threshold value, F_τ, representing the maximum value of F permissible; and issuing a warning if the determined value of F exceeds F_τ.

22. A method as in claim 21 where the function of T in the first term is non-linear.

23. A method as in claim 22 where the non-linear function is a derivative of ΔT.

24. A method as in claim 23 where the derivative of ΔT is with respect to distance.

25. A method as in claim 23 where the derivative of T is with respect to time.

26. A method as in claim 25 where the hazard factor is defined as

F = K*d(ΔT)/dt + J/A_s * /(ΔT)'

where K and J are constants, A_s is the speed of the aircraft.

27. A method as in claim 23 where a predetermined number of successive values of the derivative of Δ T are obtained and averaged to obtain the value of the derivative employed in the first term of the hazard factor.

28. A method as in claim 27 where said predetermined number is three.

29. A method as in claim 21 where the function of Δ in the first term is linear.

30. A method as in claim 29 where the hazard factor is defined as

F = K* ΔT + J/A_s * J (ΔT η

where K and J are constants, A_g is the speed of the aircraft.

31. A method as in claim 21 where the function of Δ T in the second term is non-linear.

32. A method as in claim 31 where the non-linear function is the square root of at least ΔT.

33. A method as in claim 21 where the function of T in the second term is linear.

34. A method as in claim 21 where the second term is an inverse function of A_s , the speed of the aircraft.

35. A method as in claim 21 where A is measured by wobbling.

36. A method as in claim 21 where Δ^τ s obtained by measuring each of the first and second temperatures and then calculating Δ by subtracting one of the measured temperatures from the other.

37. A method as in claim 36 where each of said temperatures is measured by a bow tie detecting element.

38. A method as in claim 36 where one of said temperatures is obtained by means for measuring the air immediately outside the aircraft.

39. A method as in claim 21 where said ΔT in the second term of the hazard factor is obtained by summing a predetermined number of samples of a time derivative of Δ to thus obtain substantial smoothing of the aforesaid ΔT used in the second term.

40. A method as in claim 39 where said predetermined number of samples of the ΔT derivative are applied to a first in, first out register means having said predetermined number of storage units and where each unit of the register is applied to a summing means to effect said summing step.

41. A method as in claim 40 including detecting whether the aircraft has encountered a front rather than low level wind shear and substantially zeroing the contents of all storage units of the first in, first out register means upon detection of said front.

42. A method as in claim 41 where said front is a cold front.

43. A method as in claim 41 where said detecting step is effected by determining whether an average of the derivative of ΔT is substantially within Z of zero for X seconds where Z and X are predetermined numbers.

44. A method as in claim 43 where Z is a function of the thermal noise level of the air outside the aircraft.

45. A method as in claim 40 including detecting that the aircraft has entered a low level windshear (LLWS) event and adding the value of the sample of the

Δ derivative most recently outputted from the first out, first in register means to the value of the Δ derivative about to read into the register means upon detection that the aircraft has entered the LLWS event.

46. A method as in claim 45 where said detecting step is effected by determining whether the average of the derivative of for a predetermined period of time is greater than the noise band of the air outside the aircraft.

47. A method as in claim 40 where said derivatives of ΔT may assume either plus or minus values.

48. A method as in claim 47 where said values are obtained by clipping the derivative of Δ T .

SUBSTIT

49. A method as in claim 48 where all storage units of the first in, first out register means are zeroed in response to the sum of the samples crossing zero.

50. A method as in claim 39 where said predetermined number is twenty.

51. A method as in claim 21 where said Δ in the second term of the hazard factor is obtained by integrating a derivative of ΔT .

52. A method as in claim 21 where said hazard factor is F = K*d(ΔT)/dt (+/-) J/A_s *JΪ£d ( Δ T) /dt/' where K and J are constants and where the sign of the second term is determined by the sign of d (ΔT)/dt where ^d( ΔT)/dt is the sum of a predetermined number of d( T) dt values.

53. A method as in claim 21 where T is obtained from infrared measurements so that said hazard index formula may be used in an infrared - predictive system for detecting LLWS.

54. A method as in claim 21 where ^τ s obtained from successive outside air temperature measurements made by temperature measurement means mounted on the aircraft so that the hazard index formula may be used in a thermal-reactive system for detecting LLWS.

55. A method as in claims 53 or 54 where a further value of said hazard index is obtained by an inertial system.

56. A method as in claims 21 where a first value of the hazard index is obtained by an infrared- predictive system and a second value thereof is obtained by a thermal reactive system to provide enhanced accuracy in determining the hazard index.

57. A method as in claim 56 where a third value of said hazard index is obtained by an inertial system.

58 . A method of obtaining information about weather related events or other temperature related events comprising the steps of : providing at least one detecting means to obtain a temperature difference signal indicative of the dif ference between ( a) an air temperature at a distance remote the detecting means and (b) the air temperature at a distance near the detecting means ; pos itioning the detec ting means in a substantially stationary position with respect to the event; and processing the temperature difference signal while the detecting means is in said stationary position to obtain said information.

59. A method as in claim 58 where said detecting means is located on an aircraft and said processing step includes determining whether said temperature dif ference signal exceeds a first predetermined threshold.

60 . A method as in claim 59 where said first predetermined threshold corresponds to the potential presence of already formed low level windshear at said remote distance from the aircraft.

61. A method as in claim 59 where processing step include s determining whether said temperature difference signal is less than a further predetermined threshold.

62. A method as in claim 61 where said further predetermined threshold corresponds to potential lower level windshear engulfing the aircraft from the side or above.

63. A method as in claim 59 including obtaining a value of the change of said temperature difference signal with respect to time and determining whether the last-mentioned value exceeds a second predetermined, threshold.

64. A method as in claim 63 where said second predetermined threshold corresponds to the potential presence of potential low level windshear currently forming at said remote distance from the aircraft.

65. A method as in claim 63 including determining whether the value of the change of said temperature difference with respect to time is less than a further predetermined threshold.

66. A method as in claim 65 where said further predetermined threshold corresponds to (a) low level wind shear potentially moving in the intended aircraft motion or (b) potential low level wind shear forming at said remote distance from the aircraft.

67. A method as in claim 59 including obtaining a value of the lapse rate where the lapse rate is the slope of the temperature versus altitude curve of the air and determining whether said value of the lapse rate is outside a predetermined range.

68. A method as in claim 67 where said lapse rate being outside the predetermined range corresponds to the potential presence of low level wind shear currently forming at said remote distance from the aircraft.

69. A method as in claim 59 including obtaining a value of the change of said temperature near the aircraft with respect to time and determining whether the las t-mentioned value exceeds a further predetermined threshold.

70 . A method as in claim 69 where said further predetermined threshold corresponds to potential low level windshear engulfing the aircraft from ahead.

71. A method as in claim 58 where said detecting means is located on an aircraft and where the method includes determining whether the speed of the aircraft is less than a predetermined speed and using said processing step , in response to the aircraft speed being less than the predetermined speed , to detect whether low level windshear is present and using an alternative proces sing step , in response to the aircraft speed being greater than said predetermined speed , to detect whether low level windshear is present.

72 . A method as in claim 71 where said alternative processing step includes at least taking the derivative of the detecting means output signal.

73 . A method as in claim 72 where said predetermined speed is about 60 knots (nautical miles /hour) .

74 . A method as in claim 58 including taking a time derivative of said temperature difference signal and where the time derivative is processed to obtain said information.

75. A method as in claim 58 where said stationary position is fixed with respect to the earth.

76. A method as in claim 58 where said detecting means is located on an aircraft having an airspeed A_s where A_R ^ A_g^O) where A_R is a predetermined reference airspeed.

77. A method as in claim 76 where the predetermined reference airspeed is about 60 knots.

78. A method as in claim 58 where said detecting means is located on an aircraft and said processing step includes determining whether potential low level wind shear is present by calculating a hazard index F where F is a function of at least (a) a first term which is a function of at least Δ T where ΔT is said temperature difference signal and (b) a second term which is also a function of at least ΔT and an inverse function of the speed of the aircraft where the functions of Δ of the first and second terms are different and can not be linearly combined; establishing a predetermined value of the said speed of the aircraft in lieu of the actual speed of the aircraft which is substantially stationary with respect to the weather event; comparing the calculated hazard factor F with a predetermined threshold value, F_τ, representing the maximum value of F permissible; and issuing a warning if the determined value of F exceeds F_τ.

79. A method as in claim 78 where said predetermined value of the speed of the aircraft is about 135 knots.

80. A method as in claim 78 where the function of T in the first term is a time derivative of ΔT.

81. A method as in claim 58 where said temperature difference signal is obtained by wobbling means.

82. A method as in claim 58 where said temperature difference signal is obtained by measuring each of the remote and near air temperatures and then calculating said temperature difference signal by. subtracting one of the measured temperatures from the other.

83. A method as in claim 82 where said temperatures are obtained by a bow tie detecting element.

84. A method as in claim 82 where the near temperature is obtained by means for measuring the air immediately outside the aircraft.

85. A method of detecting a temperature dependent phenomenon from an aircraft comprising the steps of: providing at least one detecting means to obtain a temperature difference signal indicative of the difference between the (a) the air temperature at a distance remote the aircraft and (b) the air temperature at a distance near the aircraft; directing the detecting means in different directions with respect to the aircraft; and processing said signal to determine whether the detecting means has been directed toward said temperature dependent phenomenon.

86. A method as in claim 85 where said processing step includes determining whether the temperature difference signal exceeds a predetermined threshold.

87. A method as in claim 85 where said processing step includes obtaining a deviation of said temperature difference signal with respect to a reference deviation.

88. A method as in c_.aim 87 where said deviation is a standard deviation.

89. A method as in claim 85 where said directing of the detecting means is effected by moving a mirror disposed on the outside of the aircraft and associated with the detecting means.

90. A method as in claim 89 where said mirror is movable in orthogonal planes so that said directing can be effected both to the right and left of the aircraft and/or up and down with respect to the aircraft.

91. A method as in claim 85 where said processing step includes detecting a change in said signal that exceeds a predetermined threshold corresponding to the presence of the jet stream.

92. A method as in claim 85 where said temperature dependent phenomenon is the jet stream.

93. A method as in claim 92 where the jet stream is detected at altitudes higher than a predetermined altitude.

94. A method as in claim 93 where said predetermined altitude is about 35,000 feet.

95. A method as in claim 85 where said temperature dependent phenomenon is the heat given off by the engines and/or exhaust of other aircraft to thus provide a lead/follow protection capability.

96. A method as in claim 95 where said aircraft is of a first size and at least some of the other aircraft are substantially larger in size than the first aircraft to thus provide a capability for a smaller aircraft to avoid the exhaust remaining after a larger aircraft has passed through the intended path of the smaller aircraft.

97. A method as in claim 96 where the smaller aircraft is preparing for takeoff.

98. A method as in claim 96 where the smaller aircraft is in flight.

99. A method as in claim 85 where said temperature dependent phenomenon is the presence particles in the air which potentially are volcanic ash.

100. A method as in claim 99 where said particles are detected at altitudes lower than a predetermined altitude .

101. A method as in claim 100 where said predetermined altitude is about 35,000 feet.

102. A method as in claim 85 where said temperature dependent phenomena is an object on the aircraft or the ground.

103. A method as in claim 102 where said object is the engine of the aircraft.

104. A method as in claim 103 where the engine is tested while the aircraft is in flight.

105. A method as in claim 85 where said temperature dependent phenomenon is low level windshear behind the aircraft when it is ready to take off.

106. A method as in claim 85 where said temperature difference signal is obtained by wobbling means.

107. A method as in claim 85 where said temperature difference signal is obtained by measuring each of the remote and near air temperatures and then calculating said temperature difference signal by subtracting one of the measured temperatures from the other.

108. A method as in claim 107 where said temperatures are obtained by a bow tie detecting element.

109. A method as in claim 107 where the near temperature is obtained by means for measuring the air immediately outside the aircraft.

110. A method of detecting the presence of a weather related event with respect to an aircraft comprising the steps of: obtaining a temperature measurement which is a function of the temperature of the air in front of the aircraft; obtaining a value of the lapse rate where the lapse rate is the slope of the temperature versus altitude curve of the air; obtaining a value of look distance where said look distance is the distance from the aircraft that the measurement would occur if the aircraft were flying horizontal; obtaining the pitch angle of the aircraft; calculating an error temperature from the lapse rate, the look distance, and the pitch angle; correcting the measured temperature with the error temperature to obtain a corrected temperature which corresponds to what the temperature would be at the look distance from the aircraft if the aircraft were horizontal ; and proc e s s ing the correc ted temperature to detect the presence of the weather related event .

111. A method as in claim 110 where said weather related event is low level wind shear .

112. A method as in claim 110 where said weather related event is clear air turbulence .

113 . A method as in claim 110 where s aid temperature measurement is obtained as Δ T where Δ T is the difference between a first temperature near the aircraft and a second temperature remote from the aircraft .

114. A method as in claim 110 where the lapse rate is a predetermined value thereof .

115 . A method as in claim 114 where s aid predetermined value is 6.0 to 6.5 ^* Celsius/km.

116 . A method as in claim 114 where s aid predetermined value is utilized in response to the aircraft ascending.

117 . A method as in claim 110 where the lapse rate is obtained from measured values of (a) a first temperature outside the aircraft at a first altitude thereof and (b ) a second temperature outside the aircraft at a second altitude thereof .

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118. A method as in claim 117 where the obtained lapse rate is utilized in response to the aircraft descending.

119. A method as in claim 110 where the error temperature is calculated as the product of the lapse rate, the look distance, and the tangent of the pitch angle.

120. A method using a thermal reactive system for determining low level windshear comprising the steps of: determining whether a low level wind shear (LLWS) index is exceeded to detect the presence of the LLWS; and adjusting the value of the index to compensate for lapse rate during take off or landing of the aircraft.

121. A method as in claim 120 where said index value is normally 0.15 and is adjusted to 0.13 during landing and to 0.17 during take off.

122. A method of obtaining look distance where look distance is the distance a temperature measurement occurs in front of an infrared spectrometer comprising the steps of: obtaining a first temperature measurement in front of the spectrometer and a second temperature at a predetermined vertical angle with respect to the first temperature measurement; obtaining a value of the lapse rate where the lapse rate is the slope of the temperature versus altitude curve of the air; and obtaining the look distance by dividing the difference between the first and second temperatures by the value of the lapse rate.

123. A method of detecting clear air turbulence (CAT) with respect to an aircraft comprising the steps of : obtaining a temperature difference signal representative of the difference between carbon dioxide frequencies of 13.5 and 15 microns ; processing said temperature difference signal to determine whether said clear air turbulence is present.

124 . A method as in claim 123 where said proces sing step includes obtaining the standard deviation of the temperature dif ference signal and comparing it with a reference standard deviation to determine whether the clear air turbulence is present.

125. A method of detecting clear air turbulence (CAT) with respect to an aircraft comprising the steps of : obtaining a temperature difference signal representative of the di f ference between H₂0 frequencies in the 17 to 21 microns region; processing said temperature difference signal to determine whether said clear air turbulence is present.

126 . A method as in claim 125 where s aid processing step includes obtaining the standard deviation of the temperature difference signal and comparing it with a reference standard deviation to determine whether the clear air turbulence is present.

127. A method of detecting particles in the air which are potentially volcanic ash comprising the steps of : providing at least one detecting means to obtain a temperature difference signal indicative of the difference between the (a) the air temperature at a distance remote the aircraft and (b) the air temperature at a distance near the aircraft; processing said signal to determine whether said temperature difference signal is substantially less than an expected temperature difference signal and thus detect the potential presence of said particles .

128. A method as in claim 127 where said particles are detected at altitudes lower than a predetermined altitude .

129. A method as in claim 128 where said predetermined altitude is about 35,000 feet.

130. A method of clear air turbulence detection comprising: using water based wavelengths to detect said clear air turbulence; and switching to carbon dioxide based wavelengths in the presence of clouds to detect said clear air turbulence.