CA1241082A - Warning system for tactical aircraft - Google Patents
Warning system for tactical aircraftInfo
- Publication number
- CA1241082A CA1241082A CA000449621A CA449621A CA1241082A CA 1241082 A CA1241082 A CA 1241082A CA 000449621 A CA000449621 A CA 000449621A CA 449621 A CA449621 A CA 449621A CA 1241082 A CA1241082 A CA 1241082A
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- Prior art keywords
- aircraft
- warning
- altitude
- recited
- tactical
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- 238000011084 recovery Methods 0.000 claims abstract description 32
- 230000001133 acceleration Effects 0.000 claims description 6
- 230000002401 inhibitory effect Effects 0.000 claims 4
- 230000007423 decrease Effects 0.000 claims 1
- 230000003247 decreasing effect Effects 0.000 claims 1
- 230000035484 reaction time Effects 0.000 description 6
- 238000010586 diagram Methods 0.000 description 4
- 230000001960 triggered effect Effects 0.000 description 2
- 230000009471 action Effects 0.000 description 1
- 230000009977 dual effect Effects 0.000 description 1
- 238000000034 method Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000004044 response Effects 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
Classifications
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
- G01C5/00—Measuring height; Measuring distances transverse to line of sight; Levelling between separated points; Surveyors' levels
- G01C5/005—Measuring height; Measuring distances transverse to line of sight; Levelling between separated points; Surveyors' levels altimeters for aircraft
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01P—MEASURING LINEAR OR ANGULAR SPEED, ACCELERATION, DECELERATION, OR SHOCK; INDICATING PRESENCE, ABSENCE, OR DIRECTION, OF MOVEMENT
- G01P1/00—Details of instruments
- G01P1/07—Indicating devices, e.g. for remote indication
- G01P1/08—Arrangements of scales, pointers, lamps or acoustic indicators, e.g. in automobile speedometers
Abstract
WARNING SYSTEM FOR TACTICAL AIRCRAFT
Abstract of the Disclosure A system for warning the pilot of a tactical aircraft in a tactical maneuver such as a dive of the point at which recovery must be immediately initiated to prevent the aircraft from descending below a minimum desired altitude monitors the altitude, descent rate and pitch angle of the aircraft, and issues a voice warning when recovery must be initiated. The system also provides warnings of dangerous flight conditions during non-tactical or common route segments of a flight such as the take-off, landing and cruise seg-ments of a flight. Circuitry responsive to the arming of the weapons of the aircraft enables the tactical warning portion of the system and disables certain functions of the non-tactical portion of the system that may cause false warnings to be generated during tactical maneuvers.
Abstract of the Disclosure A system for warning the pilot of a tactical aircraft in a tactical maneuver such as a dive of the point at which recovery must be immediately initiated to prevent the aircraft from descending below a minimum desired altitude monitors the altitude, descent rate and pitch angle of the aircraft, and issues a voice warning when recovery must be initiated. The system also provides warnings of dangerous flight conditions during non-tactical or common route segments of a flight such as the take-off, landing and cruise seg-ments of a flight. Circuitry responsive to the arming of the weapons of the aircraft enables the tactical warning portion of the system and disables certain functions of the non-tactical portion of the system that may cause false warnings to be generated during tactical maneuvers.
Description
t~ '~
BAC~GROUND OF THE INVENTION
Field of the Invention This invention relates generally to ground proximity warning systems, and more particularly to a system that provides a warning to the pilot of an aircraft in a tactical maneuver such as a dive that recovery must be initiated immediately to prevent the aircraft from going below a minimum desired altitude selected by the MDA, or radio altitude "bug", setting. The system also provides warnings of dangerous flight conditions during non-tactical or common route segments of a flight such as take-off, landing and cruise.
DescriPtion of the Prior Art Ground proximity warning systems that provide warnings of various dangerous flight profiles are known. However, such systems are generally designed for transport aircraft, and are designed to provide warnings primarily in the take-off and landing phases of a flight to prevent inadvertent controlled flight into terrain. Such systems are not designed to provide warnings during tactical maneuvers such as dives wherein the aircraft is deliberately flown toward the ground at high speed and at a steep angle, and consequently would provide inadequate or false warnings under such conditions.
SUMMARY OF THE INVENTION
Accordingly, it is an object of the present invention to provide a warning system that overcomes i~
:~2~ S ~ '?~
many of the disadvantages of the prior art warning systems.
It is another object of the present invention to provide a warning of a dangerous flight condition during tactical maneuvers as well as during common route segments such as take-off, landing and cruise.
It is another object of the present invention to provide a warning to the pilot of an aircraft in a dive that recovery must be initiated immediately in order to prevent the aircraft from going below a minimum desired altitude.
Tactical aircraft present special problems in the design of ground proximity warning systems that warn the pilot of the aircraft of a dangerous flight condition. One such problem occurs because of the dual mode of operaton of such aircraft. Typically, such aircraft have a tact$cal or weapons delivery mode which includes maneuvers such as dives and various low level maneuvering, as well as a non-tactical or common route segments mode of operation that includes maneuvers such as take-off, landing and cruise. Because maneuvers encountered in the tactical mode of operation of an aircraft are substantially different than the maneuvers typically encountered in the non-tactical mode, a system designed for non-tactical or transport aircraft will not provide adequate warnings during some tactical maneuvers, while providing false warnings during others. Similarly, a system designed specifi-cally to provide warnings during tactical maneuvers, wi~l not provide adequate warnings during the less severe, non-tactical maneuvers of a flight.
Therefore, in accordance with a preferred embodiment of the invention, there is provided a system that provides warnings of dangerous flight conditons in tactical as well as non-tactical phases of flight.
-3~
'rhe system according to the invention monitors the ~pitch angle ana barometric descent rate of the aircraft in the tactical mode of operation, and calculates the laltitude at which recovery must be initiated to prevent the aircraft from going below a minimum desired altitude selected by the MDA, or radio altitude "bug" setting.
When the minimum recovery altitude is reached, the warning system issues a specific aural warning that is distinct from non-tactical warnings, such as, for example, "ABORT", to indicate that recovery must be initiated immediately. The tactical warnings are enabled only during tactical phases of flight, and certain non-tactical warnings are disabled during the tactical phases.
DESCRIPTION OF THE DRAWING
These and other objects and advantages of the present invention will become readily apparent upon consideration of the following detailed descrip-tion and attached drawing, wherein:
FIG. 1 is a logical functional block diagram of the warning system according to the invention;
FIG. 2 shows an aircraft in a dive, and illustrates the dynamics of dive recovery;
FIG. 3 is a graph showing the altitude loss of an aircraft during a 4G recovery from a dive as a function of descent rate and pitch angle;
FIG. 4 is a graph showing the altitude loss of an aircraft during a 4G recovery from a dive as a function of airspeed and pitch angle; and FIG. 5 is a block diagram of a ground proxi-mity warning system according to the invention usable for both tactical and non-tactical maneuvers.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
Referring now to the drawing, with particular attention to FIG. 1, there is illustrated an embodiment of the tactical warning portion of the ground proximity warning system according to the invention generally designated by the reference numeral 10. The system 10 according to the invention is illustrated in ~IG.
1 in logical block diagram form as a series of gates, comparators and the like for purposes of illustration;
however, it should be understood that the actual imple-mentation of the logic can be other than as shown in FIG. 1, with various digital and analog implementations being possible. The signals used by the system as described include radio altitude, barometric altitude rate, airspeed and pitch angle of the aircraft, as well as signals indicating whether the weapons are armed, along with various validity signals. Depending on the type of aircraft in which the warning system is installed, the signals shown in FIG. 1 can be obtained from individual instruments such as a baro-metric altimeter 12, a barometric altitude rate circuit 14, a radio altimeter 16, a pitch gyroscope 18, an airspeed signal source 20 such as an air data computer or airspeed indicator and various discrete circuit elements indicating whether the weapons are armed and whether there is weight on the wheels. Alternatively, these signals may be obtained from a digital data bus in certain newer aircraft, The tactical warning portion of the system according to the invention is designed to be operative only during a tactical phase of flight, such as a bombing run or ground attack where a steep diYe condi-tion is encountered. Accordingly, the system isenabled only when conditions indicative of such a phase of operation are present. In the present embod-iment, these conditions are that the weight of the aircraft is off the wheels, that the airspeed is greater than a predetermined airspeed, such as, for example, 300 kr,ots, and that the weapons are armed.
The enabling function is provided by an A~D
gate 22 that receives signals indicating that the weapons are armed and that the weight is off the wheels, as well as a signal from a comparator 24 coupled to the airspeed signal source 20. The comparator 24 compares the signal from the airspeed signal source 20 with a reference signal representative of a predetermined airspeed, for example, 300 knots, and indicates when the airspeed of the aircraft exceeds a predetermined airspeed, e.g. 300 knots. In addition, the AND gate 22 receives signals indicating that the radio altimter is not inhibited and that the barometric altimeter is not inhibited in order to inhibit the system in the event of a faulty instrument.
When all of the above conditions are met, the gate 22 enables another AND gate 26 to place the gate 26 under the control of a comparator 28 which receives a signal representative of the pitch angle of the aircraft from the pitch gyroscope 8, a signal representative of the barometric descent rate of the aircraft from the rate circuit 14, as well as a signal representative of the minimum descent altitude set by the MDA setting. Based on these input signals, the comparator 28 determines theminimum recovery altitude, or a warning altitude, Hw, at which recovery must be initiated to prevent the aircraft going below the minimum descent altitude. The comparator also receives a signal representative of the altitude above ground of the aircraft from the radio altimeter 16, and pro-yides a signal to the gate 26 in the event that thesignal received from the radio altimeter 16 is repre-sentative of an altitude below the minimum recovery altitude. ~nce such a signal is received, the gate 26 initiates the operation of the warning generator, preferably a digital voice warning generator, thereby causing the generator 30 to generate a voice warning.
The voice warning should be a specific warning, such as, for example, ~ABORT" to indicate to the pilot exactly what action must be taken to avoid a dangerous flight condition. The voice warning is applied, either directly or indirectly, to a transducer 32, such as, a loudspeaker or an earphone, to be communicated to the pilot.
In addition to the speed and pitch of the aircraft, there are other factors tha~ determine the altitude at which the dive recovery warning must be generated. These factors are related to the perfor-mance of the aircraft and include the reaction time of the aircraft and pilot, and the number of G's the aircraft can sustain during a pull up. In a standard recovery maneuver, the path of travel of the aircraft approximates an arc of a circle at the bottom of the pull-up maneuver. The radius of the circle is deter-mined by the number of G's that the aircraft can generate during the pull-up maneuver, with the radius of curvature of the arc being inversely proportional to the number of G's generated.
Referring to FIG. 2, there is shown an air-craft 50 that has just passed through the minimum recovery warning altitude, Hw, and is beginning its recovery. Assuming that the recovery is satisfactory, the aircraft will level off at the minimum descent altitude, MDA, and begin its climb as shown by the dashed line. The recovery path will be a circle to a first approximation, and the radius of the circle, R, will be determined by the number of G's that the air-craft can generate during its recovery.
For an object traveling in a circle, the normal acceleration, i.e., the r2dial, inwardly directed acceleration, is directly proportional to the square of velocity of the object and inversely ~2~ 6 proportional to the radius of the circular path the object is traveling. Stated mathematically, this relationship becomes:
v2 (1) an = -R
where an is the inwardly directed normal acceleration experienced by the object during the turn, V is the velocity of the object and R is the radius of curvature of the path of travel of the object.
If the descent angle of the aircraft 50, as represented by the dive angle or approximated by the pitch angle, is equal to e, the descent rate of the aircraft can be readily be related to the velocity of the aircraft using simple trigonometric functions, specifically:
BAC~GROUND OF THE INVENTION
Field of the Invention This invention relates generally to ground proximity warning systems, and more particularly to a system that provides a warning to the pilot of an aircraft in a tactical maneuver such as a dive that recovery must be initiated immediately to prevent the aircraft from going below a minimum desired altitude selected by the MDA, or radio altitude "bug", setting. The system also provides warnings of dangerous flight conditions during non-tactical or common route segments of a flight such as take-off, landing and cruise.
DescriPtion of the Prior Art Ground proximity warning systems that provide warnings of various dangerous flight profiles are known. However, such systems are generally designed for transport aircraft, and are designed to provide warnings primarily in the take-off and landing phases of a flight to prevent inadvertent controlled flight into terrain. Such systems are not designed to provide warnings during tactical maneuvers such as dives wherein the aircraft is deliberately flown toward the ground at high speed and at a steep angle, and consequently would provide inadequate or false warnings under such conditions.
SUMMARY OF THE INVENTION
Accordingly, it is an object of the present invention to provide a warning system that overcomes i~
:~2~ S ~ '?~
many of the disadvantages of the prior art warning systems.
It is another object of the present invention to provide a warning of a dangerous flight condition during tactical maneuvers as well as during common route segments such as take-off, landing and cruise.
It is another object of the present invention to provide a warning to the pilot of an aircraft in a dive that recovery must be initiated immediately in order to prevent the aircraft from going below a minimum desired altitude.
Tactical aircraft present special problems in the design of ground proximity warning systems that warn the pilot of the aircraft of a dangerous flight condition. One such problem occurs because of the dual mode of operaton of such aircraft. Typically, such aircraft have a tact$cal or weapons delivery mode which includes maneuvers such as dives and various low level maneuvering, as well as a non-tactical or common route segments mode of operation that includes maneuvers such as take-off, landing and cruise. Because maneuvers encountered in the tactical mode of operation of an aircraft are substantially different than the maneuvers typically encountered in the non-tactical mode, a system designed for non-tactical or transport aircraft will not provide adequate warnings during some tactical maneuvers, while providing false warnings during others. Similarly, a system designed specifi-cally to provide warnings during tactical maneuvers, wi~l not provide adequate warnings during the less severe, non-tactical maneuvers of a flight.
Therefore, in accordance with a preferred embodiment of the invention, there is provided a system that provides warnings of dangerous flight conditons in tactical as well as non-tactical phases of flight.
-3~
'rhe system according to the invention monitors the ~pitch angle ana barometric descent rate of the aircraft in the tactical mode of operation, and calculates the laltitude at which recovery must be initiated to prevent the aircraft from going below a minimum desired altitude selected by the MDA, or radio altitude "bug" setting.
When the minimum recovery altitude is reached, the warning system issues a specific aural warning that is distinct from non-tactical warnings, such as, for example, "ABORT", to indicate that recovery must be initiated immediately. The tactical warnings are enabled only during tactical phases of flight, and certain non-tactical warnings are disabled during the tactical phases.
DESCRIPTION OF THE DRAWING
These and other objects and advantages of the present invention will become readily apparent upon consideration of the following detailed descrip-tion and attached drawing, wherein:
FIG. 1 is a logical functional block diagram of the warning system according to the invention;
FIG. 2 shows an aircraft in a dive, and illustrates the dynamics of dive recovery;
FIG. 3 is a graph showing the altitude loss of an aircraft during a 4G recovery from a dive as a function of descent rate and pitch angle;
FIG. 4 is a graph showing the altitude loss of an aircraft during a 4G recovery from a dive as a function of airspeed and pitch angle; and FIG. 5 is a block diagram of a ground proxi-mity warning system according to the invention usable for both tactical and non-tactical maneuvers.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
Referring now to the drawing, with particular attention to FIG. 1, there is illustrated an embodiment of the tactical warning portion of the ground proximity warning system according to the invention generally designated by the reference numeral 10. The system 10 according to the invention is illustrated in ~IG.
1 in logical block diagram form as a series of gates, comparators and the like for purposes of illustration;
however, it should be understood that the actual imple-mentation of the logic can be other than as shown in FIG. 1, with various digital and analog implementations being possible. The signals used by the system as described include radio altitude, barometric altitude rate, airspeed and pitch angle of the aircraft, as well as signals indicating whether the weapons are armed, along with various validity signals. Depending on the type of aircraft in which the warning system is installed, the signals shown in FIG. 1 can be obtained from individual instruments such as a baro-metric altimeter 12, a barometric altitude rate circuit 14, a radio altimeter 16, a pitch gyroscope 18, an airspeed signal source 20 such as an air data computer or airspeed indicator and various discrete circuit elements indicating whether the weapons are armed and whether there is weight on the wheels. Alternatively, these signals may be obtained from a digital data bus in certain newer aircraft, The tactical warning portion of the system according to the invention is designed to be operative only during a tactical phase of flight, such as a bombing run or ground attack where a steep diYe condi-tion is encountered. Accordingly, the system isenabled only when conditions indicative of such a phase of operation are present. In the present embod-iment, these conditions are that the weight of the aircraft is off the wheels, that the airspeed is greater than a predetermined airspeed, such as, for example, 300 kr,ots, and that the weapons are armed.
The enabling function is provided by an A~D
gate 22 that receives signals indicating that the weapons are armed and that the weight is off the wheels, as well as a signal from a comparator 24 coupled to the airspeed signal source 20. The comparator 24 compares the signal from the airspeed signal source 20 with a reference signal representative of a predetermined airspeed, for example, 300 knots, and indicates when the airspeed of the aircraft exceeds a predetermined airspeed, e.g. 300 knots. In addition, the AND gate 22 receives signals indicating that the radio altimter is not inhibited and that the barometric altimeter is not inhibited in order to inhibit the system in the event of a faulty instrument.
When all of the above conditions are met, the gate 22 enables another AND gate 26 to place the gate 26 under the control of a comparator 28 which receives a signal representative of the pitch angle of the aircraft from the pitch gyroscope 8, a signal representative of the barometric descent rate of the aircraft from the rate circuit 14, as well as a signal representative of the minimum descent altitude set by the MDA setting. Based on these input signals, the comparator 28 determines theminimum recovery altitude, or a warning altitude, Hw, at which recovery must be initiated to prevent the aircraft going below the minimum descent altitude. The comparator also receives a signal representative of the altitude above ground of the aircraft from the radio altimeter 16, and pro-yides a signal to the gate 26 in the event that thesignal received from the radio altimeter 16 is repre-sentative of an altitude below the minimum recovery altitude. ~nce such a signal is received, the gate 26 initiates the operation of the warning generator, preferably a digital voice warning generator, thereby causing the generator 30 to generate a voice warning.
The voice warning should be a specific warning, such as, for example, ~ABORT" to indicate to the pilot exactly what action must be taken to avoid a dangerous flight condition. The voice warning is applied, either directly or indirectly, to a transducer 32, such as, a loudspeaker or an earphone, to be communicated to the pilot.
In addition to the speed and pitch of the aircraft, there are other factors tha~ determine the altitude at which the dive recovery warning must be generated. These factors are related to the perfor-mance of the aircraft and include the reaction time of the aircraft and pilot, and the number of G's the aircraft can sustain during a pull up. In a standard recovery maneuver, the path of travel of the aircraft approximates an arc of a circle at the bottom of the pull-up maneuver. The radius of the circle is deter-mined by the number of G's that the aircraft can generate during the pull-up maneuver, with the radius of curvature of the arc being inversely proportional to the number of G's generated.
Referring to FIG. 2, there is shown an air-craft 50 that has just passed through the minimum recovery warning altitude, Hw, and is beginning its recovery. Assuming that the recovery is satisfactory, the aircraft will level off at the minimum descent altitude, MDA, and begin its climb as shown by the dashed line. The recovery path will be a circle to a first approximation, and the radius of the circle, R, will be determined by the number of G's that the air-craft can generate during its recovery.
For an object traveling in a circle, the normal acceleration, i.e., the r2dial, inwardly directed acceleration, is directly proportional to the square of velocity of the object and inversely ~2~ 6 proportional to the radius of the circular path the object is traveling. Stated mathematically, this relationship becomes:
v2 (1) an = -R
where an is the inwardly directed normal acceleration experienced by the object during the turn, V is the velocity of the object and R is the radius of curvature of the path of travel of the object.
If the descent angle of the aircraft 50, as represented by the dive angle or approximated by the pitch angle, is equal to e, the descent rate of the aircraft can be readily be related to the velocity of the aircraft using simple trigonometric functions, specifically:
(2) Sin e wherein Hb is the descent rate of the aircraft, more specifically the barometric descent rate. In the illustrated embodiment of the invention, the barometric descent rate is used in the calculation; however, closure rate or radio altitude rate can be used in certain embodiments.
If the dive angle of the aircraft is then by geometric principles, the angle between a radius that is perpendicular to the path of travel of the aircraft and the vertical is also e (FIG. 2).
Consequently, the distance between the center 52 of the circle and the warning height, Hw, is equal to R cos e, and the distance between the warning height Hw and the minimum descent altitude, MDA, is equal to R(l-cos e). The latter relationship permits the amount of altitude loss, ~H, between the warning height, Hw, and the minimum descent altitude, MDA, to be ~-V~f~ 2 calculated as a function of the dive angle of the ,aircraft and the radius, R, of the circular path the ,aircraft is traveling.
Since the radius of the recovery path is a function of the velocity of the aircraft and the number of G's it can generate during a pull up, the altitude loss between the warning altitude and the MDA can be calculated in terms of the velocity of the aircraft and the number of G's being generated during pull up, }0 Moreover, because the descent rate of the aircraft is a function of the velocity of the aircraft and the dive angle, the altitude loss, ~H, during recovery can be calculated as a function of descent rate and the number of G's generated in the pull up.
This is accomplished as follows. Taking relationship l2), squaring it and making appropriate trigonometric substitutions, the following relationship results:
Hb2 Hb2 Hb2
If the dive angle of the aircraft is then by geometric principles, the angle between a radius that is perpendicular to the path of travel of the aircraft and the vertical is also e (FIG. 2).
Consequently, the distance between the center 52 of the circle and the warning height, Hw, is equal to R cos e, and the distance between the warning height Hw and the minimum descent altitude, MDA, is equal to R(l-cos e). The latter relationship permits the amount of altitude loss, ~H, between the warning height, Hw, and the minimum descent altitude, MDA, to be ~-V~f~ 2 calculated as a function of the dive angle of the ,aircraft and the radius, R, of the circular path the ,aircraft is traveling.
Since the radius of the recovery path is a function of the velocity of the aircraft and the number of G's it can generate during a pull up, the altitude loss between the warning altitude and the MDA can be calculated in terms of the velocity of the aircraft and the number of G's being generated during pull up, }0 Moreover, because the descent rate of the aircraft is a function of the velocity of the aircraft and the dive angle, the altitude loss, ~H, during recovery can be calculated as a function of descent rate and the number of G's generated in the pull up.
This is accomplished as follows. Taking relationship l2), squaring it and making appropriate trigonometric substitutions, the following relationship results:
Hb2 Hb2 Hb2
(3) v2 =
sin2e (l-coS2 e ) (l+cos e)tl- cos e) Rearranging relationship (1) and substituting relation- 5 ship (3) therein we obtain:
v2
sin2e (l-coS2 e ) (l+cos e)tl- cos e) Rearranging relationship (1) and substituting relation- 5 ship (3) therein we obtain:
v2
(4) R = _ = Hb2 an an(l+cos e ) (l-cos e ) Multiplying the relationship (4) by (l-cos e ), the altitude loss during the pull up, ~H, is obtained as follows:
Hb2
Hb2
5 (5) H = -an (l+cos e _g~ Q8~
The last-mentioned relationship defines the altitude loss of the aircraft between the time that recovery from the dive is initiated and the aircraft attains level flight; however, it does not take into account the reaction time of the pilot and aircraft, nor is the minimum desired altitude considered. These factors must also be entered into the equation defining the warning height, Hw, to assure that the warning is given soon enough to permit the pilot to recover safely.
The minimum descent altitude can readily be taken into account by adding the MDA setting to the altitude loss. Reaction time can be taken into account by adding a term equal to the reaction time of the pilot and aircraft, for example, 2 seconds, multiplied by lS the descent rate to the equation. Thus, the warning height Hw is given by the following equation:
The last-mentioned relationship defines the altitude loss of the aircraft between the time that recovery from the dive is initiated and the aircraft attains level flight; however, it does not take into account the reaction time of the pilot and aircraft, nor is the minimum desired altitude considered. These factors must also be entered into the equation defining the warning height, Hw, to assure that the warning is given soon enough to permit the pilot to recover safely.
The minimum descent altitude can readily be taken into account by adding the MDA setting to the altitude loss. Reaction time can be taken into account by adding a term equal to the reaction time of the pilot and aircraft, for example, 2 seconds, multiplied by lS the descent rate to the equation. Thus, the warning height Hw is given by the following equation:
(6) H - Hb2 + 2Hb + MDA
an tl + cos e wherein the factor multiplying the Hb term represents the 2 second reaction time of the pilot and the air-craft.
The above equation can be used to define a warning altitude, Hw, for various types of aircraft having various performance characteristics. However, a 4G pull up is one that is typically used during tactical maneuvers because it does not subject the pilot to excessive stress. Thus, by using 4G's as the normal acceleration, an, and solving the above equation for a 4G recovery and a 2 second reaction time using numerical techniques, the following result is obtained Hw = (Hb)2 *(0.005176+0.0000236~( ~p 0.0000321~( ep-40)+2Hb + MDA
wherein:
Hw is the radio altitude warning height in feet Hb is the barometric altitude rate in feet per second ep is the pitch angle in degrees MDA is the minimum desired altitude in feet.
In the above equation the terms (ep - 11) and (ep - 40) are not permitted to go negative. Thus, if the pitch angle is less than 11 degrees or less than 40 degrees, respectively, the (ep - 11) and ~p - 40) terms are set to zero. Also, the pitch angle of the aircraft is used instead of the dive angle to represent the descent angle because the pitch angle is readily obtained from a gyroscope or the like, and the dive angle is not; however, the actual dive angle can be used in the calc~lation. The dive angle differs from the pitch angle by the angle of attack of the wings, which is, typically only a few degrees. Thus, using the pitch angle instead of the dive angle in the calculation provides a satisfactory approximation of the warning height. Also, the numerical approxima-tion of the trigonometric equation defining the warning height is used because it permits the warning system according to the invention to make simple multiplica-tions rather than much more complex trigonometric calculations, thereby providing a faster response time.
The recovery envelope of the dive recovery mode of the warning system according to the invention for a 4G recovery is illustrated in FIG. 3. FIG. 3 shows the altitude loss, ~H, as a function of descent rate for various pitch angles. As can be seen from FIG. 3, the altitude loss, ~H, and consequently, the warning height, Hw, increases as the closure rate increases and as the pitch angle increases. However, c~
because the cosine function changes slowly for small angles, the curve of FIG. 3 is relatively independent of pitch angle at lower pitch angles, i.e., dive angles below 40 degrees. Thus, the same curve can be used to define the warning boundary for pitch angles of approximately 40 degrees and below.
Also, as previously stated, descent rate and airspeed are related by the sine of the dive angle.
Consequently, the warning envelope can also be defined in terms of airspeed rather than descent rate as is illustrated in FIG. 4, which shows the altitude loss, H, as a function of airspeed and pitch angle. Thus, in an alternative embodiment of the present invention, a signal representative of airspeed, rather than the descent rate, can be applied to the comparator 28 and the curves of FIG. 4, rather than those of FIG. 3, can be used to determine the altitude loss and the warning height.
Referring to FIG. 5, there is shown a portion of the dive recovery warning system utilized in con-junction with systems that generate warnings of a dangerous flight condition un~er various non-tactical flight conditions. In the block diagram of FIG. 5, like numbers are used to indicate the same components illustrated in FIG. 1, and primed numbers are used to designate components of the non-tactical portion of the warning system that are analogous to components of the tactical portion of the system that are desig-nated by a like unprimed number. Thus, the non-tactical portion of the warning system 10' is analogousto the tactical portion of the warning system 10, and the non-tactical warning generator 30' is analogous to the tactical warning generator 30.
The system 10' illustrated in FIG. 5 also utilizes various signals representative of various ~4~0~
flight parameters of the aircraft including airspeed, radio altitude, barometric altitude rate, radio altitude rate, a signal representative of the minimum descent ialtitude, MDA, a signal representative of the arming of the weap~ns, a signal representative of the position of the landing gear, as well as various validity signals (not shown). As in the case of the system 10, the system 10' receives the signals either from various individual instruments and discrete circuit elements, or from a digital data bus. The system 10' analyzes the received signals and provides a warning function in the event that a dangerous flight condition exists.
Typical functions performed by the system 10' include the functions described in copending Canadian applications filed concurrently herewith by the same inventors as the inventors of the present invention entitled ~NEGATIVE CLIMB AFTER TAKE-OFF WARNING SYSTEM" -449,618, "SYSTEM FOR ALERTING A PILOT OF A DANGEROUS FLIGHT
PROFILE DURING LOW LEVEL MANEUVERINGn - 449,619, "EXCESSIVE
TERRAIN CLOSURE WARNING SYSTEM~ - 449,620, "EXCESSIVE ~ESCENT
RATE WARNING SYSTEM FOR TACTICAL AIRCRAFT" - 44~,622, and "WARNING SYSTEM FOR AIRCRAFT LANDING WITH LANDING GEAR UP" -449,623. All of the above applications are assigned to the same assignee as the assignee of the present invention.
The system 10' may contain one or more of the warning functions disclosed in the above-identiied applications, or other warning functions, and is opera-tive during the non-tactical phases of aircraft opera-tion. The system 10' controls the operation of a non-tactical warning generator 30', and causes the generator 30' to generate various voice warnings that are applied to the transducer 32, or other suitable transducer, to inform the pilot of a dangerous flight condition during non-tactical phases of flight opera-.
-13~ Q~'~
l~ion. Typical of such warnings are ~DO~'T SINK", which indicates to a pilot that he is sinking after ltake-off, "TOO LOW", which indicates to a pilot that Ihe is flying below the minimum descent altitude, ~TERRAIN~, which warns a pilot of excessive terrain closure during low altitude flight, "SINK RATE", which indicates to a pilot that he is sinking too rapidly on approach to a landing, and "TOO LOW GEAR", which warns a pilot if he should attempt to lana with the landing gear up.
While all of these warnings are useful during the non-tactical phases of flight, some of these warnings can be triggered falsely during tactical phases of flight, thereby causing nuisance warnings.
In particular, the "TERRAIN" warning and the ~SINK
RATE" warning are likely to be generated during tacti-cal maneuvers because both the radio altitude rate and the barometric altitude rate are high during tacti-cal maneuvers. Consequently, those non-tactical modes most likely to be falsely triggered be inhibited during tactical maneuvers.
Consequently, the system 10' is provided with a weapons not armed input that inhibits the gèneration of any warnings, such as, for example, the "TERRAIN~ and "DON'T SINK" warnings, and other warnings which would cause nuiscance warnings to be generated during tactical maneuvers. Such a signal can be con-veniently obtained from a weapons arming switch 40 which is used to arm the weapons of the aircraft.
Such a switch can be used to provide a WEAPONS NOT
ARMED signal to the system 10' to inhibit any warnings that would be a nuisance during tactical flight, as well as to enable the tactical warning system 10 during the tactical phases of flight.
- 14 - ~24~08~
1 Obviously, many modifications and variations of the present invention are possible in light of the above teachings. Thus, ii is to be understood that, within the scope of the appended claims, the invention may be practiced otherwise than as specifically described above.
an tl + cos e wherein the factor multiplying the Hb term represents the 2 second reaction time of the pilot and the air-craft.
The above equation can be used to define a warning altitude, Hw, for various types of aircraft having various performance characteristics. However, a 4G pull up is one that is typically used during tactical maneuvers because it does not subject the pilot to excessive stress. Thus, by using 4G's as the normal acceleration, an, and solving the above equation for a 4G recovery and a 2 second reaction time using numerical techniques, the following result is obtained Hw = (Hb)2 *(0.005176+0.0000236~( ~p 0.0000321~( ep-40)+2Hb + MDA
wherein:
Hw is the radio altitude warning height in feet Hb is the barometric altitude rate in feet per second ep is the pitch angle in degrees MDA is the minimum desired altitude in feet.
In the above equation the terms (ep - 11) and (ep - 40) are not permitted to go negative. Thus, if the pitch angle is less than 11 degrees or less than 40 degrees, respectively, the (ep - 11) and ~p - 40) terms are set to zero. Also, the pitch angle of the aircraft is used instead of the dive angle to represent the descent angle because the pitch angle is readily obtained from a gyroscope or the like, and the dive angle is not; however, the actual dive angle can be used in the calc~lation. The dive angle differs from the pitch angle by the angle of attack of the wings, which is, typically only a few degrees. Thus, using the pitch angle instead of the dive angle in the calculation provides a satisfactory approximation of the warning height. Also, the numerical approxima-tion of the trigonometric equation defining the warning height is used because it permits the warning system according to the invention to make simple multiplica-tions rather than much more complex trigonometric calculations, thereby providing a faster response time.
The recovery envelope of the dive recovery mode of the warning system according to the invention for a 4G recovery is illustrated in FIG. 3. FIG. 3 shows the altitude loss, ~H, as a function of descent rate for various pitch angles. As can be seen from FIG. 3, the altitude loss, ~H, and consequently, the warning height, Hw, increases as the closure rate increases and as the pitch angle increases. However, c~
because the cosine function changes slowly for small angles, the curve of FIG. 3 is relatively independent of pitch angle at lower pitch angles, i.e., dive angles below 40 degrees. Thus, the same curve can be used to define the warning boundary for pitch angles of approximately 40 degrees and below.
Also, as previously stated, descent rate and airspeed are related by the sine of the dive angle.
Consequently, the warning envelope can also be defined in terms of airspeed rather than descent rate as is illustrated in FIG. 4, which shows the altitude loss, H, as a function of airspeed and pitch angle. Thus, in an alternative embodiment of the present invention, a signal representative of airspeed, rather than the descent rate, can be applied to the comparator 28 and the curves of FIG. 4, rather than those of FIG. 3, can be used to determine the altitude loss and the warning height.
Referring to FIG. 5, there is shown a portion of the dive recovery warning system utilized in con-junction with systems that generate warnings of a dangerous flight condition un~er various non-tactical flight conditions. In the block diagram of FIG. 5, like numbers are used to indicate the same components illustrated in FIG. 1, and primed numbers are used to designate components of the non-tactical portion of the warning system that are analogous to components of the tactical portion of the system that are desig-nated by a like unprimed number. Thus, the non-tactical portion of the warning system 10' is analogousto the tactical portion of the warning system 10, and the non-tactical warning generator 30' is analogous to the tactical warning generator 30.
The system 10' illustrated in FIG. 5 also utilizes various signals representative of various ~4~0~
flight parameters of the aircraft including airspeed, radio altitude, barometric altitude rate, radio altitude rate, a signal representative of the minimum descent ialtitude, MDA, a signal representative of the arming of the weap~ns, a signal representative of the position of the landing gear, as well as various validity signals (not shown). As in the case of the system 10, the system 10' receives the signals either from various individual instruments and discrete circuit elements, or from a digital data bus. The system 10' analyzes the received signals and provides a warning function in the event that a dangerous flight condition exists.
Typical functions performed by the system 10' include the functions described in copending Canadian applications filed concurrently herewith by the same inventors as the inventors of the present invention entitled ~NEGATIVE CLIMB AFTER TAKE-OFF WARNING SYSTEM" -449,618, "SYSTEM FOR ALERTING A PILOT OF A DANGEROUS FLIGHT
PROFILE DURING LOW LEVEL MANEUVERINGn - 449,619, "EXCESSIVE
TERRAIN CLOSURE WARNING SYSTEM~ - 449,620, "EXCESSIVE ~ESCENT
RATE WARNING SYSTEM FOR TACTICAL AIRCRAFT" - 44~,622, and "WARNING SYSTEM FOR AIRCRAFT LANDING WITH LANDING GEAR UP" -449,623. All of the above applications are assigned to the same assignee as the assignee of the present invention.
