BACKGROUND 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 if the descent rate of the aircraft is exces- sive for the altitude at which the aircraft is flying. The criteria for determining whether or not such a warning is to be generated are alterable as a function of the configuration of the aircraft. Such a warning system is particularly useful for carrier based anti- submarine aircraft such as the Lockheed S3-A. Description of the Prior Art

Systems for providing the pilot of an air¬ craft with a warning in the event that the aircraft is descending too rapidly are Known. Examples of such systems are disclosed in United States Patent Nos. 3,946,358; 3,947,808; 3,958,219 and 4,215,334, all assigned to the same assignee as the assignee of the present invention.

While all of the above systems serve the basic function of alerting the pilot if the descent rate of the aircraft exceeds a predetermined safe descent rate determined by the altitude above ground of the aircraft, the above systems are designed to operate with transport aircraft whose operational and

flight conditions are such that the approach and land¬ ing conditions are readily predictable. However, for aircraft such as, for example, military and tactical aircraft whose operational and flight conditions can vary widely, systems designed for aircraft having predictable flight parameters tend to provide false or nuisance warnings where no unsafe condition actual¬ ly exists, and may provide less than the desired warn¬ ing times under other conditions. SUMMARY OF THE INVENTION

Accordingly, it is an object of the present invention to provide a ground proximity warning system that overcomes many of the disadvantages of the prior art warning systems. It is another object of the present inven¬ tion to provide a warning to the pilot of an aircraft if the descent rate of the aircraft is excessive for the altitude above ground and flight configuration o the aircraft. It is yet another object of the present invention to provide a ground proximity warning system for alerting the pilot of an aircraft if its descent rate is dangerously high, and where the criteria for determining whether the descent rate is excessively high are altered as a function of flight configuration of the aircraft.

It is another object of the present inven¬ tion to provide a warning system that generates a warning if the descent rate of the aircraft is exces- sively high for the altitude at which the aircraft is flying, but does not generate an appreciable number of false or nuisance warnings when used in a manuever- able aircraft.

It is yet another object of the present invention to provide a ground proximity warning system

particularly useful for tactical aircraft wherein the criteria used to determine whether a warning should be generated are adjusted depending on the flight configuration of the aircraft. Briefly, in accordance with a preferred embodiment of the invention, there is provided a system that monitors the altitude above ground and the descent rate of an aircraft and issues a first aural warning if the descent rate exceeds a predetermined safe limit for the radio altitude at which the aircraft is flying. If the descent rate exceeds the predetermined safe limit by a predetermined margin, a second distinct aural warning is generated to warn the pilot of a particularly dangerous flight condition. The criteria for generating the warnings are altered as a function of the flight configuration of the aircraft, for example, whether the landing gear is up or down and whether the aircraft is flying a tactical or non-tac¬ tical segment of a mission to optimi2e the warning criteria for the operational conditions of the air¬ craft.

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 block diagram of the warning system according to the invention;

FIG. 2 is a graph of the descent rate re- quired to ge.nerate the two warnings as a function of radio altitude when the aircraft is operating in a gear up configuration;

FIG. 3 is a graph of the descent rates re¬ quired to generate the two warnings as a function of

radio altitude when the aircraft is operating with its landing gear down; and

FIG. 4 is a graph of the descent rate requir¬ ed to generate the two warnings as a function of radio altitude when the aircraft is operating in a tactical mode and the landing gear is up.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

Referring now to the drawing with particular attention to FIG. 1, there is illustrated an erabodi- ment of the configuration responsive 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 FIG. 1 in logical bloc* diagram form as a series of gates, comparators and the like for purposes of illustration; however, it should be understood actual implementation 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 and barometric altitude rate, a signal indicative of a tactical mission, for example, a mission selector switch, a landing gear position switch, a weight on wheels switch and various validity signals. Depending on the type of aircraft on which the warning system is installed, the signals shown in FIG. 1 can be obtained from individual instruments such as an air data computer 12, a radio altimeter 14, a gear position switch 16, a mission selector switch 18 and a weight on wheels switch 20. Alterna- tively, the -barometric rate signal may be obtained by differentiating a barometric altitude signal from a barometric altimeter, or the 2-velocity signal from an inertial navigation system may be used to obtain descent rate information. Also, rather than obtaining the signals from individual instruments.

the signals can be obtained from a digital data bus in certain newer aircraft.

As previously stated, the system according to the invention compares the altitude above ground with the descent rate, preferably barometric descent rate or Z-velocity, and issues a warning if the descent rate is excessive for the altitude at which the aircraft is flying. The comparison function is provided by two mode comparators 22 and 24 which, in the present embodiment, compare the radio altitude received from the radio altimeter 14 via a filter 26 which removes high frequency components from the radio altitude signal. The radio altitude signal is compared with the barometric rate signal that is applied to the comparators 22 and 24 via a low pass filter 28 and a pair of function generators 30 and 32, respec¬ tively. The radio altitude signal is also compared with a 2450 foot reference signal and a 30 foot refer¬ ence signal by a pair of comparators 34 and 36 to provide enabling singals to the system when the air¬ craft is flying between 30 feet and 2450 feet. Although, in the embodiment illustrated in FIG. 1, the comparators and function generators have ύeen illustrated as separate components for purposes of clarity, the comparison function required to generate the appropriate warnings could be provided by a single comparator, or otherwise. For example, the function generation and comparison functions could be combined in a single block. In addition to the comparison and warning generation functions discussed above, the system utilizes a plurality of gates and discrete switches to enable the system only during predetermined phases of flight. These enabling functions are provided by the weight on wheels switch which cooperates with an

inverter 38 and a delay circuit 40 to close a switch 42 to permit the barometric rate signal to be applied to the filter 28 only when there is no weight on the wheels of the aircraft, that is, the aircraft is off the ground, in addition, a series of AND gates 44, 46 and 48 and delay circuits 51 and 53 enable the system only when the signal from the radio altimeter and the air data computer are valid, when there is no weight on the wheels and when the aircraft is flying within a predetermined range of altitudes, for example, between 30 and 2450 feet.

