Patent Publication Number: US-6906641-B2

Title: Apparatus, method and computer program product for helicopter enhanced ground proximity warning system

Description:
CROSS REFERENCES TO RELATED APPLICATIONS 
   This application claims priority from U.S. Provisional application Ser. No. 60/207,515 titled “Terrain Floor Delta Height for Helicopter EGPWS Based on Ground Speed,” filed May 26, 2000; and from U.S. Provisional application Ser. No. 60/207,740 titled “Terrain Awareness Display Coloring for Helicopter EGPWS Based on Ground Speed,” filed May 26, 2000; and from U.S. Provisional application Ser. No. 60/207,998 titled “Look Ahead Distance for Helicopter EGPWS Based on Stopping Distance,” filed May 26, 2000; the entire specifications of which are herein incorporated by reference. 
   This application is also related to U.S. Provisional application Ser. No. 60/232,967, titled: “Tail Strike Algorithm for Helicopters” and filed Sep. 14, 2000 and to application Ser. No. 09/865,365 filed the same day herewith and titled: “Method, Apparatus and Computer Program Product for Helicopter Tail Strike Warning”. 

   BACKGROUND OF THE INVENTION 
   The present invention provides a ground proximity warning system and method for rotary wing aircraft such as helicopters, gyrocopters, and tilt rotors and more particularly to logic and displays useful in a helicopter enhanced ground proximity warning system, or EGPWS. 
   Ground proximity warning systems, or GPWS, provide aural and visual warnings of conditions when the aircraft is in potentially hazardous proximity to terrain, and/or in a flight condition apparently inappropriate given the aircraft&#39;s position relative to terrain. Earlier generation ground proximity warning systems sensed dangerous approach to terrain by using a radar altimeter to sense height above the ground. The rate at which height above ground changes, is compared with a predefined envelope(s) to determine if a dangerous condition exists. Classic GPWS systems also contain additional alert functions called ‘modes’ that alert to other potentially hazardous conditions based on flight regime. Examples of GPWS devices are contained in U.S. Pat. Nos.: 3,715,718; 3,936,796; 3,958,218; 3,944,968; 3,947,808; 3,947,810; 3,934,221; 3,958,219; 3,925,751; 3,934,222; 4,060,793; 4,030,065; 4,215,334; and 4,319,218. 
   Later generation GPWS devices, called EGPWS devices or terrain awareness systems (TAWS), include a stored terrain database that compares the position of the aircraft in three dimensional space with the stored terrain information to identify potential conflicts. EGPWS devices may also include all the functionality and modes of the classic GPWS devices. Examples of EGPWS-type devices include U.S. Pat. Nos.: 4,646,244; 5,839,080; 5,414,631; 5,448,563; 5,661,486; 4,224,669; 6,088,634; 6,092,009; 6,122,570 and 6,138,060. 
   In certain EGPWS designs, the position of the terrain relative to the aircraft may be shown on a display in the cockpit. In some displays, the terrain is color-coded according to the degree of hazard. For example, green colored terrain usually depicts nonhazardous terrain below the aircraft. Yellow colored terrain usually depicts terrain that is in proximity to the aircraft and/or which may cause the ground proximity system to generate a precautionary alert. Red colored terrain usually depicts terrain at or above the aircraft altitude or for which the ground proximity warning system will issue a warning from which evasive action must be taken. U.S. Pat. Nos.: 5,839,080 and 6,138,060 describe some terrain cockpit displays. U.S. Pat. No. 5,936,522 describes a terrain display having vertical and plan views. 
   The above referenced systems have been primarily developed for fixed wing aircraft. Rotary wing aircraft and aircraft capable of hover present unique challenges for ground proximity alerting due to the different flight profiles flown and the unique capabilities of rotary wing aircraft. For example, unlike fixed wing aircraft, rotary wing aircraft can cease all forward motion while still remaining airborne. Rotary wing aircraft can also descend straight down from a hover to land on all sorts of terrain, and need not make a gradual descent and approach to land as in the case of fixed wing aircraft. 
   U.S. Pat. Nos. 5,781,126 titled “Ground Proximity Warning System and Methods for Rotary Wing Aircraft;” 5,666,110 titled “Helicopter Enhanced Descent After Take-off Warning for GPWS;” and 6,043,759 titled “Air Ground Logic System and Method for Rotary Wing Aircraft;” and co-pending application Ser. No. 08/844,116 titled: “Systems and Methods for Generating Altitude Callouts for Rotary Wing Aircraft,” each address various issues associated with applying ground proximity warning technology to rotary wing aircraft and are each incorporated herein by reference. These patents are applicable to both conventional and enhanced ground proximity warning designs for use in helicopters, however, these patents address the particularities of modifying various of the “modes” for use in helicopters. Specifically, U.S. Pat. No. 5,781,126 includes a barometric altitude rate detector including a controller for adjusting this rate detector to account for downwash of the rotary wing. U.S. Pat. No. 5,666,110 discloses a descent after take-off protection envelope. U.S. Pat. No. 6,043,759 discloses a logic method and device for determining when the helicopter is in the airborne or ground state which assists with preventing nuisance alarms during helicopter autorotations. Ser. No. 08/844,116 discloses a device and method for generating altitude call outs during helicopter landing operations. 
   None of the above mentioned patents account for modifying the terrain look ahead logic or the associated terrain display of an EGPWS type device to account for the unique flying performance of helicopters and other rotary wing craft. 
   SUMMARY OF THE INVENTION 
   The present invention recognizes the problems in enhanced ground proximity alerting for rotary wing aircraft such as, for example, helicopters, gyrocopters and tilt rotors when in the rotor mode, hereinafter generically and interchangeably referred to as “helicopter(s)” or “rotary wing aircraft”. In particular, the present invention recognizes that certain categories of aircraft such as rotary wing aircraft and airships are capable of executing an avoidance maneuver by coming to a stop in a hover. The present invention further recognizes that the terrain threats as depicted on the display should preferably also be modified to reflect the unique, yet normal operating capabilities of such aircraft. In addition, the present invention recognizes the unique landing characteristics of rotary wing aircraft. 
   According to one aspect of the present invention, a helicopter EGPWS uses a look ahead distance to define terrain that is a threat to the aircraft. If terrain is located within the boundaries of the warning envelope, an alert is given to the pilot. In a preferred embodiment of the present invention, this look ahead distance is based upon the nominal distance required to halt forward momentum and enter a hover. 
   According to another aspect of the present invention, a helicopter EGPWS permits the helicopter to land at any location, including off-airport locations, without incurring nuisance alarms. In a preferred embodiment of the invention, the helicopter EGPWS uses groundspeed in conjunction with helicopter vertical speed to define a terrain floor below which a proximity warning will be given. In this manner, the helicopter executing a safe landing, for example, on a hillside will not incur an unwanted terrain proximity alert. 
   According to yet another aspect of the present invention, the present invention recognizes that a display color coded to alert pilots of fixed wing aircraft to potential hazards may result in displaying non-threatening terrain as a hazard to pilots of rotary wing aircraft. The EGPWS and EGPWS display of the present invention color codes terrain based upon flight regimes associated with modes of operation typical for rotary wing aircraft. 
   Further details and operation of the present invention will be described below with reference to the drawings. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIGS. 1A-1B  are a top level block diagram of an EGPWS computer for use on helicopters according to an embodiment of the present invention; 
       FIG. 2  is a functional block diagram of an EGPWS computer according to an embodiment of the present invention; 
       FIG. 3  diagrams six basic warning modes for a helicopter ground proximity warning system according to an embodiment of the present invention; 
       FIG. 4  is a side view of terrain caution and warning envelopes according to an embodiment of the present invention; 
       FIG. 5  is a perspective view of terrain caution and warning envelopes according to an embodiment of the present invention; 
       FIG. 6  illustrates a look ahead distance for use in determining terrain caution and warning envelopes for aircraft capable of hover according to an embodiment of the present invention; 
       FIG. 7  graphs a function useful for obtaining terrain floor ΔH for on-airport landings according to an embodiment of the present invention; 
       FIG. 8  graphs a function useful for obtaining terrain floor ΔH for helicopters according to an embodiment of the present invention; 
       FIG. 9  is a graphical illustration of a cut-off angle correction boundary for a level flight condition according to an embodiment of the present invention; 
       FIG. 10  is a graphical illustration of a cut-off angle correction boundary for a condition where the flight path angle of the aircraft is greater than a predefined reference or datum according to an embodiment of the present invention; 
       FIG. 11  is similar to  FIG. 9  but for a condition where the flight path angle of the aircraft is less than a predefined reference plane or datum and also illustrates a BETA sink rate enhancement boundary according to an embodiment of the present invention; 
       FIG. 12  is a graphical illustration of a cut-off altitude boundary according to an embodiment of the present invention; 
       FIG. 13  illustrates a terrain caution and warning envelopes for a condition where the aircraft is descending according to an embodiment of the present invention; 
       FIG. 14  illustrates a terrain caution and warning envelopes for a condition where the aircraft is climbing according to an embodiment of the present invention; 
       FIG. 15  illustrates a look up caution and warning envelopes for a condition for detecting precipitous terrain ahead of the aircraft according to an embodiment of the present invention; 
       FIG. 16  is a functional block diagram for a system useful for asserting a signal representative of altitude loss due to pilot reaction time, ALPT, as well as altitude loss due to a pull up maneuver, ALPU, according to an embodiment of the present invention; 
       FIG. 17  is a functional block diagram of a terrain display system according to an embodiment of the present invention; 
       FIG. 18  is a block diagram of a display system according to an embodiment of the present invention. 
       FIG. 19  illustrates depiction of background terrain on a display according to an embodiment of the present invention; 
       FIG. 20  illustrates a background terrain display color and dot pattern densities according to an embodiment of the present invention; and 
       FIG. 21  illustrates an alternative background terrain display color and dot pattern densities according to an embodiment of the present invention. 
   

   DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
   System Overview 
   U.S. Pat. No. 5,839,080, incorporated herein by reference, describes an EGPWS device manufactured by Honeywell International Inc., and suitable for use with the present invention. Additional EGPWS features applicable to rotary wing aircraft are described in U.S. Pat. Nos. 6,138,060; 6,122,570; 6,092,009; 6,088,634; as well as in copending application Ser. Nos.: 09/099,822; 09/074,953; 09/103,349; 09/255,670 and 09/496,297, each of which is incorporated by reference.  FIGS. 1A and 1B  provide a top level description of such a system in block diagram form. The terrain awareness system  20  utilizes navigation information from a global positioning system  22  and/or a flight management system (FMS) and/or inertial navigation system. Navigation information may also be received from other aviation navigation systems such as, for example: DME/DME, VOR/DME, RNAV, and LORAN systems. System  20  further includes a terrain/obstacle database  24 , and/or an airport database  26  (collectively “terrain”) and a corrected barometric altitude signal which may be obtained from an air data computer or barometric altimeter  28 . In a preferred embodiment of the invention, altimetry information can be obtained in accordance with the techniques described in co-pending application Ser. No. 09/255,670 titled “Method and Apparatus for Determining Altitude” and/or co-pending application No. titled, “Device, Method and Computer Program Product for Altimetry System” filed Feb. 2, 2001. 
   The latitude and longitude of the current aircraft position are applied to an airport and terrain search algorithm, indicated by a block  29  which includes location search logic for determining the terrain data, as well as the airport data surrounding the aircraft. Example search logic is described in U.S. Pat. Nos. 4,675,823 and 4,914,436 assigned to the assignee of the present invention and incorporated herein by reference as well as in U.S. Pat. No. 5,839,080. The navigational position data, along with the terrain and airport data are supplied to a threat assessment function  30  which provides both terrain advisory and terrain warning signals based upon the position and flight path vector of the aircraft. Function  30  may provide both aural and/or visual warnings when a hazardous condition is believed to exist. Aural warnings may be provided by voice generator  32  and speaker  34 . Visual warnings may be provided by a moving map or display  36 . Display  36  may comprise any cockpit display, such as, for example, a weather radar display, a TCAS display, an Electronic Flight Instrument System (EFIS) display or a Honeywell UDI display. The terrain and obstacles depicted on display  36  may be colored according to the degree of threat in a manner to be described in greater detail below. 
     FIG. 2  contains a more detailed functional block diagram of EGPWS computer  20  of the present invention. The EGPWS computer  20  as shown in  FIGS. 1A-1B  and in  FIG. 2 , may be implemented as executable code, an analog or digital electronic circuit, on a PCMCIA card, as programmable logic and/or as a general purpose processor. In a preferred embodiment of the invention, warning computer  20  is implemented as a line replaceable unit (LRU) containing a microprocessor. Database  24  is included on a PCMCIA card which may be loaded into the LRU and also used to provide periodic system upgrades. 
   In the embodiment of  FIG. 2 , local terrain processing function  29  receives as input the aircraft position and heading data and retrieves from database(s)  24  and  26 , the terrain, obstacle and/or runway data in the vicinity of the aircraft. As described in U.S. Pat. No. 5,839,080, herein incorporated by reference for all purposes, terrain processing function  29  extracts and formats the local topographical data and terrain features to create a set of elevation matrix overlays that are positioned with respect to the current aircraft location. Each matrix element includes as data, the highest terrain altitude with respect to mean sea level contained within that element&#39;s area. Terrain processing function  29  optionally retrieves any obstacle data associated with the matrix elements as well as retrieving data for the runway nearest the current location of the aircraft. Runway data includes the runway endpoint locations and may be processed to include nearest runway center position, nearest runway threshold position, and nearest runway altitude. 
   The threat detection and terrain display processing function  30  receives as input the terrain data as processed by function  29  as well as the current aircraft position, altitude, attitude, speed and track information. When a potential hazard to the aircraft exists, function  30  controls the output of an alert which may include an aural warning, the illumination of a lamp, and/or depiction of the threat on display  36 . Results of the threat detection process are combined with background terrain data/obstacle matrix data and data for the nearest runway and formatted into a matching set of display matrix overlays for display on display  36 . A display output processor  50  receives the set of display matrix overlays from function  30  as well as the aircraft position and heading information to drive display  36 . A display control logic  52  controls the display range and the activation of display of terrain data on the chosen display. In the embodiment of  FIG. 2 , display control logic  52  controls an external switch  54  that may be used to switch a weather radar display from display of weather to display of terrain. 
     FIG. 3  diagrams the six basic warning modes for the helicopter ground proximity warning system of the present invention. The various modes provide aural and visual alerts and warnings including warnings for: unsafe proximity to terrain, deviation below ILS glide slope, excessive bank angle, onset of severe wind shear, altitude awareness. Mode one, for example, provides pilots with alert/warnings for high descent rates into terrain. In this mode, a warning device compares the altitude above ground of the aircraft with the descent rate, preferably barometric descent rate, and issues a warning if the descent rate is excessive for the altitude at which the aircraft is flying. A more complete description of an exemplary warning device for indicating excessive descent rate can be found in U.S. Pat. No. 4,551,723, the complete disclosure of which has previously been incorporated herein by reference. Mode two provides warnings for excessive closure rates to terrain with respect to altitude (AGL), phase of flight and speed. Mode three provides warnings for significant altitude loss after takeoff or low altitude go around as described in U.S. Pat. No. 5,666,110. A complete description of the system can be found in U.S. Pat. No. 4,818,992, the complete disclosure of which has previously been incorporated herein by reference. 
   Mode four provides alerts and warnings for insufficient terrain clearance with respect to phase of flight and speed. Mode five provides glide slope alerts when the airplane is below 1,000 ft. AGL with the gear down and the glide slope deviation exceeds a threshold number of dots below the ILS glide slope. Mode six provides callouts for descent through predefined altitudes (AGL). In particular, mode six is utilized during autorotation when the aircraft has lost all or partial engine power. Each of the various warning modes provides at least an aural alert for a particular hazard as shown. 
   According to another embodiment of the invention, Mode  6  includes a capability for alerting the pilot of a helicopter of a possible tail strike condition. Details of the Mode  6  tail strike alert are contained in copending U.S. application Ser. No. 09/865,365 filed the same day herewith and incorporated herein by reference. 
   Terrain Caution and Warning Envelopes 
   In addition to the warning modes depicted in  FIG. 3 , and as described above in connection with  FIGS. 1A-1B  and  FIG. 2 , the EGPWS device of the present invention utilizes a plurality of caution and alert envelopes to warn of potential terrain hazards. If the aircraft penetrates the caution envelope boundary, the aural message “ Caution Terrain, Caution Terrain ”  is generated, and alert discretes are provided for activation of visual annunciators. Terrain located within the caution envelope is shown in solid yellow color on display  36 . If terrain proximate the aircraft penetrates the warning envelope boundary, the aural message “Warning Terrain” is generated, and alert discretes are provided for activation of visual annunciators. Terrain located within the warning envelope is shown in solid red color on display  36 . 
   The caution and warning envelopes are obtained using a terrain floor and a “look ahead” distance to define a volume which is calculated as a function of groundspeed and flight path angle.  FIG. 4  shows a simplified side view of caution and warning envelopes  100  and  101  according to an embodiment of the present invention.  FIG. 5  shows a perspective view of the caution and warning envelopes. 
   According to a preferred embodiment of the invention, terrain and caution alerting can be provided for both a “normal” mode and a “low level” mode of operation. The low level mode is applicable to low altitude flight in daylight VFR conditions. Use of the low level caution and warning logic is selectable by the pilot and reduces nuisance warnings by accounting for the unique low altitude operations of rotary wing aircraft. The normal mode comprises the default condition and is used for cruise, night time and instrument flight rules operations. The pilot may select between the caution and warning envelopes of the normal mode and the alert envelopes and/or display of the low level mode via a cockpit switch or display touch screen. 
   The look ahead distance of the caution and warning terrain protection envelopes is taken in a direction along the groundtrack of the aircraft. To reduce nuisance warnings, the look ahead distance may have a maximum value. Otherwise, potentially threatening terrain along the current flight path of the aircraft relatively far from the current position could produce nuisance warnings. Two different look ahead distances (LAD) are utilized. The first LAD is used for a terrain caution signal. A second LAD is used for terrain warning signals which require immediate evasive action. 
   In aircraft not capable of hover, and in the normal mode, the LAD for a terrain advisory condition is considered first in determining the LAD because it is assumed that the pilot could make a turn at any time at a turning radius R. For a fixed wing aircraft as fully described in U.S. Pat. No. 5,839,080, the total look ahead time is equal to the sum of the look ahead time T 1  of a single turn of radius R; the look ahead time T 2  for terrain clearance at the top of the turn plus a predetermined reaction time T 3 . In a helicopter, or other aircraft capable of hover using the low level mode, however, the look ahead distance can instead be based on the distance required to bring the aircraft to a stop or hover plus the distance covered during a nominal reaction time. In the case of a helicopter, the distance required to transition from cruise to hover using, for example, a 10° pitch up at constant altitude may be used.  FIG. 6  diagrams the look ahead distance for the helicopter according to the present invention. 
   The LAD can be expressed as:
 
LAD=Transition Distance to Hover+Reaction Time Distance  Eq. (1).
 
   Assuming the aircraft comes to a stop, the total distance covered when transitioning to hover is governed by the standard equation: 
             s   =       1   2     ⁢     at   2               Eq   .           ⁢     (   2   )               
 
   where: 
   s=distance to stop 
   a=deceleration 
   t=time. 
   The average velocity during the deceleration interval t is: 
                 V   _     =       G   ⁢           ⁢   S     2       ,           Eq   .           ⁢     (   3   )               
 
   where GS =groundspeed. 
   The distance covered can then be written as: 
             s   =         G   ⁢           ⁢   S     2     ⁢   t             Eq   .           ⁢     (   4   )               
 
   Substituting Eq. (4) back into Eq. (2); and solving for time, t, yields: 
             t   =       G   ⁢           ⁢   S     a             Eq   .           ⁢     (   5   )               
 
   Eq. 5 can then be used to develop an expression for stopping distance as given in Eq. (6) below: 
             s   =         (   GS   )     2       2   ⁢   a               Eq   .           ⁢     (   6   )               
 
   For a nominal pitch up of 10° the following equation can be used to solve for a:
 
 ma=mg  (tan 10°)  Eq. (7)
 
   or, 
             a   =       g   ⁢           ⁢   tan   ⁢           ⁢   10   ⁢   °     =         (     68682   ⁢           ⁢     nm     hr   2         )     ⁢     (   0.18   )       =     12363   ⁢           ⁢     nm     hr   2                     Eq   .           ⁢     (   8   )               
 
   The above derivations have the advantage of making the calculation of LAD independent of the aircraft mass and hence independent of aircraft type. 
   For a predetermined reaction time T 3 , for example, 10 seconds, the look ahead distance LAD in nautical miles for a terrain advisory signal can be determined simply by multiplying the ground speed of the aircraft (V) by the reaction time T 3  and adding this value to the stopping distance as shown in Eq. (9). 
             LAD   =       (         (     G   ⁢           ⁢   S     )     2       2   ⁢   a       )     +     G   ⁢           ⁢     S   ⁡     (     T   3     )                   Eq   .           ⁢     (   9   )               
 
   Table I lists the resulting LAD for a 10° constant altitude transition to hover and a reaction time of 5 seconds: 
   
     
       
         
             
           
             
               TABLE I 
             
           
          
             
                 
             
             
               LAD For 10° Pitch, 
             
             
               Constant Altitude Hover And 5 Second Reaction Time 
             
          
         
         
             
             
             
          
             
                 
               GROUNDSPEED (kts) 
               LAD (nm) 
             
             
                 
                 
             
             
                 
                80 
               0.55 
             
             
                 
               100 
               0.78 
             
             
                 
               120 
               1.06 
             
             
                 
                 
             
          
         
       
     
   
   The LAD may optionally be additionally bounded by an upper limit and a lower limit. The lower limit may be a configurable amount, for example; either 0.35, 1 or 1 ½ nautical miles at relatively low speeds e.g. speeds less than 40 knots, for example, and 4 nautical miles at higher speeds, for example, greater than 250 knots. The LAD may also be limited to a fixed amount regardless of the speed when the distance to the nearest runway is less than a predetermined amount, for example, 2 nautical miles, except when the aircraft altitude is greater than 500 feet, relative to the runway. 
   In a preferred embodiment of the invention, the LAD for a terrain warning signal is taken as ½ the terrain caution signal LAD. Optionally, and as described more fully in U.S. Pat. No. 5,839,080, the terrain warning LAD may be given by equation (10) below.
 
LAD=k 1 *LAD (terrain caution), k 2 *LAD (terrain warning), K 3  LAD (terrain look-up advisory)  Eq. (10)
 
   where k 1 =1.5, except when the LAD is limited at its lower limit, in which case k 1 =1, k 3 =1 and where k 2 =0.5, k 3 =2. 
   As shown in  FIGS. 4 and 5 , the caution and advisory envelopes are additionally defined by a terrain floor boundary. In a fixed wing aircraft, the terrain floor relates to a distance AH below the aircraft and is proportional to the distance to the closest runway to prevent nuisance warnings when the aircraft is taking off and landing, while providing adequate protection in other modes of operation.  FIG. 7  illustrates the terrain floor AH used for airport landings. The terrain floor boundary below the aircraft is essentially based upon providing 100 feet clearance per nautical mile from the runway as identified by runway centerpoint  72  and endpoint  80 , limited to 150 feet. In a preferred embodiment of the invention, distance D from runway centerpoint  72  is equal to 12 nm. 
   However, in aircraft capable of landing safely off-airport, such as a helicopter, nuisance alarms will occur when the helicopter lands on, for example, a hillside or other safe but off-airport/helipad location. For this reason, the present invention utilizes the terrain floor ΔH of FIG.  8 . In  FIG. 8 , the horizontal axis represents the groundspeed while the vertical axis represents the ΔH terrain floor boundary beneath the aircraft. The terrain floor ΔH boundary beneath the aircraft is limited such that the segment commencing at point  78  beings at 0 feet and the segment  82  never goes above a predetermined maximum, for example, 150 feet. The groundspeed corresponding to point  78  is preferably the airspeed corresponding to the landing or touchdown speed in zero wind conditions. Therefore, as illustrated in  FIG. 8 , so long as the pilot continuously slows the aircraft while descending, no terrain alert will be given by the present invention. Such conditions are indicative of an approach to land and are not likely to be indicative of a controlled flight into terrain accident. 
   For helicopters with retractable landing gear, the terrain floor delta height function of  FIG. 8  can be additionally coupled to logic that detects when the gear is deployed. 
   If the gear is not deployed, the curve of  FIG. 8  can be disabled and the ΔH curve of  FIG. 7  utilized. The Mode  4  “Too Low Gear” warning will also sound. 
   In order to avoid spurious warnings when the aircraft over flies a ridge at relatively low altitudes, the terrain advisory and warning boundaries may additionally include cut-off boundaries, for example, as illustrated in  FIGS. 9 ,  10  and  11 . Without the cut-off boundaries, warnings would be given, although the terrain is virtually below the aircraft and no terrain is visible ahead. In  FIG. 9 , the cut-off boundary  126  begins at a predetermined cut-off offset  128  below the aircraft and extends in a direction in front of the aircraft at a predetermined envelope cut-off angle  130 . The envelope cut-off angle  130  is equal to the flight path angle y plus a configurable predetermined cut-off angle, described and illustrated as 6°. For level flight as shown in  FIG. 9 , the cut-off boundary  126  extends from the cut-off offset  128  in the direction of the envelope cut-off angle  130  toward the front of the aircraft to a point  132  where it intersects a terrain caution boundary or terrain warning boundary, identified with the reference numeral  134 . For level flight, as shown in  FIG. 9 , the flight path angle y is zero. Thus, the cut-off boundary  126  illustrated in  FIG. 9  will extend from the cut-off offset  128  along an angle equal to the cut-off angle. 
   The cut-off boundary  126  extends from the cut-off offset  129  to the point  132  where it intersects the terrain caution boundary  134 . The warning boundary is then selected to be the highest of the terrain caution boundary  134  and the envelope cut-off boundary  126 . Thus, for the example illustrated in  FIG. 9 , the terrain caution boundary would include the cut-off boundary  126  up to the point  132 , where the envelope cut-off boundary  124  intersects the warning envelope  126 . From the point  132  forward, the normal terrain caution boundary  134 , corresponding, for example, to a THETA 1  slope is utilized. Thus, if either a terrain caution or terrain warning boundary is below the cut-off boundary  126 , the cut-off boundary  126  becomes the new boundary for the caution or warning signal. 
   The cut-off boundary may additionally include a cut-off altitude. The cut-off altitude is an altitude relative to the nearest runway elevation; set at, for example, 500 feet. Altitudes below the cut-off altitude are ignored by the terrain advisory and terrain warning computations. An advantage to using a cut-off altitude is that nuisance warnings on final approach, due to altitude errors, terrain database resolution and accuracy errors are minimized. 
   However, the use of a cut-off altitude during certain conditions, such as an approach to an airport on a bluff (i.e. Paine Field) at a relatively low altitude or even at an altitude below the airport altitude, may compromise system performance. More particularly, during such conditions, the use of a cut-off altitude may prevent a terrain warning from being generated because the threatening terrain may be below the cut-off altitude. In order to account for such situations and for the ability of helicopter type aircraft to land safely off the airport, the system selects the lower of two cut-off altitudes; a nearest runway cut-off altitude (NRCA) and a cut-off altitude relative to aircraft (CARA). The NRCA is a fixed cut-off altitude, relative to the nearest runway. The CARA is an altitude below the instantaneous aircraft altitude (ACA) by an amount essentially equivalent the ΔH terrain floor boundary of  FIG. 7  or  8 , whichever is smaller. 
   Equations (11) and (12) below set forth the NRCA and CARA. As mentioned above, the absolute cut-off altitude (ACOA) is the lower of the NRCA and CARA as set forth in equation ( 13 ).
 
NRCA=COH+RE,  Eq. (11)
 
   where COH relates to the cut-off height and is a fixed configurable value, initially set between 400 feet and 500 feet; and RE equals the runway elevation.
 
CARA=ACA-ΔH-DHO,  Eq. (12)
 
   where ACA is the instantaneous aircraft altitude; ΔH is the smaller of the terrain floor of  FIG. 7  or  8  and DHO is a configurable bias, set to, for example, 50 feet.
 
ACOA=lower of CARA, NRCA,  Eq. (13)
 
   For landings in the vicinity of an airport runway contained in the database, however, a point, DH 1 , exists for which the ACOA is forced to be equal to NRCA independent of the aircraft altitude. The point DH 1  is related to COH, AΔand DHO such that on a nominal three (3) degree approach slope, CARA is equal to NRCA when the aircraft is at a distance equal to a distance DH 1  from the airport, as illustrated in Table II below: 
   
     
       
         
             
           
             
               TABLE II 
             
           
          
             
                 
             
             
               Relationship Between COH, DH1 and Runway Distance 
             
          
         
         
             
             
             
          
             
               COH 
               DH1 
               DISTANCE TO RUNWAY 
             
             
               (feet) 
               (feet) 
               (n mile) 
             
             
                 
             
          
         
         
             
             
             
          
             
               300 
                50 
               1 
             
             
               400 
               100 
               1.5 
             
             
               500 
               150 
               2 
             
             
                 
             
          
         
       
     
   
   The point DH 1  forces the cut-off altitude (COH) above the runway whenever the aircraft is close to the runway to ensure robustness against nuisance warnings caused by altitude errors and terrain database resolution margins to disable the terrain caution and warning indications when the aircraft is within the airport perimeter. There are trade-offs between nuisance warnings and legitimate warnings. In particular, the lower COH, the closer the caution and warning indications are given, making the system more nuisance prone. As indicated above, for a COH of 400, terrain caution and warning indications are effectively disabled when the aircraft is closer than 1.5 nm from an airport runway. 
     FIG. 12  illustrates the operation of the alternative cut-off altitude boundaries. In particular,  FIG. 12  illustrates a condition when the COH is set to 300 feet with DH 1  equal to 50 feet. The cut-off altitude for an area from the runway, for example, greater than 4 nautical miles, is 300 feet, as indicated by the segment  278  when the glide slope angle is less than a predetermined angle, for example, 3°. As the aircraft gets closer to the runway, the COH is lowered, as illustrated by the segment  288 , until the aircraft is within one (1) nautical mile of the runway, at which point the COH is forced to be 300 feet, which effectively disables any terrain caution and warning indications when the aircraft is closer than one (1) nautical mile from the runway, as represented by the segment  282 . During a condition when the aircraft is on, for example, a 3° glide slope angle to the database runway, the ACOA is forced to be the NRCA. The NRCA is illustrated by the segment  282  of FIG.  12 . 
   The resulting terrain threat envelopes are illustrated in  FIGS. 13-15 . The terrain caution and warning envelopes may be thought of as including two parts: a look-ahead/look-down boundary for detecting terrain ahead or below the aircraft as shown in FIGS.  13  and  14 ; and a look-up boundary for detecting precipitous high terrain ahead of the aircraft which may be difficult to clear as shown in FIG.  15 . 
   In  FIG. 13 , caution and warning envelopes are illustrated for a condition when the aircraft is descending (i.e. γ&lt;0). During such a condition, the first segment of the terrain caution envelope, identified with the reference numeral  300 , corresponds to the ΔH terrain floor boundary. To determine the bottom segment  302  of the terrain caution envelope, the flight path angle y is compared with a configurable datum, THETA 1 , for example 0°. During descent conditions, the flight path angle y will thus be less than THETA 1 . Thus, the look-ahead/look-down terrain advisory boundary segment will extend from the ΔH terrain floor boundary segment  300  along the angle THETA 1  to the look-ahead distance for a terrain advisory (LAD). The final segment  304  extends vertically upward from the segment  302  along the LAD. 
   The terrain caution boundary may also be modified by a BETA sink rate enhancement. In this embodiment, the BETA sink rate enhancement ensures that an advisory indication always precedes a warning indication when the aircraft descends into or on top of terrain. The BETA sink rate enhancement is determined as a function of the flight path angle γ and two (2) configurable constants KBETA and GBIAS. The BETA sink rate enhancement BETA 1  for a look-ahead/look-down terrain advisory is provided in equation (14) below for a condition when the aircraft is descending. The angle BETA 1  is given by:
 
BETA 1 =KBETA*(γ-GBIAS),  Eq. (14)
 
   where GBIAS is a configurable constant, selected for example, to be zero (0) and KBETA is also a configurable constant selected, for example, to be 0.5. In the embodiment of  FIG. 13 , the BETA sink rate enhancement BETA 1  for the look-ahead/look-down terrain advisory boundary provides an advisory warning at a distance of ½ LAD. The BETA sink rate enhancement BETA 1  results in a segment  306  which extends from the ΔH terrain floor segments  300  at an angle equal to γ/2 up to ½ of the LAD. Beyond ½ LAD, a LAD, a segment  308  extends at the angle THETA 1  to a distance equal to the LAD. A vertical segment  310  extends along the LAD to connect the segments  308  to the segment  304 . 
   In  FIG. 13 , the cut-off boundary is identified with the reference numeral  312 . The cut-off boundary  312  extends from a vertical distance  314  below the aircraft along a cut-off angle up to the point of intersection  316  with the terrain advisory boundary. For distances less than the intersection  316 , the cut-off boundary  312  forms the terrain caution envelope boundary. For distances beyond the point of intersection  316 , the boundaries  306  and  308  form the terrain caution boundaries. 
   The terrain warning boundary includes the segment  300  extending from the aircraft along the ΔH terrain floor. A bottom segment  318  connects to the segment  300  and extends along a BETA sink rate enhancement angle BETA 2 , where angle BETA 2  is given by the equation:
 
BETA 2 =KBETA 2 *(γ-GBIAS),  Eq. (15)
 
   where GBIAS is a configurable constant selected, for example, to be 0 and KBETA 2  is also a configurable constant selected, for example, to be 0.25. For such values of the constants KBETA 2  and GBIAS, the BETA enhancement angle KBETA 2  will be ¼*γ. Thus, the segment  318  extends from the segment  300  at an angle equal to ¼*γ up to ½ the LAD. A vertical segment  320  extends along a distance equal to ½*LAD from the segment  318  to define the terrain warning boundary. 
   The terrain warning boundaries are also limited by the cut-off boundary  312 . Thus, the cut-off boundary  312  forms the terrain warning boundary up to a point  322 , where the cut-off boundary  312  intersects the lower terrain warning boundary  318 . At distances beyond the point of intersection  322 , the segment  318  forms the lower terrain warning boundary up to a distance equal to ½ of the LAD. 
   The terrain advisory and terrain warning boundaries for a condition when the aircraft is climbing (i.e. γ&gt;0) is illustrated in FIG.  14 . During such a condition, the BETA sink rate enhancement angles BETA 1  and BETA 2  are set to a configurable constant, for example, zero (0). The terrain advisory boundary during the climbing condition is formed by extending a vertical segment  324  from the aircraft for a distance below the aircraft equal to the ΔH terrain floor of  FIG. 7  or  8 , whichever is smaller. A segment  326  is extended from the segment  324  to the LAD at an angle equal to the flight path angle γ. At a point  328  where the segment  326  intersects a position equal to ½ of the LAD, a vertical segment  330  is extended up from the segment  326 , forming a first vertical boundary for the terrain advisory condition. The line segment  326  from the point  328  to the LAD forms the lower terrain advisory boundary while a line segment  332  extending vertically upward from the line segment  326  along the LAD forms a second vertical boundary. 
   For the exemplary condition illustrated, a cut-off boundary  334  does not intersect the terrain caution boundaries. Thus, the terrain caution boundaries for the exemplary condition illustrated is formed by the segments  330  and  332  and that portion of the line segment  326  between the line segments  330  and  332 . 
   The terrain warning boundaries for the climbing condition of  FIG. 14  include the vertical segment  324  which extends from the aircraft to vertical distance equal to the ΔH terrain floor of  FIG. 7  or  8 , whichever is smaller, below the aircraft forming a first vertical boundary. For a condition when the aircraft is climbing, the line segment  326  extends from the segment  324  at the flight path angle γ to form the lower terrain warning boundary. The vertical segment  330  at a distance equal to ½ of the LAD forms the second vertical terrain warning boundary. 
   The cut-off boundary  334  limits a portion of the terrain warning boundary. In particular, the cut-off boundary  334  forms the lower terrain warning boundary up to a point  340 , where the cut-off boundary  334  intersects the line segment  326 . Beyond the point  340 , a portion  342  of the line segment  340  forms the balance of the lower terrain warning boundary up to a distance equal to ½ of the LAD. 
   A look-up terrain advisory and terrain warning boundaries are illustrated in FIG.  15  and are applicable for the condition where the aircraft is not slowing for an off-airport landing. As will be discussed in more detail below, the look-up terrain advisory and warning boundaries start at altitudes DHYEL 2  and DHRED 2 , respectively, below the aircraft. These altitudes DHYEL 2  and DHRED 2  are modulated by the instantaneous sink rate (i.e. vertical speed, HDOT of the aircraft). The amount of modulation is equal to the estimated altitude loss for a pull-up maneuver, for example, at ¼ G (e.g 8 ft/sec 2 ) at the present sink rate. The altitudes DHRED 2  and DHYEL 2  are dependent upon the altitude loss during a pull-up maneuver (ALPU) and the altitude loss due to reaction time (ALRT) as given by:
 
ALRT=HDOT*T R ,  Eq. (16)
 
   where HDOT equals the vertical acceleration of the aircraft in feet/sec. and T R  equals the total reaction time of the pilot in seconds. 
   Assuming a pull-up maneuver is initiated at time T I , the altitude loss due to the pull-up maneuver ALPU may be determined by integrating the vertical velocity HDOT with respect to time as set forth below.
 
HDOT(t)=α*t+HDOT 0   Eq. (17)
 
   where “a” equals the pull-up acceleration and HDOT. 0  is a constant. 
   Integrating both sides of equation (17) yields the altitude loss as a function of time H(t) as provided in equation (18) below. 
               H   ⁡     (   t   )       =         1   2     ⁢   a   ⁢           ⁢     t   2       +       (     HDOT   0     )     ⁢   t               Eq   .           ⁢     (   18   )               
 
Assuming a constant acceleration during the pull-up maneuver, the time t until vertical speed reaches zero is given by equation (19). 
             t   =       -     HDOT   0       a             Eq   .           ⁢     (   19   )               
 
   Substituting equation (19) into equation (18) yields equation (20). 
               HDOT   0   2       2   ⁢   a             Eq   .           ⁢     (   20   )               
 
   Equation (20) thus represents the altitude loss during the pull-up maneuver. 
   An exemplary block diagram for generating the signals ALRT and ALPU is illustrated in FIG.  16 . In particular, a signal representative of the vertical velocity of the aircraft HDOT, available, for example, from a barometric altimeter rate circuit (not shown), is applied to a filter  350  in order to reduce nuisance warnings due to turbulence. The filter  350  may be selected with a transfer function of 1/(TAUDOT*S+1); where TAUDOT is equal to one second. The output of the filter  350  is a signal HDOTf, which represents the filtered instantaneous vertical speed; positive during climbing and negative during descent. 
   To obtain the altitude loss due to reaction time signal ALTR, the signal HDOTf is applied to a multiplier  352 . Assuming a pilot reaction time T r , for example, of 5 seconds, a constant  354  equal to 5 seconds is applied to another input of the multiplier  352 . The output of the multiplier  352  represents the signal ALTR, which is positive when HDOTf is negative and set to zero if the signal HDOTf is positive, which represents a climbing condition. More particularly, the signal HDOTf is applied to a comparator  356  and compared with a reference value, for example, zero. If the comparator  356  indicates that the signal HDOTf is negative, the signal HDOTf is applied to the multiplier  352 . During climbing conditions, the signal HDOTf will be positive. During such conditions, the comparator  356  will apply a zero to the multiplier  352 . 
   The altitude loss due to the pull-up maneuver signal ALPU is developed by a square device  358 , a divider  360  and a multiplier  362 . The filtered instantaneous vertical speed signal HDOTf is applied to the square device  358 . Assuming a constant acceleration during the pull-up maneuver, for example, equal to 8 feet/sec 2  (0.25  g ), a constant is applied to the multiplier  362  to generate the signal  2   a.  This signal,  2   a,  is applied to the divider  360  along with the output of the square device  350 . The output of the divider  360  is a signal (HDOTf) 2/2a, which represents the altitude loss during a pull-up maneuver signal ALPU. 
   These signals ALRT and ALPU are used to modulate the distance below the aircraft where terrain advisory and terrain warning boundaries begin during a look-up mode of operation. More particularly, during such a mode of operation, the portion of the ΔH terrain floor segment of the terrain caution envelope contributed by DHYEL 2  is modulated by the signals ALRT and ALPU while the ΔH terrain floor of the terrain warning boundary DHRED 2  is modulated by the signal ALPU as indicated in equations (21) and (22), respectively.
 
DHYEL 2 =¾*ΔH+ALRT+ALPU  Eq. (21)
 
DHRED 2 =½*Δ+ALPU,  Eq. (22)
 
   where ΔH represents the terrain floor of FIG.  8 ? as discussed above. 
   Thus, in  FIG. 15 , the look-up terrain caution boundary begins at a point  364  below the aircraft; equal to DHYEL 2 . If the flight path angle γ is less than a configurable datum THETA 2 , a terrain advisory boundary  366  extends from the point  364  to the advisory LAD at an angle equal to THETA 2 . Should the flight path angle y be greater than THETA 2 , the lower advisory boundary, identified with the reference numeral  368 , will extend from the point  364  at an angle equal to the flight path angle γ. 
   Similarly, the look-up terrain warning boundary begins at point  370  below the aircraft; equal to DHRED 2 . If the flight path angle γ is less than the angle THETA 2 , a warning boundary  372  extends from the point  370  at angle THETA 2  to the warning LAD. Should the flight path angle γ be greater than THETA 2 , a warning boundary  374  will extend at an angle equal to the flight path angle γ between the point  370  and the warning LAD. 
   Terrain Display 
   A display system, generally identified with the reference numeral  400 , is illustrated in  FIGS. 17 and 18 .  FIG. 17  contains a top level functional diagram of the warning computer/display interface, while  FIG. 18  represents a simplified block diagram for implementation of the display system  400  in accordance with the present invention. The display system  400  may include a microprocessor, for example, an Intel type 80486, 25 MHz type microprocessor (“ 486 ”) and an Analog Devices type Digital Signal Processor (DSP). The DSP is primarily used for calculating the RHO/THETA conversions to off load the  486  microprocessor or to facilitate display of terrain delta on a display device typically used to display weather radar data in a RHO/THETA format. Additional, details on the construction of the terrain data display overlay matrices, terrain data files and display drivers may be found in U.S. Pat. No. 5,839,080 which has been incorporated herein by reference. 
   The display system  400  is used to provide a visual indication of terrain which penetrates the terrain caution and terrain warning envelopes as well as for display of terrain in the current vicinity of the aircraft. The background terrain information may be displayed using dot patterns whose density varies as a function of the elevation of the terrain relative to the altitude of the aircraft and color coded according to the degree of threat. Terrain located within the caution and warning envelopes may be displayed in solid colors, such as yellow and red. 
     FIG. 19  illustrates how the terrain background information can be shown on the display  36 . As will be discussed in more detail below, the elevation of the terrain relative to the altitude of the aircraft is shown as a series of colored dot patterns whose density varies as a function of the distance between the aircraft and the terrain. The colors are used to distinguish between terrain caution and terrain warning indications. For example, red may be used to represent a terrain warning indication while yellow or amber is used to represent a terrain advisory indication. By using colored shapes for terrain threat indications and dot patterns of variable density for the terrain background information, clutter of the display is minimized. 
     FIG. 20  illustrates display  36  color and dot pattern densities suitable for use with helicopters and other aircraft capable of off-airport landings. Terrain located more than 1500 feet below the aircraft is not shown. Terrain located between 500 and 1500 feet below the aircraft is shown in green with a dot pattern density of 16%. Terrain less than 500 feet below the aircraft but greater than a predetermined amount beneath the aircraft is shown in green with a dot pattern density of 50%. 
   The predetermined amount defines the boundary between where terrain is shown as green, indicating non-threatening terrain, and where terrain is shown as yellow indicating a potential hazard. According to the present invention, this predetermined value is ascertained as a function of the aircraft phase of flight. For example, during cruise, the yellow/green boundary is located 250 feet below the aircraft. The cruise phase can be identified by monitoring the groundspeed of the aircraft. In a preferred embodiment of the invention, cruise is identified as a condition where the groundspeed is equal to or exceeds 90 kts. During approach conditions, the yellow/green boundary is located at the aircraft altitude. In a preferred embodiment of the invention, groundspeeds between 40 and 90 kts are indicative of the approach condition. During hover or landing conditions, the yellow/green boundary is located at the current aircraft altitude. In a preferred embodiment of the invention, the hover condition is defined as a groundspeed below 40 kts. Preferably, the groundspeeds defining the hover and cruise conditions also form corner points  78  and  82  of the ΔH curve of FIG.  8 . Terrain depicted in yellow above the yellow/green boundary is shown with a dot density of 25%. For terrain elevations located between 250 feet and 500 feet above the aircraft current altitude, the background terrain is colored yellow using a 50% dot density. Terrain elevations greater than 500 feet above the current aircraft elevation are colored red using a 50% dot density. Terrain located in the caution advisory envelope is colored solid yellow, while terrain located within the warning envelope is colored solid red. Bodies of water located at sea level may optionally be colored in cyan, or other shade of blue. 
     FIG. 21  illustrates an alternative terrain display suitable for use with the present invention. 
   The present invention thus enables those aircraft capable of routine, safe, off-airport landings to land at a location other than an airport without flying through terrain that would otherwise be depicted as yellow or red on a terrain display for fixed wing aircraft. Overflying terrain that is depicted as yellow or red during off airport landings, can lead to complacency, and the pilot may ignore such a display in the future believing a safe condition exists when in reality it does not. The present invention thus permits the background terrain display to reflect a safe approach to an off airport landing by matching the display of potentially hazardous terrain to phase of flight. The present invention thus also preserves the depiction of terrain as hazardous, and assists in the prevention of CFIT accidents, when the approach and landing conditions do not exist and the pilot is in hazardous proximity to terrain. 
   The invention has now been described with reference to the preferred embodiments. Variations and modifications will be readily apparent to those of skill in the art. For this reason, the invention is to be interpreted in view of the claims.