The system 10' may contain one or more of the warning functions disclosed in the above-identiied applications, or other warning functions, and is opera-tive during the non-tactical phases of aircraft opera-tion. The system 10' controls the operation of a non-tactical warning generator 30', and causes the generator 30' to generate various voice warnings that are applied to the transducer 32, or other suitable transducer, to inform the pilot of a dangerous flight condition during non-tactical phases of flight opera-.
-13~ Q~'~
l~ion. Typical of such warnings are ~DO~'T SINK", which indicates to a pilot that he is sinking after ltake-off, "TOO LOW", which indicates to a pilot that Ihe is flying below the minimum descent altitude, ~TERRAIN~, which warns a pilot of excessive terrain closure during low altitude flight, "SINK RATE", which indicates to a pilot that he is sinking too rapidly on approach to a landing, and "TOO LOW GEAR", which warns a pilot if he should attempt to lana with the landing gear up.
While all of these warnings are useful during the non-tactical phases of flight, some of these warnings can be triggered falsely during tactical phases of flight, thereby causing nuisance warnings.
In particular, the "TERRAIN" warning and the ~SINK
RATE" warning are likely to be generated during tacti-cal maneuvers because both the radio altitude rate and the barometric altitude rate are high during tacti-cal maneuvers. Consequently, those non-tactical modes most likely to be falsely triggered be inhibited during tactical maneuvers.
Consequently, the system 10' is provided with a weapons not armed input that inhibits the gèneration of any warnings, such as, for example, the "TERRAIN~ and "DON'T SINK" warnings, and other warnings which would cause nuiscance warnings to be generated during tactical maneuvers. Such a signal can be con-veniently obtained from a weapons arming switch 40 which is used to arm the weapons of the aircraft.
Such a switch can be used to provide a WEAPONS NOT
ARMED signal to the system 10' to inhibit any warnings that would be a nuisance during tactical flight, as well as to enable the tactical warning system 10 during the tactical phases of flight.
- 14 - ~24~08~
1 Obviously, many modifications and variations of the present invention are possible in light of the above teachings. Thus, ii is to be understood that, within the scope of the appended claims, the invention may be practiced otherwise than as specifically described above.
Claims (18)
1. A system for indicating to the pilot of an aircraft in a dive the altitude at which recovery must be initiated to avoid descending below a predetermined altitude, comprising:
means responsive to the descent rate and descent angle of the aircraft for determining the minimum altitude at which recovery must be initiated to avoid descending below the predetermined altitude; and means responsive to the altitutde of the aircraft and to said determining means for providing a warning to the pilot when recovery must be initiated.
means responsive to the descent rate and descent angle of the aircraft for determining the minimum altitude at which recovery must be initiated to avoid descending below the predetermined altitude; and means responsive to the altitutde of the aircraft and to said determining means for providing a warning to the pilot when recovery must be initiated.
2. A system as indicated in claim 1 wherein said descent rate and descent angle responsive means includes means responsive to the barometric descent rate of the aircraft.
3. A system as recited in claim 1 wherein said descent rate and descent angle responsive means includes means responsive to the pitch angle of the aircraft.
4. A system as recited in claim 1 wherein said means responsive to the altitude of the aircraft includes means responsive to the altitude above ground of the aircraft.
5. A system as recited in claim 1, wherein the aircraft is an aircraft having a tactical and a non-tactical phase of flight, and further including:
(a) added warning means, responsive to flight conditions of the aircraft during a non-tactical phase of flight, for generating a warning indicative of a dangerous flight condition; and (b) means for inhibiting the generation of a warning by said added warning means when the aircraft is in a tactical phase of flight.
(a) added warning means, responsive to flight conditions of the aircraft during a non-tactical phase of flight, for generating a warning indicative of a dangerous flight condition; and (b) means for inhibiting the generation of a warning by said added warning means when the aircraft is in a tactical phase of flight.
6. A ground proximity warning system as recited in claim 5 wherein said inhibiting means includes means responsive to a weapons arming signal.
7. A ground proximity warning system as recited in claim 5 wherein said added warning generating means is responsive to the altitude of the aircraft above the ground for generating a warning if the altitutde above ground decreases at an excessive rate.
8. A ground proximity warning system as recited in claim 7 wherein said inhibiting means includes means for preventing the generation of the warning that the altitude above ground is decreasing excessively rapidly.
9. A ground proximity warning system as recited in claim 5 wherein said added warning generating means is responsive to the descent rate of the aircraft for
9. A ground proximity warning system as recited in claim 5 wherein said added warning generating means is responsive to the descent rate of the aircraft for
Claim 9 continued....
generating a warning if the aircraft is descending at an excessive rate.
generating a warning if the aircraft is descending at an excessive rate.
10. A ground proximity warning system as recited in claim 9 wherein said inhibiting means includes means for preventing the generation of the warning that the aircraft is descending excessively rapidly when the weapons are armed.
11. A system as recited in claim 1, further including means for setting the predetermined minimum altitude.
12. A system as recited in claim 4 wherein said means for providing a signal representative of the aircraft above ground includes a radio altimeter.
13. A system as recited in claim 1 further including means reponsive to the speed of the aircraft for enabling the generation of a warning only when the speed of the aircraft exceeds a predetermined speed.
14. A system as recited in claim 13 wherein said speed signal providing means includes an air speed indicator, and said enabling means includes means for enabling the generation of a warning only when the airspeed of the aircraft exceeds a predetermined airspeed.
15. A system as recited in claim 14 wherein said predetermined airspeed is approximately 300 knots.
16. A system as recited in claim 1 further including means for providing a signal representative of the airspeed of the aircraft, wherein said warning generating means is further responsive to the airspeed of the aircraft for generating the warning.
17. A system as recited in claim 1 wherein said warning generating means further includes means for generating a specific voice warning.
18. A warning system as recited in claim 1 wherein said altitude loss is a function of the normal acceleration of the aircraft during recovery, and wherein said determining means includes means for determining the altitude loss for a predetermined normal acceleration.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US49459283A | 1983-05-13 | 1983-05-13 | |
| US494,592 | 1983-05-13 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1241082A true CA1241082A (en) | 1988-08-23 |
Family
ID=23965106
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000449621A Expired CA1241082A (en) | 1983-05-13 | 1984-03-14 | Warning system for tactical aircraft |
Country Status (15)
| Country | Link |
|---|---|
| JP (1) | JPS59216791A (en) |
| AU (1) | AU547207B2 (en) |
| BE (1) | BE899645A (en) |
| CA (1) | CA1241082A (en) |
| CH (1) | CH659983A5 (en) |
| DE (1) | DE3417885A1 (en) |
| ES (1) | ES8507405A1 (en) |
| FI (1) | FI74253C (en) |
| FR (1) | FR2545930B1 (en) |
| GB (2) | GB2139589B (en) |
| GR (1) | GR79932B (en) |
| IL (1) | IL71350A (en) |
| IT (1) | IT1177718B (en) |
| NL (1) | NL8401534A (en) |
| SE (1) | SE8402469L (en) |
Families Citing this family (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5001476A (en) * | 1983-05-13 | 1991-03-19 | Sundstrand Data Control, Inc. | Warning system for tactical aircraft |
| US4980684A (en) * | 1983-06-10 | 1990-12-25 | Sundstrand Data Controls, Inc. | Warning system for tactical rotary wing aircraft |
| CA1243405A (en) * | 1985-02-22 | 1988-10-18 | Michael M. Grove | Configuration responsive descent rate warning system for aircraft |
| CA1243118A (en) * | 1985-02-22 | 1988-10-11 | Michael M. Grove | Aircraft terrain closure warning system with descent rate based envelope modification |
| CA1243119A (en) * | 1985-02-22 | 1988-10-11 | Michael M. Grove | Aircraft terrain warning system with configuration modified warning and improved mode switching |
| US4916448A (en) * | 1988-02-26 | 1990-04-10 | The United States Of America As Represented By The Secretary Of The Air Force | Low altitude warning system for aircraft |
| US5864307A (en) * | 1996-02-19 | 1999-01-26 | Gec Marconi Limited | Aircraft terrain advisory system |
| CN105203075B (en) * | 2015-09-15 | 2017-07-28 | 北京安达维尔航空设备有限公司 | Radio altimeter data fusion height-finding system and the high method of survey |
| FR3063715B1 (en) | 2017-03-09 | 2019-04-12 | Dassault Aviation | AIRCRAFT FLOOR PROXIMITY WARNING SYSTEM, AVIONICS AND ASSOCIATED METHOD |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2736878A (en) * | 1954-01-26 | 1956-02-28 | Jr Edward B Boyle | Dive-bombing breakaway computer |
| US3715718A (en) * | 1970-08-11 | 1973-02-06 | Sundstrand Data Control | Ground proximity warning system utilizing radio and barometric altimeter combination |
-
1984
- 1984-03-14 CA CA000449621A patent/CA1241082A/en not_active Expired
- 1984-03-26 IL IL71350A patent/IL71350A/en unknown
- 1984-04-10 AU AU26688/84A patent/AU547207B2/en not_active Ceased
- 1984-04-27 CH CH2079/84A patent/CH659983A5/en not_active IP Right Cessation
- 1984-05-08 GR GR74650A patent/GR79932B/el unknown
- 1984-05-08 SE SE8402469A patent/SE8402469L/en not_active Application Discontinuation
- 1984-05-09 GB GB08411770A patent/GB2139589B/en not_active Expired
- 1984-05-10 JP JP59091971A patent/JPS59216791A/en active Pending
- 1984-05-11 ES ES532431A patent/ES8507405A1/en not_active Expired
- 1984-05-11 NL NL8401534A patent/NL8401534A/en not_active Application Discontinuation
- 1984-05-11 FI FI841912A patent/FI74253C/en not_active IP Right Cessation
- 1984-05-11 BE BE0/212923A patent/BE899645A/en not_active IP Right Cessation
- 1984-05-11 IT IT48178/84A patent/IT1177718B/en active
- 1984-05-11 FR FR8407320A patent/FR2545930B1/en not_active Expired
- 1984-05-14 DE DE19843417885 patent/DE3417885A1/en not_active Ceased
-
1986
- 1986-05-06 GB GB08611004A patent/GB2174665B/en not_active Expired
Also Published As
| Publication number | Publication date |
|---|---|
| FI841912A0 (en) | 1984-05-11 |
| GB2174665A (en) | 1986-11-12 |
| GB8611004D0 (en) | 1986-06-11 |
| NL8401534A (en) | 1984-12-03 |
| CH659983A5 (en) | 1987-03-13 |
| SE8402469D0 (en) | 1984-05-08 |
| ES532431A0 (en) | 1985-09-01 |
| FI841912A (en) | 1984-11-14 |
| ES8507405A1 (en) | 1985-09-01 |
| AU547207B2 (en) | 1985-10-10 |
| FI74253B (en) | 1987-09-30 |
| JPS59216791A (en) | 1984-12-06 |
| SE8402469L (en) | 1984-11-14 |
| GB2139589A (en) | 1984-11-14 |
| IT8448178A0 (en) | 1984-05-11 |
| GB2174665B (en) | 1987-11-25 |
| GB8411770D0 (en) | 1984-06-13 |
| IL71350A (en) | 1990-02-09 |
| IT1177718B (en) | 1987-08-26 |
| DE3417885A1 (en) | 1984-11-15 |
| FI74253C (en) | 1988-01-11 |
| BE899645A (en) | 1984-11-12 |
| GR79932B (en) | 1984-10-31 |
| GB2139589B (en) | 1987-11-25 |
| FR2545930B1 (en) | 1988-01-15 |
| AU2668884A (en) | 1984-11-15 |
| FR2545930A1 (en) | 1984-11-16 |
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| MKEX | Expiry |