In addition, the system includes a pair of aural warning generators such as a sink rate warning generator 50 and a pull up warning generator 52 that are activated by signals received from the gates 46 and 48 via a pair of delay circuits 54 and 56 and apply either a "SINK RATE" or a "PULL UP" warning to a transducer 58.

In operation, whenever the aircraft is flying between 30 and 2450 feet the gates 44, 46 and 48 are enabled and the switch 42 is closed so that the baro¬ metric rate signal is applied to the function genera¬ tors 30 and 32. Under these conditions, if the com¬ bination of descent rate and altitude above ground is such that the warning criteria established by the function generator 30 and the comparator 22 are exceed¬ ed, the sink rate comparator 22 will apply a signal to the AND gate 46, thereby causing the AND gate 46 to apply a warning initiation signal to the sink rate generator 50 after delay circuit 54 has operated. Similarly, if the combination of descent rate and altitude above ground is such that the warn¬ ing criteria established by the function generator 32 and the comparator 24 are exceeded, the comparator 24 will apply an enabling signal to the AND gate 48 to apply a warning initiation signal to the pull up warn-

ing generator 52, after the delay circuit 56 has operated. The sink rate and pull up warning generators 50 and 52 are preferably digital voice generators that apply the warnings to the transducer 58 either directly or indirectly, for example, through" the intercom system of the aircraft.

As previously stated, the criteria for determining whether a sink rate or a pull up warning are to be generated are determined by the configura- tion of the aircraft. In the illustrated embodiment, the modification of the warning criteria is accomplish¬ ed by modifying the characteristics of the function generators 30 and 32 by discrete signals from the gear position switch 16 and the mission selector switch 18. The mission selector switch is actuated by the pilot whenever a tactical mission, such as an anti-submarine reconnaissance or attack mission is being flown. In the illustrated embodiment, the func¬ tion generators 30 and 32 are altered to modify the barometric rate signal from the filter 28 in order to obtain the desired warning criteria; however, in al¬ ternative embodiments the radio altitude signal could be altered instead.

When the aircraft is flying with its gear up, the function provided by the function generator

30 is selected such that the "SINK RATE" warning will be generated whenever the relationship between the descent rate and the radio altitude is such that it falls within the envelope below the line 100 (FIG. 2) . The "PULL UP" warning will be initiated if the envelope below the line 102 is penetrated. Because of the delays provided by the delay 54 (0.8 second) and the delay 56 (1.6 seconds) the actual warning will not occur until the envelope illustrated by the dashed lines 100' and 102' is penetrated. Thus, the delays serve to reduce nuisance warnings in instances

where the envelopes 100 and 102 are only momentarily penetrated.

Whenever the aircraft lowers its landing gear, and is not in a tactical mode, the warning criteria are altered as is shown in FIG. 3 to make the warning system more sensitive at higher altitudes. Thus, with the landing gear down, the "SINK RATE" warn¬ ing will be given at a descent rate of 9800 feet per minute at an altitude of 2450 feet as opposed to the 11,800 feet per minute descent rate required at the same altitude with the landing gear up. Similarly, the pull up warning will be given at 12,550 feet per minute at 2450 feet rather than at 14,700 feet per minute. This takes into account the reduced recovery capabil- ity of an aircraft when the landing gear is down.

Whenever an aircraft is in a tactical phase of flight, such as, for example, when it is attacking a target, higher descent rates are encountered than are encountered in a non-tactical phase of flight. Thus, the warning criteria must be modified according¬ ly, in order to avoid nuisance warnings, and the cri¬ teria illustrated in FIG. 4 are utilized to provide additional maneuvering room for the aircraft before a warning is generated. These criteria are used only when both the landing gear is up and the mission selector switch indicates a tactical mode of opera¬ tion.

The tactical mode criteria require a 13,800 foot per minute and a 16,700 foot per minute descent rate before the "SINK RATE" and "PULL UP" warnings, respectively, are generated when the altitude of the aircraft is 2450 feet. In addition, the descent rates required for a warning to be generated at low altitudes, for example, below 450 feet, are also sub- stantially increased. For example, at 450 feet a

3800 foot per minute descent rate is required before the "SINK RATE" warning is generated, and at 30 feet of radio altitude, a descent rate of 2960 feet per minute is required. To generate a "PULL UP" warning at 30 feet of radio altitude, a 3474 feet per minute descent rate is required, and 3800 feet per minute is required at 300 feet of radio altitude. Thus, the system permits the aircraft to maneuver at low alti¬ tudes, for example, as it does during submarine recon- naissance missions without generating nuisance warn¬ ings.

Obviously, many modifications and variations of the present invention are possible in light of the above teachings. Thus, it is to be understood that, within the scope of the appended claims, the invention may be practiced otherwise than as specifically des¬ cribed above.

What is claimed and desired to be secured by Letters Patent of the United States is: